Next Generation Spacecraft, Crew Exploration Vehicle

Home , Aerospike engine, Reusable spacecraft, Space Launch Initiative, VentureStar, Vernier thruster

A Special Bibliography From the NASA Scientific and Technical Information Program

Includes research on reusable launch vehicles, aerospace planes, shuttle replacement, crew/cargo transfer vehicle, related X-craft, orbital space plane, and next generation launch technology.

January 2004 Next Generation Spacecraft, Crew Exploration Vehicle A Special Bibliography from the NASA Scientiﬁc and Technical Information Program

JANUARY 2004

20040004384 Nebraska Univ., Lincoln, NE, USA Design of Experiments for the Thermal Characterization of Metallic Foam Crittenden, Paul E.; Cole, Kevin D.; Optimal Experiment Design for Thermal Characterization of Functionally Graded Materials; August 18, 2003; In English Contract(s)/Grant(s): NAG1-01087; No Copyright; Avail: CASI; A03, Hardcopy Metallic foams are being investigated for possible use in the thermal protection systems of reusable launch vehicles. As a result, the performance of these materials needs to be characterized over a wide range of temperatures and pressures. In this paper a radiation/conduction model is presented for heat transfer in metallic foams. Candidates for the optimal transient experiment to determine the intrinsic properties of the model are found by two methods. First, an optimality criterion is used to find an experiment to find all of the parameters using one heating event. Second, a pair of heating events is used to determine the parameters in which one heating event is optimal for finding the parameters related to conduction, while the other heating event is optimal for finding the parameters associated with radiation. Simulated data containing random noise was analyzed to determine the parameters using both methods. In all cases the parameter estimates could be improved by analyzing a larger data record than suggested by the optimality criterion. Author Metal Foams; Heat Transfer

20040003713 NASA Langley Research Center, Hampton, VA, USA Development of Structural Health Management Technology for Aerospace Vehicles Prosser, W. H.; [2003]; In English; JANNAF 39th Combustion/27th Airbreathing Propulsion/21st Propulsion Systems Harzards/3rd Modeling and Simulation Joint Subcommittee Meeting, 1-5 Sep. 2003, Colorado Springs, CO, USA; Original contains color illustrations; No Copyright; Avail: CASI; A02, Hardcopy As part of the overall goal of developing Integrated Vehicle Health Management (IVHM) systems for aerospace vehicles, NASA has focused considerable resources on the development of technologies for Structural Health Management (SHM). The motivations for these efforts are to increase the safety and reliability of aerospace structural systems, while at the same time decreasing operating and maintenance costs. Research and development of SHM technologies has been supported under a variety of programs for both aircraft and spacecraft including the Space Launch Initiative, X-33, Next Generation Launch Technology, and Aviation Safety Program. The major focus of much of the research to date has been on the development and testing of sensor technologies. A wide range of sensor technologies are under consideration including ﬁber-optic sensors, active and passive acoustic sensors, electromagnetic sensors, wireless sensing systems, MEMS, and nanosensors. Because of their numerous advantages for aerospace applications, most notably being extremely light weight, ﬁber-optic sensors are one of the leading candidates and have received considerable attention. Author Aerospace Vehicles; Systems Health Monitoring; Systems Engineering; Technology Utilization; Remote Sensors

20040001677 Aerospace Corp., USA Liquid Engine Test Facilities Assessment Adams, Michael J.; Emdee, Jeffery L.; George, Daweel J.; Peinemann, Manfred; May 03, 2002; In English Contract(s)/Grant(s): F04701-00-C-0009 Report No.(s): NASA/SE-2002-05-00041-SSC; No Copyright; Avail: CASI; A04, Hardcopy The John C. Stennis Space Center (SSC) requested The Aerospace Corporation to examine the current testing capability of all existing large liquid engine test facilities located in the USA. That information along with projected liquid rocket engine development was used to examine future liquid rocket engine testing facilities needs in the coming decade. Current domestic liquid engine test facilities capabilities, when examined against engine concepts for the coming decade, indicate there are

1 ample facilities offering altitude simulation during test. In addition, it was observed that many contractor facilities have limited ambient test capability of larger thrust engines under current consideration. Finally, it was concluded that diminished contractor participation engine development testing will drive this activity to the government sector. Only three facilities are seen as key contributors to engine testing in the coming decade, namely John C. Stennis Space Center (SSC), Marshall Space Flight Center (MSFC), and Air Force Research Laboratory (AFRL). Past rocket engine test experience was evaluated as a possible resource for projecting future engine test needs. A database comprised of various engine models and the level of testing performed to flight qualify those systems for their first flight was constructed. For comparison purposes in this study, development and qualification efforts were totaled and treated as one test program. Based on experience with past Air Force programs, the time on the test stand accounts for typically 50% or more of the total program time. Historical data show that the time to design and develop new engines has increased over the last 40 years, most likely due to scarcer resources in today’s funding environment. Derived from text Engine Tests; Liquid Propellant Rocket Engines; Test Facilities; Aerospace Systems

20040001153 NASA Langley Research Center, Hampton, VA, USA Cryopumping in Cryogenic Insulations for a Reusable Launch Vehicle Johnson, Theodore F.; Weiser, Erik S.; Grimsley, Brian W.; Jensen, Brian J.; [2003]; In English; No Copyright; Avail: CASI; A03, Hardcopy Testing at cryogenic temperatures was performed to verify the material characteristics and manufacturing processes of reusable propellant tank cryogenic insulations for a Reusable Launch Vehicle (RLV). The unique test apparatus and test methods developed for the investigation of cryopumping in cryogenic insulations are described. Panel level test specimens with various types of cryogenic insulations were subjected to a specific thermal profile where the temperature varied from -262 C to 21 C. Cryopumping occurred if the interior temperature of the specimen exhibited abnormal temperature fluctuations, such as a sudden decrease in temperature during the heating phase. Author Cryogenic Temperature; Cryopumping; Insulation; Reusable Launch Vehicles; Manufacturing

20040000905 NASA Stennis Space Center, Bay Saint Louis, MS, USA Improved Testing Capability and Adaptability Through the Use of Wireless Sensors Solano, Wanda M.; April 14, 2003; In English; 2003 Propulsion Measurement Sensor Development Workshop, 13-15 May 2003, Huntsville, AL, USA; Original contains color illustrations Report No.(s): SE-2003-04-00020-SSC; No Copyright; Avail: CASI; A02, Hardcopy From the ﬁrst Saturn V rocket booster (S-II-T) testing in 1966 and the routine Space Shuttle Main Engine (SSME) testing beginning in 1975, to more recent test programs such as the X-33 Aerospike Engine, the Integrated Powerhead Development (IPD) program, and the Hybrid Sounding Rocket (HYSR), Stennis Space Center (SSC) continues to be a premier location for conducting large-scale testing. Central to each test program is the capability for sensor systems to deliver reliable measurements and high quality data, while also providing a means to monitor the test stand area to the highest degree of safety and sustainability. Sensor wiring is routed along piping and through cable trenches, making its way from the engine test area, through the test stand area and to the signal conditioning building before ﬁnal transfer to the test control center. When sensor requirements lie outside the reach of the routine sensor cable routing, the use of wireless sensor networks becomes particularly attractive due to their versatility and ease of installation. As part of an on-going effort to enhance the testing capabilities of Stennis Space Center, the Test Technology and Development group has found numerous applications for its sensor-adaptable wireless sensor suite. While not intended for critical engine measurements or control loops, in-house hardware and software development of the sensor suite can provide improved testing capability for a range of applications including the safety monitoring of propellant storage barrels and as an experimental test-bed for embedded health monitoring paradigms. Author Engine Tests; Engine Monitoring Instruments; Instrument Packages

20040000874 Lockheed Martin Space Operations, Bay Saint Louis, MS, USA NASA Stennis Space Center V and V Capabilities Overview Ryan, Robert; October 22, 2001; In English; MAPPS-ASPRS Conference: Measuring the Earth-Digital Elevation Technologies and Applications, 31 Oct. - 2 Nov. 2001, Saint Petersburg, FL, USA Contract(s)/Grant(s): NAS13-650 Report No.(s): SE-2001-10-00062-SSC; No Copyright; Avail: CASI; A03, Hardcopy

2 The objective are provide independent field measurement ground capability for satellite and aircraft mounted in-flight sensor validation. NASA Stennis is considering expanding its Verification and Validation (V&V) range to support LIDAR and digital camera characterization. Soliciting input for requirements and suggestion. Derived from text Digital Cameras; Optical Radar; Aircraft Instruments

20040000627 NASA Glenn Research Center, Cleveland, OH, USA Update on the Development and Capabilities of Unique Structural Seal Test Rigs DeMange, Jeffrey J.; Dunlap, Patrick H., Jr.; Steinetz, Bruce M.; Breen, Daniel P.; Robbie, Malcolm G.; 2002 NASA Seal/Secondary Air System Workshop; September 2003; Volume 1; In English; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy High temperature structural seals are necessary in many aerospace and aeronautical applications to minimize any detrimental effects originating from undesired leakage. The NASA Glenn Research Center has been and continues to be a pioneer in the development and evaluation of these types of seals. The current focus for the development of structural seals is for the 3rd Generation Reusable Launch Vehicle (RLV), which is scheduled to replace the current space shuttle system by 2025. Speciﬁc areas of development under this program include seals for propulsion systems (such as the hypersonic air-breathing ISTAR engine concept based upon Rocket Based Combined Cycle technology) and control surface seals for spacecraft including the autonomous rescue X-38 Crew Return Vehicle and the X-37 Space Maneuver Vehicle. Derived from text X-38 Crew Return Vehicle; X-37 Vehicle; Sealing; Control Surfaces; Air Breathing Engines; Reusable Launch Vehicles

20040000626 NASA Glenn Research Center, Cleveland, OH, USA Third Generation RLV Structural Seal Development Programs at NASA GRC Dunlap, Patrick H., Jr.; Steinetz, Bruce M.; DeMange, Jeffrey J.; 2002 NASA Seal/Secondary Air System Workshop; September 2003; Volume 1; In English; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy NASA is currently developing technologies for the 3rd Generation Reusable Launch Vehicle (RLV) that is being designed to enter service around the year 2025. In particular, NASA’s Glenn Research Center (GRC) is working on advanced high temperature structural seal designs including propulsion system and control surface seals. Propulsion system seals are required along the edges of movable panels in advanced engines, while control surface seals seal the edges and hinge lines of movables flaps and elevons on the vehicle. The overall goal is to develop reusable, resilient seals capable of operating at temperatures up to 2000 F. High temperature seal preloading devices (e.g., springs) are also being evaluated as a means of improving seal resiliency. In order to evaluate existing and potential new seal designs, GRC has designed and is installing several new test rigs capable of simulating the types of conditions that the seals would endure during service including temperatures, pressures, and scrubbing. Two new rigs, the hot compression test rig and the hot scrub test rig, will be used to perform seal compression and scrub tests for many cycles at temperatures up to 3000 F. Another new test rig allows simultaneous flow and scrub tests to be performed on the seals at room temperature to evaluate how the flow blocking performance of the seals varies as they accumulate damage during scrubbing. This presentation will give an overview of these advanced seal development efforts. Author Reusable Launch Vehicles; Structural Design; Sealing; Control Surfaces

20030112990 NASA Marshall Space Flight Center, Huntsville, AL, USA Using CFD as Rocket Injector Design Tool: Recent Progress at Marshall Space Flight Center Tucker, Kevin; West, Jeff; Williams, Robert; Lin, Jeff; Rocker, Marvin; Canabal, Francisco; Robles, Bryan; Garcia, Robert; Chenoweth, James; [2003]; In English; 5th International Symposium on Liquid Space Propulsion Long Life Combustion Devices Technology, 27-30 Oct. 2003, Chatanooga, TN, USA; Copyright; Avail: CASI; A03, Hardcopy The choice of tools used for injector design is in a transitional phase between exclusive reliance on the empirically based correlations and extensive use of computational ﬂuid dynamics (CFD). The Next Generation Launch Technology (NGLT) Program goals emphasizing lower costs and increased reliability have produced a need to enable CFD as an injector design tool in a shorter time frame. This is the primary objective of the Staged Combustor Injector Technology Task currently under way at Marshall Space Flight Center (MSFC). The documentation of this effort begins with a very brief status of current injector design tools. MSFC’s vision for use of CFD as a tool for combustion devices design is stated and discussed with emphasis on the injector. The concept of the Simulation Readiness Level (SRL), comprised of solution ﬁdelity, robustness and accuracy, is introduced and discussed. This quantitative measurement is used to establish the gap between the current state of

3 demonstrated capability and that necessary for regular use in the design process. MSFC’s view of the validation process is presented and issues associated with obtaining the necessary data are noted and discussed. Three current experimental efforts aimed at generating validation data are presented. The importance of uncertainty analysis to understand the data quality is also demonstrated. First, a brief status of current injector design tools is provided as context for the current effort. Next, the MSFC vision for using CFD as an injector design tool is stated. A generic CFD-based injector design methodology is also outlined and brieﬂy discussed. Three areas where MSFC is using injector CFD analyses for program support will be discussed. These include the Integrated Powerhead Development (IPD) engine which uses hydrogen and oxygen propellants in a full ﬂow staged combustion (FFSC) cycle and the TR-107 and the RS84 engine both of which use RP-1 and oxygen in an ORSC cycle. Finally, an attempt is made to objectively summarize what progress has been made at MSFC in enabling CFD as an injector design tool. Author Computational Fluid Dynamics; Rocket Engine Design; Research Facilities; Injectors

20030112784 NASA Glenn Research Center, Cleveland, OH, USA Computations of Internal and External Axisymmetric Nozzle Aerodynamics at Transonic Speeds Dalbello, Teryn; Georgiadis, Nicholas; Yoder, Dennis; Keith, Theo; November 2003; In English; 39th Combustion/27th Airbreathing Propulsion/21st Propulsion Systems Hazards/Third Modeling and Simulation Joint Subcommittee Meeting, 1-5 Dec. 2003, Colorado Springs, CO, USA; Original contains color and black and white illustrations Contract(s)/Grant(s): WBS 22-708-73-22 Report No.(s): NASA/TM-2003-212731; E-14256; NAS 1.15:212731; Copyright; Avail: CASI; A03, Hardcopy Computational Fluid Dynamics (CFD) analyses of axisymmetric circular-arc boattail nozzles have been completed in support of NASA’s Next Generation Launch Technology Program to investigate the effects of high-speed nozzle geometries on the nozzle internal flow and the surrounding boattail regions. These computations span the very difficult transonic flight regime, with shock-induced separations and strong adverse pressure gradients. External afterbody and internal nozzle pressure distributions computed with the Wind code are compared with experimental data. A range of turbulence models were examined in Wind, including an Explicit Algebraic Stress model (EASM). Computations on two nozzle geometries have been completed at freestream Mach numbers ranging from 0.6 to 0.9, driven by nozzle pressure ratios (NPR) ranging from 2.9 to 5. Results obtained on converging-only geometry indicate reasonable agreement to experimental data, with the EASM and Shear Stress Transport (SST) turbulence models providing the best agreement. Calculations completed on a converging- diverging geometry involving large-scale internal flow separation did not converge to a true steady-state solution when run with variable timestepping (steady-state). Calculations obtained using constant timestepping (time-accurate) indicate less variations in flow properties compared with steady-state solutions. This failure to converge to a steady-state solution was found to be the result of difficulties in using variable time-stepping with large-scale separations present in the flow. Nevertheless, time-averaged boattail surface pressure coefficient and internal nozzle pressures show fairly good agreement with experimental data. The SST turbulence model demonstrates the best over-all agreement with experimental data. Author Convergent-Divergent Nozzles; Nozzle Flow; Nozzle Geometry; Boattails

20030112231 NASA Marshall Space Flight Center, Huntsville, AL, USA Orbital Space Plane Cost Credibility Creech, Steve; January 9, 2003; In English; 54th International Astronautical Congress, 29 Sep. - 3 Oct. 2003, Bremen, Germany; No Copyright; Avail: Other Sources; Abstract Only NASA’s largest new start development program is the Orbital Space Plane (OSP) Program. The program is currently in the formulation stage. One of the critical issues to be resolved, prior to initiating full-scale development, is establishing cost credibility of NASA s budget estimates for development, production, and operations of the OSP. This paper will discuss the processes, tools, and methodologies that NASA, along with its industry partners, are implementing to assure cost credibility for the OSP program. Results of benchmarking of current tools and the development of new cost estimating capabilities and approaches will be discussed. Author Cost Estimates; Aerospace Planes

20030111907 NASA Marshall Space Flight Center, Huntsville, AL, USA NASA’s Next Generation Launch Technology Program - Strategy and Plans Hueter, Uwe; [2003]; In English; 54th International Astronautical Congress, 29 Sep. - 3 Oct. 2003, Bremen, Germany Report No.(s): 1AC-03-V.5.01; No Copyright; Avail: CASI; A02, Hardcopy

4 The National Aeronautics and Space Administration established a new program office, Next Generation Launch Technology (NGLT) Program Office, last year to pursue technologies for future space launch systems. NGLT will fund research in key technology areas such as propulsion, launch vehicles, operations and system analyses. NGLT is part of NASA s Integrated Space Technology Plan. The NGLT Program is sponsored by NASA s Office of Aerospace Technology and is part of the Space Launch Initiative theme that includes both NGLT and Orbital Space Plane. NGLT will focus on technology development to increase safety and reliability and reduce overall costs associated with building, ﬂying and maintaining the nation s next-generations of space launch vehicles. These investments will be guided by systems engineering and analysis with a focus on the needs of National customers. Author Aerospace Engineering; Aerospace Planes; Spacecraft Launching

20030111798 NASA Marshall Space Flight Center, Huntsville, AL, USA Orbital Space Plane Program Status Dumbacher, Daniel L.; [2003]; In English; 54th International Astronautical Congress, 29 Sep. - 3 Oct. 2003, Bremen, Germany; No Copyright; Avail: Other Sources; Abstract Only The Orbital Space Plane Program is an integral part of NASA’s Integrated Space Transportation Program (ISTP). The ISTP consists of three major programs: Space Shuttle, Orbital Space Plane, and Next Generation Launch Technology. The Orbital Space Plane (OSP) Program will develop a new Crew Transfer Vehicle (CTV) with multipurpose utility for the Agency. The CTV will complement and back up the Space Shuttle by taking crews to and from the International Space Station (ISS), as well as enable a transition path to future reusable launch vehicle systems. In the CTV development cycle, around 2010 it will be used as a Crew Return Vehicle (CRV). The OSP will be launched on an Evolved Expendable Launch Vehicle (EELV). NASA is in the process of establishing Level 1 Requirements and initiating concept studies. Ongoing ﬂight demonstrators will continue, while new ﬂight demonstrator projects will begin. The OSP Program contains two elements: (1) Technology and Demonstrations, and (2) Design, Development, and Production. The OSP Design, Development, and Production element will enter the Formulation Phase in FY03. Per NASA Procedures and Guidelines 7120.5B, the Formulation Phase will be utilized to establish the Program schedule and budget plans. Current budget planning is based on Phase A concept studies being conducted in FY03 and FY04, preliminary design activities conducted in FY04 and FY05, and a Preliminary Design Review in FY05. An OSP full-scale development decision will be made in FY05. At that point, a conclusion to proceed will result in the OSP Program transitioning from the Formulation Phase to the Development Phase. Author Aerospace Planes; NASA Space Programs; Launch Vehicles; Space Shuttles; Space Shuttle Orbiters; Reusable Launch Vehicles

20030111797 Lockheed Martin Space Mission Systems and Services, Littleton, CO, USA Orbital Space Plane (OSP) Program McKenzie, Patrick M.; September 19, 2003; In English; 54th International Astronautical Congress, 29 Sep. - 3 Oct. 2003, Bremen, Germany Contract(s)/Grant(s): NAS8-01098; No Copyright; Avail: CASI; A02, Hardcopy Lockheed Martin has been an active participant in NASA’s Space Launch Initiative (SLI) programs over the past several years. SLI, part of NASA’s Integrated Space Transportation Plan (ISTP), was restructured in November of 2002 to focus the overall theme of safer, more afford-able space transportation along two paths - the Orbital Space Plane Program and the Next Generation Launch Technology programs. The Orbital Space Plane Program has the goal of providing rescue capability from the International Space Station by 2008 and transfer capability for crew (and limited cargo) by 2012. The Next Generation Launch Technology program is combining research and development efforts from the 2nd Generation Reusable Launch Vehicle (2GRLV) program with cutting-edge, advanced space transportation programs (previously designated 3rd Generation) into one program aimed at enabling safe, reliable, cost-effective reusable launch systems by the middle of the next decade. Lockheed Martin is one of three prime contractors working to bring Orbital Space Plane system concepts to a system deﬁnition level of maturity by December of 2003. This paper and presentation will update the international community on the progress of the’ OSP program, from an industry perspective, and provide insights into Lockheed Martin’s role in enabling the vision of a safer, more affordable means of taking people to and from space. Author Space Transportation; International Space Station; Systems Analysis; Aerospace Planes

5 20030111778 NASA Marshall Space Flight Center, Huntsville, AL, USA Orbital Space Plane Program Flight Demonstrators Status Turner, Susan G.; [2003]; In English; 54th International Astronautical Congress, 29 Sep. - 3 Oct. 2003, Bremen, Germany; No Copyright; Avail: Other Sources; Abstract Only Under the Orbital Space Plane Program, NASA is currently pursuing the maturation of technologies via three ﬂight demonstrators - DART (Demonstration of Autonomous Rendezvous Technology), X-37, and PAD (Pad Abort Demonstrator). Flight demonstrators provide the opportunity to test key technologies in their actual working environment. These ﬂight demonstrators are required at this stage to mature technologies needed to support full-scale development design of a future competitively selected Orbital Space. Derived from text Flight Tests; Flight Simulators; Autonomy; Aerospace Planes

20030108286 NASA Glenn Research Center, Cleveland, OH, USA Durability of Membrane Electrode Assemblies (MEAs) in PEM Fuel Cells Operated on Pure Hydrogen and Oxygen Stanic, Vesna; Braun, James; Hoberecht, Mark; [2003]; In English; International Energy Conversion Engineering Conference, 17-21 Aug. 2003, Portsmouth, VA, USA Contract(s)/Grant(s): NAS3-02093; 22-721-22-09; No Copyright; Avail: CASI; A02, Hardcopy Proton exchange membrane (PEM) fuel cells are energy sources that have the potential to replace alkaline fuel cells for space programs. Broad power ranges, high peak-to-nominal power capabilities, low maintenance costs, and the promise of increased life are the major advantages of PEM technology in comparison to alkaline technology. The probability of PEM fuel cells replacing alkaline fuel cells for space applications will increase if the promise of increased life is verified by achieving a minimum of 10,000 hours of operating life. Durability plays an important role in the process of evaluation and selection of MEAs for Teledyne s Phase I contract with the NASA Glenn Research Center entitled Proton Exchange Membrane Fuel cell (PEMFC) Power Plant Technology Development for 2nd Generation Reusable Launch Vehicles (RLVs). For this contract, MEAs that are typically used for H2/air operation were selected as potential candidates for H2/O2 PEM fuel cells because their catalysts have properties suitable for O2 operation. They were purchased from several well-established MEA manufacturers who are world leaders in the manufacturing of diverse products and have committed extensive resources in an attempt to develop and fully commercialize MEA technology. A total of twelve MEAs used in H2/air operation were initially identified from these manufacturers. Based on the manufacturers specifications, nine of these were selected for evaluation. Since 10,000 hours is almost equivalent to 14 months, it was not possible to perform continuous testing with each MEA selected during Phase I of the contract. Because of the lack of time, a screening test on each MEA was performed for 400 hours under accelerated test conditions. The major criterion for an MEA pass or fail of the screening test was the gas crossover rate. If the gas crossover rate was higher than the membrane intrinsic permeability after 400 hours of testing, it was considered that the MEA had failed the test. Three types of MEAs out of the nine total membranes failed the test. The evaluation results showed that fuel cell operating conditions (current, pressure, stoichiometric flow rates) were the parameters that influenced the durability of MEAs. In addition, the durability test results indicated that the type of membrane was also an important parameter for MEA durability. At accelerated test conditions, the MEAs with casted membranes failed during the 400 hour test. However, the MEAs prepared from the casted membrane with support as well as extruded membranes, both passed the 400h durability test at accelerated operating test conditions. As a result of the MEA accelerated durability tests, four MEAs were selected for further endurance testing. These tests are being carried out with four-cell stacks under nominal fuel cell operating conditions. Author Durability; Electrodes; Membranes; Protons; Manufacturing; Hydrogen Oxygen Fuel Cells

20030108283 NASA Glenn Research Center, Cleveland, OH, USA Performance Evaluation of Electrochem’s PEM Fuel Cell Power Plant for NASA’s 2nd Generation Reusable Launch Vehicle Kimble, Michael C.; Hoberecht, Mark; [2003]; In English; International Energy Conversion Engineering Conference, 17-21 Aug. 2003, Portsmouth, VA, USA Contract(s)/Grant(s): NAS3-02092; 22-721-22-09; No Copyright; Avail: CASI; A02, Hardcopy NASA’s Next Generation Launch Technology (NGLT) program is being developed to meet national needs for civil and commercial space access with goals of reducing the launch costs, increasing the reliability, and reducing the maintenance and operating costs. To this end, NASA is considering an all- electric capability for NGLT vehicles requiring advanced electrical power generation technology at a nominal 20 kW level with peak power capabilities six times the nominal power. The proton exchange membrane (PEM) fuel cell has been identiﬁed as a viable candidate to supply this electrical power; however, several

6 technology aspects need to be assessed. Electrochem, Inc., under contract to NASA, has developed a breadboard power generator to address these technical issues with the goal of maximizing the system reliability while minimizing the cost and system complexity. This breadboard generator operates with dry hydrogen and oxygen gas using eductors to recirculate the gases eliminating gas humidification and blowers from the system. Except for a coolant pump, the system design incorporates passive components allowing the fuel cell to readily follow a duty cycle profile and that may operate at high 6:1 peak power levels for 30 second durations. Performance data of the fuel cell stack along with system performance is presented to highlight the benefits of the fuel cell stack design and system design for NGLT vehicles. Author Reusable Launch Vehicles; Protons; Membranes; Fuel Cell Power Plants; Performance Prediction; NASA Space Programs; Fuel Cells

20030107869 NASA Stennis Space Center, Bay Saint Louis, MS, USA Engineering the Future of Full-Scale Propulsion Testing Ryan, H. M.; Solano, W.; Holland, R.; Saint Cyr, W.; Rahman, S.; [2001]; In English; 39th AIAA Aerospace Sciences Meeting and Exhibit, 8-11 Jan. 2001, Reno, NV, USA Report No.(s): AIAA Paper 2001-0746; SE-2001-01-00001-SSC; Copyright; Avail: CASI; A03, Hardcopy Year 2000 has been an active one for rocket propulsion testing at the NASA John C. Stennis Space Center. This paper highlights several major test facilites for large-scale propulsion devices, and summarizes the varied nature of the recent test projects conducted at the Stennis Space Center (SSC) such as the X-33 Aerospike Engine, Ultra Low Cost Engine (ULCE) thrust chamber program, and the Hybrid Sounding Rocket (HYSR) program. Further, an overview of relevant engineering capabilities and technology challenges in conducting full-scale propulsion testing are outlined. Author Engine Tests; Full Scale Tests; Propellant Tests; Test Facilities

20030106071 Lockheed Martin Space Systems Co., Huntsville, AL, USA Orbital Space Plane (OSP) Program at Lockheed Martin Ford, Robert; September 19, 2003; In English; Space 2003, 24 Sep.2003, Long Beach, CA, USA Contract(s)/Grant(s): NAS8-01098; No Copyright; Avail: CASI; A02, Hardcopy Lockheed Martin has been an active participant in NASA’s Space Launch Initiative (SLI) programs over the past several years. SLI, part of NASA’s Integrated Space Transportation Plan (ISTP), was restructured in November 2002 to focus the overall theme of safer, more affordable space transportation along two paths the Orbital Space Plane (OSP) and the Next Generation Launch Technology programs. The Orbital Space Plane program has the goal of providing rescue capability from the International Space Station by 2008 or earlier and transfer capability for crew (and contingency cargo) by 2012. The Next Generation Launch Technology program is combining research and development efforts from the 2d Generation Reusable Launch Vehicle (2GRLV) program with cutting-edge, advanced space transportation programs (previously designated 31d Generation) into one program aimed at enabling safe, reliable, cost-effective reusable launch systems by the middle of the next decade. Lockheed Martin is one of three prime contractors working to bring Orbital Space Plane system concepts to a system design level of maturity by December 2003. This paper and presentation will update the aerospace community on the progress of the OSP program, from an industry perspective, and provide insights into Lockheed Martin’s role in enabling the vision of a safer, more affordable means of taking people to and from space. Author Aerospace Planes; Space Transportation; Reusable Launch Vehicles; NASA Space Programs; Spacecraft Orbits; Spacecraft Conﬁgurations

20030093628 Vanderbilt Univ., TN, USA Impedance-Based Structural Health Monitoring for Composite Laminates at Cryogenic Environments Tseng, Kevin; The 2002 NASA Faculty Fellowship Program Research Reports; April 2003; In English Contract(s)/Grant(s): NAG8-1859; No Copyright; Avail: CASI; C01, CD-ROM; A01, Hardcopy One of the important ways of increasing the payload in a reusable launch vehicle (RLV) is to replace heavy metallic materials by lightweight composite laminates. Among various parts and systems of the RLV, this project focuses on tanks containing cryogenic fuel. Historically, aluminum alloys have been used as the materials to construct fuel tanks for launch vehicles. To replace aluminum alloys with composite laminates or honeycomb materials, engineers have to make sure that the composites are free of defects before, during, and after launch. In addition to robust design and manufacturing procedures, the

7 performance of the composite structures needs to be monitored constantly.In recent years, the impedance-based health monitoring technique has shown its promise in many applications. This technique makes use of the special properties of smart piezoelectric materials to identify the change of material properties due to the nucleation and progression of damage. The piezoceramic patch serves as a sensor and an actuator simultaneously. The piezoelectric patch is bonded onto an existing structure or embedded into a new structure and electrically excited at high frequencies. The signature (impedance or admittance) is extracted as a function of the exciting frequency and is compared with the baseline signature of the healthy state. The damage is quantiﬁed using root mean square deviation (RMSD) in the impedance signatures with respect to the baseline signature. A major advantage of this technique is that the procedure is nondestructive in nature and does not perturb the properties and performance of the materials and structures. This project aims at applying the impedance-based nondestructive testing technique to the damage identiﬁcation of composite laminates at cryogenic temperature. Author Laminates; Reusable Launch Vehicles; Piezoelectric Actuators

20030093602 Rose-Hulman Inst. of Tech., Terre Haute, IN, USA Fuzzy Logic Trajectory Design and Guidance for Terminal Area Energy Management Burchett, Bradley; The 2002 NASA Faculty Fellowship Program Research Reports; April 2003; In English Contract(s)/Grant(s): NAG8-1859; No Copyright; Avail: CASI; C01, CD-ROM; A01, Hardcopy The second generation reusable launch vehicle will leverage many new technologies to make flight to low earth orbit safer and more cost effective. One important capability will be completely autonomous flight during reentry and landing, thus making it unnecessary to man the vehicle for cargo missions with stringent weight constraints. Implementation of sophisticated new guidance and control methods will enable the vehicle to return to earth under less than favorable conditions. The return to earth consists of three phases--Entry, Terminal Area Energy Management (TAEM), and Approach and Landing. The Space Shuttle is programmed to fly all three phases of flight automatically, and under normal circumstances the astronaut-pilot takes manual control only during the Approach and Landing phase. The automatic control algorithms used in the Shuttle for TAEM and Approach and Landing have been developed over the past 30 years. They are computationally efficient, and based on careful study of the spacecraft’s flight dynamics, and heuristic reasoning. The gliding return trajectory is planned prior to the mission, and only minor adjustments are made during flight for perturbations in the vehicle energy state. With the advent of the X-33 and X-34 technology demonstration vehicles, several authors investigated implementing advanced control methods to provide autonomous real-time design of gliding return trajectories thus enhancing the ability of the vehicle to adjust to unusual energy states. The bulk of work published to date deals primarily with the approach and landing phase of flight where changes in heading angle are small, and range to the runway is monotonically decreasing. These benign flight conditions allow for model simplification and fairly straightforward optimization. This project focuses on the TAEM phase of flight where mathematically precise methods have produced limited results. Fuzzy Logic methods are used to make onboard autonomous gliding return trajectory design robust to a wider energy envelope, and the possibility of control surface failures, thus increasing the flexibility of unmanned gliding recovery and landing. Derived from text Fuzzy Systems; Logic Design; Trajectories; Terminal Area Energy Management; Automatic Control

20030093594 Embry-Riddle Aeronautical Univ., USA System Identification of X-33 Neural Network Aggarwal, Shiv; The 2002 NASA Faculty Fellowship Program Research Reports; April 2003; In English Contract(s)/Grant(s): NAG8-1859; No Copyright; Avail: CASI; C01, CD-ROM; A01, Hardcopy Modern flight control research has improved spacecraft survivability as its goal. To this end we need to have a failure detection system on board. In case the spacecraft is performing imperfectly, reconfiguration of control is needed. For that purpose we need to have parameter identification of spacecraft dynamics. Parameter identification of a system is called system identification. We treat the system as a black box which receives some inputs that lead to some outputs. The question is: what kind of parameters for a particular black box can correlate the observed inputs and outputs? Can these parameters help us to predict the outputs for a new given set of inputs? This is the basic problem of system identification. The X33 was supposed to have the onboard capability of evaluating the current performance and if needed to take the corrective measures to adapt to desired performance. The X33 is comprised of both rocket and aircraft vehicle design characteristics and requires, in general, analytical methods for evaluating its flight performance. Its flight consists of four phases: ascent, transition, entry and TAEM (Terminal Area Energy Management). It spends about 200 seconds in ascent phase, reaching an altitude of about 180,000 feet and a speed of about 10 to 15 Mach. During the transition phase which lasts only about 30 seconds, its altitude may increase to about 190,000 feet but its speed is reduced to about 9 Mach. At the beginning of this phase, the Main Engine

8 is Cut Off (MECO) and the control is reconfigured with the help of aerosurfaces (four elevons, two flaps and two rudders) and reaction control system (RCS). The entry phase brings down the altitude of X33 to about 90,000 feet and its speed to about Mach 3. It spends about 250 seconds in this phase. Main engine is still cut off and the vehicle is controlled by complex maneuvers of aerosurfaces. The last phase TAEM lasts for about 450 seconds and the altitude and speed, both are reduced to zero. The present attempt, as a start, focuses only on the entry phase. Since the main engine remains cut off in this phase, there is no thrust acting on the system. This considerably simplifies the equations of motion. We introduce another simplification by assuming the system to be linear after some non-linearities are removed analytically from our consideration. Under these assumptions, the problem could be solved by Classical Statistics by employing the least sum of squares approach. Instead we chose to use the Neural Network method. This method has many advantages. It is modern, more efficient, can be adapted to work even when the assumptions are diluted. In fact, Neural Networks try to model the human brain and are capable of pattern recognition. Derived from text Aircraft Design; Design Analysis; Flight Characteristics; Flight Control; Neural Nets; X-33 Reusable Launch Vehicle

20030093586 Missouri Univ., Rolla, MO, USA Adaptive Critic Neural Network-Based Terminal Area Energy Management and Approach and Landing Guidance Grantham, Katie; The 2002 NASA Faculty Fellowship Program Research Reports; April 2003; In English Contract(s)/Grant(s): NAG8-1859; No Copyright; Avail: CASI; C01, CD-ROM; A02, Hardcopy Reusable Launch Vehicles (RLVs) have different mission requirements than the Space Shuttle, which is used for benchmark guidance design. Therefore, alternative Terminal Area Energy Management (TAEM) and Approach and Landing (A/L) Guidance schemes can be examined in the interest of cost reduction. A neural network based solution for a ﬁnite horizon trajectory optimization problem is presented in this paper. In this approach the optimal trajectory of the vehicle is produced by adaptive critic based neural networks, which were trained off-line to maintain a gradual glideslope. Derived from text Aircraft Guidance; Aircraft Landing; Neural Nets; Reusable Launch Vehicles; Terminal Area Energy Management

20030093585 Alabama Univ., Huntsville, AL, USA An Improved RLV Stability Analysis Via a Continuation Approach Friedman, Mark; The 2002 NASA Faculty Fellowship Program Research Reports; April 2003; In English Contract(s)/Grant(s): NAG8-1859; No Copyright; Avail: CASI; C01, CD-ROM; A01, Hardcopy The continuation of Invariant Subspaces (CIS) algorithm produces a smooth orthogonal similarity transformation to block triangular form of A(s). A practical CIS algorithm and stability analysis of the X-33 are described. Derived from text Algorithms; Invariant Imbeddings; X-33 Reusable Launch Vehicle; Stability Tests; Continuity (Mathematics); Branching (Mathematics)

20030093576 Alabama Univ., Tuscaloosa, AL, USA, Alabama Univ., Huntsville, AL, USA The 2002 NASA Faculty Fellowship Program Research Reports Bland, J., Compiler; April 2003; In English; Original contains color and black and white illustrations Contract(s)/Grant(s): NAG8-1859 Report No.(s): NASA/CR-2003-212397; M-1074; NAS 1.26:212397; No Copyright; Avail: CASI; C01, CD-ROM Contents include the following: System Identification of X-33. Neural Network Advanced Ceramic Technology for Space Applications at NASA MSFC. Developing a MATLAB-Based Tool for Visualization and Transformation. Subsurface Stress Fields in Single Crystal (Anisotropic). Contacts Our Space Future: A Challenge to the Conceptual Artist Concept Art for Presentation and Education. Identification and Characterization of Extremophile Microorganisms. Significant to Astrobiology. Mathematical Investigation of Gamma Ray and Neutron. Absorption Grid Patterns for Homeland Defense-Related Fourier Imaging Systems. The Potential of Microwave Radiation for Processing Martian Soil. Fuzzy Logic Trajectory Design and Guidance for Terminal Area. Derived from text Active Control; Combustion Chambers; Cylindrical Bodies; Data Processing; Design to Cost

9 20030079974 NASA Langley Research Center, Hampton, VA, USA Hypervelocity Impact Test Results for a Metallic Thermal Protection System Karr, Katherine L.; Poteet, Carl C.; Blosser, Max L.; August 2003; In English; Original contains color and black and white illustrations Contract(s)/Grant(s): 721-21-87-07 Report No.(s): NASA/TM-2003-212440; L-18192; NAS 1.15:212440; No Copyright; Avail: CASI; A05, Hardcopy Hypervelocity impact tests have been performed on specimens representing metallic thermal protection systems (TPS) developed at NASA Langley Research Center for use on next-generation reusable launch vehicles (RLV). The majority of the specimens tested consists of a foil gauge exterior honeycomb panel, composed of either Inconel 617 or Ti-6Al-4V, backed with 2.0 in. of fibrous insulation and a final Ti-6Al-4V foil layer. Other tested specimens include titanium multi-wall sandwich coupons as well as TPS using a second honeycomb sandwich in place of the foil backing. Hypervelocity impact tests were performed at the NASA Marshall Space Flight Center Orbital Debris Simulation Facility. An improved test fixture was designed and fabricated to hold specimens firmly in place during impact. Projectile diameter, honeycomb sandwich material, honeycomb sandwich facesheet thickness, and honeycomb core cell size were examined to determine the influence of TPS configuration on the level of protection provided to the substructure (crew, cabin, fuel tank, etc.) against micrometeoroid or orbit debris impacts. Pictures and descriptions of the damage to each specimen are included. Author Hypervelocity Impact; Impact Tests; Thermal Protection; Fabrication; Intermetallics

20030077847 NASA Marshall Space Flight Center, Huntsville, AL, USA Experimental Performance Evaluation of a Supersonic Turbine for Rocket Engine Applications Snellgrove, Lauren M.; Griffin, Lisa W.; Sieja, James P.; Huber, Frank W.; July 08, 2003; In English; 39th AIAA/ASME/ SAE/ASEE Joint Propulsion Conference and Exhibit, 20-23 Jul. 2003, Huntsville, AL, USA Report No.(s): AIAA Paper 2003-4918; Copyright; Avail: CASI; A03, Hardcopy In order to mitigate the risk of rocket propulsion development, efficient, accurate, detailed fluid dynamics analysis and testing of the turbomachinery is necessary. To support this requirement, a task was developed at NASA Marshall Space Flight Center (MSFC) to improve turbine aerodynamic performance through the application of advanced design and analysis tools. These tools were applied to optimize a supersonic turbine design suitable for a reusable launch vehicle (RLV). The hot gas path and blading were redesigned-to obtain an increased efficiency. The goal of the demonstration was to increase the total-to- static efficiency of the turbine by eight points over the baseline design. A sub-scale, cold flow test article modeling the final optimized turbine was designed, manufactured, and tested in air at MSFC s Turbine Airflow Facility. Extensive on- and off- design point performance data, steady-state data, and unsteady blade loading data were collected during testing. Author Supersonic Turbines; Reusable Launch Vehicles; Turbine Blades; Aerodynamic Characteristics

20030075777 NASA Ames Research Center, Moffett Field, CA, USA Accelerating the Design of Space Vehicles Laufenberg, Larry, Editor; Infusion; August 08, 2003; In English Contract(s)/Grant(s): UPN 704-00-00; No Copyright; Avail: CASI; A01, Hardcopy One of NASA’s key goals is to increase the safety and reduce the cost of space transportation. Thus, a key element of NASA’s new Integrated Space Transportation Plan is to develop new propulsion, structures, and operations for future generations of reusable launch vehicles (RLVs). As part of this effort to develop the next RLV, the ClCT Program’s Computing, Networking, and Information Systems (CNIS) Project is developing and demonstrating collaborative software technologies that use the collective power of the NASA Grid to accelerate spacecraft design. One of these technologies, called AeroDB, automates the execution and monitoring of computational fluid dynamics (CFD) parameter studies on the NASA Grid. About the NASA Grid The NASA Grid, or Information Power Grid,. is being developed to leverage the distributed resources of NASA’s many computers. instruments, simulators, and data storage systems. The goal is to use these combined resources to sdve difficult NASA challenges, such as iimulating the entire flight of a space vehicle from ascent to descent.To realize the vision of the NASA Grid, the CNIS Project is developing the software framework and protocols for building domain-specific environments and interfaces, new Grid services based on emerging industry standards, and advanced networking and computing testbeds to support new Grid-based applications such as AeroDB. Author Reusable Launch Vehicles; Computational Fluid Dynamics

10 20030071272 Aerojet-General Corp., Sacramento, CA, USA Non-Toxic Dual Thrust Reaction Control Engine Development for On-Orbit APS Applications Robinson, Philip J.; Veith, Eric M.; [2003]; In English; 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 20-23 Jul. 2003, Huntsville, AL, USA; No Copyright; Avail: CASI; A03, Hardcopy A non-toxic dual thrust proof-of-concept demonstration engine was successfully tested at the Aerojet Sacramento facility under a technology contract sponsored by the National Aeronautics and Space Administration’s (NASA) Marshall Space Flight Center (MSFC). The goals of the NASA MSFC contract (NAS8-01109) were to develop and expand the technical maturity of a non-toxic, on-orbit auxiliary propulsion system (APS) thruster under the Next Generation Launch Technology (NGLT) program. The demonstration engine utilized the existing Kistler K-1 870 lbf LOX/Ethanol orbital maneuvering engine (Om)coupled with some special test equipment (STE) that enabled engine operation at 870 lbf in the primary mode and 25 lbf in the vernier mode. Ambient testing in primary mode varied mixture ratio (MR) from 1.28 to 1.71 and chamber pressure (P(c) from 110 to 181 psia, and evaluated electrical pulse widths (EPW) of 0.080, 0.100 and 0.250 seconds. Altitude testing in vernier mode explored igniter and thruster pulsing characteristics, long duration steady state operation (greater than 420 sec) and the impact of varying the percent fuel ﬁlm cooling on vernier performance and chamber thermal response at low PC (4 psia). Data produced from the testing provided calibration of the performance and thermal models used in the design of the next version of the dual thrust Reaction Control Engine (RCE). Author Auxiliary Propulsion; Oxygen-Hydrocarbon Rocket Engines

20030071238 Ohio Aerospace Inst., Brook Park, OH, USA Structural Seal Development DeMange, Jeff; November 29, 2002; In English Contract(s)/Grant(s): NCC3-879 Report No.(s): OAI-02-011; No Copyright; Avail: CASI; A03, Hardcopy NASA currently has aggressive goals for the development of the next generations of Reusable Launch Vehicles (RLV). Many of the goals center around signiﬁcantly reducing the time and cost of deployment (including payload) for these spacecraft, as well as substantially increasing the safety of the crew and hardware. A core component in the design of these next generation spacecraft is the propulsion system. Many concepts are currently being evaluated including Rocket Based Combined Cycle (RBCC) and Turbine Based Combined Cycle (TBCC) engine technologies, shown. Author Cost Reduction; Time Lag; Payloads; Rocket-Based Combined-Cycle Engines

20030068866 QSS Group, Inc., Cleveland, OH, USA Oxidation-Reduction Resistance of Advanced Copper Alloys Greenbauer-Seng, L., Technical Monitor; Thomas-Ogbuji, L.; Humphrey, D. L.; Setlock, J. A.; August 2003; In English; International Symposium on Corrosion Science in the 21st Century, 6-11 Jul. 2003, Manchester, UK Contract(s)/Grant(s): NAS3-98008; WBS-22-721-26-06 Report No.(s): NASA/CR-2003-212549; E-14115; NAS 1.26:212549; No Copyright; Avail: CASI; A03, Hardcopy Resistance to oxidation and blanching is a key issue for advanced copper alloys under development for NASA’s next generation of reusable launch vehicles. Candidate alloys, including dispersion-strengthened Cu-Cr-Nb, solution-strengthened Cu-Ag-Zr, and ODS Cu-Al2O3, are being evaluated for oxidation resistance by static TGA exposures in low-p(O2) and cyclic oxidation in air, and by cyclic oxidation-reduction exposures (using air for oxidation and CO/CO2 or H2/Ar for reduction) to simulate expected service environments. The test protocol and results are presented. Author Copper Alloys; Oxidation-Reduction Reactions; Oxidation Resistance; Powder Metallurgy; Degradation

20030068243 Ohio Univ., Athens, OH, USA Integrated G and C Implementation within IDOS: A Simulink Based Reusable Launch Vehicle Simulation Fisher, Joseph E.; Bevacqua, Tim; Lawrence, Douglas A.; Zhu, J. Jim; Mahoney, Michael; January 15, 2003; In English Contract(s)/Grant(s): NAS8-01105; No Copyright; Avail: CASI; A03, Hardcopy The implementation of multiple Integrated Guidance and Control (IG&C) algorithms per ﬂight phase within a vehicle simulation poses a daunting task to coordinate algorithm interactions with the other G&C components and with vehicle subsystems. Currently being developed by Universal Space Lines LLC (USL) under contract from NASA, the Integrated

11 Development and Operations System (IDOS) contains a high fidelity Simulink vehicle simulation, which provides a means to test cutting edge G&C technologies. Combining the modularity of this vehicle simulation and Simulink s built-in primitive blocks provide a quick way to implement algorithms. To add discrete-event functionality to the unfinished IDOS simulation, Vehicle Event Manager (VEM) and Integrated Vehicle Health Monitoring (IVHM) subsystems were created to provide discrete-event and pseudo-health monitoring processing capabilities. Matlab’s Stateflow is used to create the IVHM and Event Manager subsystems and to implement a supervisory logic controller referred to as the Auto-commander as part of the IG&C to coordinate the control system adaptation and reconfiguration and to select the control and guidance algorithms for a given flight phase. Manual creation of the Stateflow charts for all of these subsystems is a tedious and time-consuming process. The Stateflow Auto-builder was developed as a Matlab based software tool for the automatic generation of a Stateflow chart from information contained in a database. This paper describes the IG&C, VEM and IVHM implementations in IDOS. In addition, this paper describes the Stateflow Auto-builder. Author Algorithms; Controllers; Simulation

20030068239 Iowa State Univ. of Science and Technology, Ames, IA, USA On-Board Entry Trajectory Planning Expanded to Sub-orbital Flight Lu, Ping; Shen, Zuojun; January 15, 2003; In English; AIAA Guidance, Navigation, and Control Conference, 11-14 Aug. 2003, Austin, TX Contract(s)/Grant(s): NAS8-01105; NAG8-1637; Copyright; Avail: CASI; A03, Hardcopy A methodology for on-board planning of sub-orbital entry trajectories is developed. The algorithm is able to generate in a time frame consistent with on-board environment a three-degree-of-freedom (3DOF) feasible entry trajectory, given the boundary conditions and vehicle modeling. This trajectory is then tracked by feedback guidance laws which issue guidance commands. The current trajectory planning algorithm complements the recently developed method for on-board 3DOF entry trajectory generation for orbital missions, and provides full-envelope autonomous adaptive entry guidance capability. The algorithm is validated and veriﬁed by extensive high ﬁdelity simulations using a sub-orbital reusable launch vehicle model and difficult mission scenarios including failures and aborts. Author Reusable Launch Vehicles; Trajectory Planning; Boundary Conditions; Algorithms

20030067947 Lockheed Martin Astronautics, Denver, CO, USA Selection of Lockheed Martin’s Preferred TSTO Configurations for the Space Launch Initiative Hopkins, Joshua B.; [2002]; In English; 2002 World Space Congress (IAF/COSPAR/AIAA), 10-19 Oct. 2002, Houston, TX, USA Contract(s)/Grant(s): NAS8-01098 Report No.(s): IAC-02-V.4.03; Copyright; Avail: CASI; A02, Hardcopy Lockheed Martin is developing concepts for safe, affordable Two Stage to Orbit (TSTO) reusable launch vehicles as part of NASA s Space Launch Initiaiive. This paper discusses the options considered for the design of the TSTO, the impact of each of these options on the vehicle configuration, the criteria used for selection of preferred configurations, and the results of the selection process. More than twenty configurations were developed in detail in order to compare optioiis such as propellant choice, serial vs. parallel burn sequence, use of propellant crossfeed between stages, bimese or optimized stage designs, and high or low staging velocities. Each configuration was analyzed not only for performance and sizing, but also for cost and reliability. The study concluded that kerosene was the superior fuel for first stages, and that bimese vehicles were not attractive. Author Spacecraft Launching; Reusable Launch Vehicles; Spacecraft Configurations; Astronautics; Spacecraft Design; NASA Programs

20030067930 Rose-Hulman Inst. of Tech., Terre Haute, IN, USA Genetic Algorithm Tuned Fuzzy Logic for Gliding Return Trajectories Burchett, Bradley T.; [2003]; In English Contract(s)/Grant(s): NAG8-1915; No Copyright; Avail: CASI; A03, Hardcopy The problem of designing and ﬂying a trajectory for successful recovery of a reusable launch vehicle is tackled using fuzzy logic control with genetic algorithm optimization. The plant is approximated by a simpliﬁed three degree of freedom

12 non-linear model. A baseline trajectory design and guidance algorithm consisting of several Mamdani type fuzzy controllers is tuned using a simple genetic algorithm. Preliminary results show that the performance of the overall system is shown to improve with genetic algorithm tuning. Author Trajectory Control; Reusable Launch Vehicles; Fuzzy Systems; Genetic Algorithms

20030067885 NASA Marshall Space Flight Center, Huntsville, AL, USA NASA’s Orbital Space Plane Risk Reduction Strategy Dumbacher, Dan; [2003]; In English; AIAA/ICAS International Air and Space Symposium and Exposition: The Next 100 Years, 14-17 Jul. 2003, Dayton, OH, USA Report No.(s): AIAA Paper 2003-3332; Copyright; Avail: CASI; A02, Hardcopy This paper documents the transformation of NASA s Space Launch Initiative (SLI) Second Generation Reusable Launch Vehicle Program under the revised Integrated Space Transportation Plan, announced November 2002. Outlining the technology development approach followed by the original SLI, this paper gives insight into the current risk-reduction strategy that will enable confident development of the Nation s first orbital space plane (OSP). The OSP will perform an astronaut and contingency cargo transportation function, with an early crew rescue capability, thus enabling increased crew size and enhanced science operations aboard the International Space Station. The OSP design chosen for full-scale development will take advantage of the latest innovations American industry has to offer. The OSP Program identifies critical technologies that must be advanced to field a safe, reliable, affordable space transportation system for U.S. access to the Station and low-Earth orbit. OSP flight demonstrators will test crew safety features, validate autonomous operations, and mature thermal protection systems. Additional enabling technologies may be identified during the OSP design process as part of an overall risk-management strategy. The OSP Program uses a comprehensive and evolutionary systems acquisition approach, while applying appropriate lessons learned. Author Risk; NASA Space Programs; Aerospace Planes; Reusable Launch Vehicles; Autonomy; Orbital Maneuvering Vehicles

20030067819 NASA Marshall Space Flight Center, Huntsville, AL, USA A Novel Approach for Modeling Chemical Reaction in Generalized Fluid System Simulation Program Sozen, Mehmet; Majumdar, Alok; [2002]; In English; 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 20-23 Jul. 2003, Huntsville, AL, USA; Copyright; Avail: CASI; A01, Hardcopy The Generalized Fluid System Simulation Program (GFSSP) is a computer code developed at NASA Marshall Space Flight Center for analyzing steady state and transient flow rates, pressures, temperatures, and concentrations in a complex flow network. The code, which performs system level simulation, can handle compressible and incompressible flows as well as phase change and mixture thermodynamics. Thermodynamic and thermophysical property programs, GASP, WASP and GASPAK provide the necessary data for fluids such as helium, methane, neon, nitrogen, carbon monoxide, oxygen, argon, carbon dioxide, fluorine, hydrogen, water, a hydrogen, isobutane, butane, deuterium, ethane, ethylene, hydrogen sulfide, krypton, propane, xenon, several refrigerants, nitrogen trifluoride and ammonia. The program which was developed out of need for an easy to use system level simulation tool for complex flow networks, has been used for the following purposes to name a few: Space Shuttle Main Engine (SSME) High Pressure Oxidizer Turbopump Secondary Flow Circuits, Axial Thrust Balance of the Fastrac Engine Turbopump, Pressurized Propellant Feed System for the Propulsion Test Article at Stennis Space Center, X-34 Main Propulsion System, X-33 Reaction Control System and Thermal Protection System, and International Space Station Environmental Control and Life Support System design. There has been an increasing demand for implementing a combustion simulation capability into GFSSP in order to increase its system level simulation capability of a liquid rocket propulsion system starting from the propellant tanks up to the thruster nozzle for spacecraft as well as launch vehicles. The present work was undertaken for addressing this need. The chemical equilibrium equations derived from the second law of thermodynamics and the energy conservation equation derived from the first law of thermodynamics are solved simultaneously by a Newton-Raphson method. The numerical scheme was implemented as a User Subroutine in GFSSP. Author Computer Programs; Computerized Simulation; Chemical Reactions; Liquid Propellant Rocket Engines; Flow Visualization

13 20030067733 NASA Marshall Space Flight Center, Huntsville, AL, USA Evaluation and Characterization Study of Dual Pulse Laser-Induced Spark (DPLIS) for Rocket Engine Ignition System Application Osborne, Robin; Wehrmeyer, Joseph; Trinh, Huu; Early, James; July 07, 2003; In English; 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 20-23 Jul. 2003, Huntsville, AL, USA; Original contains black and white illustrations; Copyright; Avail: CASI; A02, Hardcopy This paper addresses the progress of technology development of a laser ignition system at NASA Marshall Space Flight Center (MSFC). Laser ignition has been used at MSFC in recent test series to successfully ignite RP1/GOX propellants in a subscale rocket chamber, and other past studies by NASA GRC have demonstrated the use of laser ignition for rocket engines. Despite the progress made in the study of this ignition method, the logistics of depositing laser sparks inside a rocket chamber have prohibited its use. However, recent advances in laser designs, the use of fiber optics, and studies of multi-pulse laser formats3 have renewed the interest of rocket designers in this state-of the-art technology which offers the potential elimination of torch igniter systems and their associated mechanical parts, as well as toxic hypergolic ignition systems. In support of this interest to develop an alternative ignition system that meets the risk-reduction demands of Next Generation Launch Technology (NGLT), characterization studies of a dual pulse laser format for laser-induced spark ignition are underway at MSFC. Results obtained at MSFC indicate that a dual pulse format can produce plasmas that absorb the laser energy as efficiently as a single pulse format, yet provide a longer plasma lifetime. In an experiments with lean H2/air propellants, the dual pulse laser format, containing the same total energy of a single laser pulse, produced a spark that was superior in its ability to provide sustained ignition of fuel-lean H2/air propellants. The results from these experiments are being used to optimize a dual pulse laser format for future subscale rocket chamber tests. Besides the ignition enhancement, the dual pulse technique provides a practical way to distribute and deliver laser light to the combustion chamber, an important consideration given the limitation of peak power that can be delivered through optical fibers. With this knowledge, scientists and engineers at Los Alamos National Laboratory and CFD Research Corporation have designed and fabricated a miniaturized, first-generation optical prototype of a laser ignition system that could be the basis for a laser ignition system for rocket applications. This prototype will be tested at MSFC in future subscale rocket ignition tests. Author Computational Fluid Dynamics; Ignition Systems; Pulsed Lasers; Rocket Engines; Laser Outputs; Spark Ignition; Fabrication

20030067586 NASA Marshall Space Flight Center, Huntsville, AL, USA X-33 LH2 Tank Failure Investigation Findings Niedermeyer, Melinda; [2003]; In English; International Conference on Composites Engineering, 20-25 Jul. 2003, New Orleans, LA, USA; No Copyright; Avail: CASI; A03, Hardcopy This viewgraph presentation provides information on the composite sandwich-honeycomb structure of the liquid hydrogen tank of the X-33 reusable launch vehicle, and describes why the the ﬁrst pressure test to determine the tank’s structural integrity failed. The presentation includes images of the tank before and after the failed test, including photomicrographs. It then reaches conclusions on the nature of the microcracks which caused the liquid hydrogen leakage. CASI X-33 Reusable Launch Vehicle; Structural Failure; Liquid Hydrogen; Microcracks; Leakage

20030067582 Northrop Grumman Corp., Redondo, CA, USA, NASA Marshall Space Flight Center, Huntsville, AL, USA Design and Testing of Non-Toxic RCS Thrusters for Second Generation Reusable Launch Vehicle Calvignac, Jacky; Dang, Lisa; Tramel, Terri; Paseur, Lila; Champion, Robert, Technical Monitor; July 22, 2003; In English; 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 20-23 Jul. 2003, Huntsville, AL, USA Contract(s)/Grant(s): NAS8-01110; NAG8-0110 Report No.(s): AIAA Paper 2003-4922; Copyright; Avail: CASI; A03, Hardcopy Under NASA sponsorship, Northrop Grumman Space Technology (NGST) designed, built and tested two non-toxic, reaction control engines, one using liquid oxygen (LOX) and liquid hydrogen (LH2) and the other using liquid oxygen and ethanol. This paper presents the design and testing of the LOX/LH2 thruster. The two key enabling technologies are the coaxial liquid-on-liquid pintle injector and the fuelcooling duct. The workhorse thruster was hotfire tested at the NASA Marshall Space Flight Center Test Stand 500 in March and April of 2002. All tests were performed at sea-level conditions. During the test program, 7 configurations were tested, including 2 combustion chambers, 3 LOX injector pintle tips, and 4 LHp injector settings. The operating conditions surveyed were 70 to 100% thrust levels, mixture ratios from 3.27 to 4.29, and LH2 duct cooling from 18.0 to 25.5% fuel flow. The copper heat sink chamber was used for 16 burns, each burn lasting from 0.4 to 10 seconds, totaling 51.4 seconds, followed by Haynes chamber testing ranging from 0.9 to 120 seconds, totaling 300.9 seconds.

14 The performance of the engine reached 95% C* efficiency. The temperature on the Haynes chamber remained well below established material limits, with the exception of one localized hot spot. These results demonstrate that both the coaxial liquid-on-liquid pintle injector design and fuel duct concepts are viable for the intended application. The thruster headend design maintained cryogenic injection temperatures while ﬁring, which validates the selected injector design approach for minimal heat soak-back. Also, off -nominal operation without adversely impacting the thermal response of the engine showed the robustness of the duct design, a key design feature for this application. By injecting fuel into the duct, the throat temperatures are manageable, yet the split of fuel through the cooling duct does not compromise the overall combstion efficiency, which indicates that, provided proper design reﬁnement, such a concept could be applied to a high-performance version of the thruster. Author Aerospace Engineering; Reusable Launch Vehicles; Engine Tests; Combustion Chambers; Combustion Effıciency; Spacecraft Performance; Thrust

20030067336 NASA Dryden Flight Research Center, Edwards, CA, USA Advanced Range Safety System for High Energy Vehicles Claxton, Jeffrey S.; Linton, Donald F.; September 2002; In English; Eleventh International Conference on Space Planes and Hypersonic Systems and Technologies, 29 Sep. - 4 Oct. 2002, Orleans, France; Original contains black and white illustrations Contract(s)/Grant(s): WU 710-35-74-00-FF-00-RSL Report No.(s): H-2508; No Copyright; Avail: CASI; A02, Hardcopy The advanced range safety system project is a collaboration between the National Aeronautics and Space Administration and the USA Air Force to develop systems that would reduce costs and schedule for safety approval for new classes of unmanned high-energy vehicles. The mission-planning feature for this system would yield flight profiles that satisfy the mission requirements for the user while providing an increased quality of risk assessment, enhancing public safety. By improving the speed and accuracy of predicting risks to the public, mission planners would be able to expand flight envelopes significantly. Once in place, this system is expected to offer the flexibility of handling real-time risk management for the high-energy capabilities of hypersonic vehicles including autonomous return-from-orbit vehicles and extended flight profiles over land. Users of this system would include mission planners of Space Launch Initiative vehicles, space planes, and other high-energy vehicles. The real-time features of the system could make extended flight of a malfunctioning vehicle possible, in lieu of an immediate terminate decision. With this improved capability, the user would have more time for anomaly resolution and potential recovery of a malfunctioning vehicle. Author Hypersonic Vehicles; Launch Vehicles; Mission Planning; Range Safety; Aerospace Planes

20030066934 NASA Dryden Flight Research Center, Edwards, CA, USA Flight Demonstration of X-33 Vehicle Health Management System Components on the F/A-18 Systems Research Aircraft Schweikhard, Keith A.; Richards, W. Lance; Theisen, John; Mouyos, William; Garbos, Raymond; December 2001; In English; 19th Digital Avionics Systems Conference, 7-11 Oct. 2000, Philadelphia, PA, USA; Original contains black and white illustrations Contract(s)/Grant(s): WU 529-35-34-E8-RR Report No.(s): NASA/TM-2001-209037; H-2435; NAS 1.15:209037; No Copyright; Avail: CASI; A03, Hardcopy The X-33 reusable launch vehicle demonstrator has identified the need to implement a vehicle health monitoring system that can acquire data that monitors system health and performance. Sanders, a Lockheed Martin Company, has designed and developed a COTS-based open architecture system that implements a number of technologies that have not been previously used in a flight environment. NASA Dryden Flight Research Center and Sanders teamed to demonstrate that the distributed remote health nodes, fiber optic distributed strain sensor, and fiber distributed data interface communications components of the X-33 vehicle health management (VHM) system could be successfully integrated and flown on a NASA F-18 aircraft. This paper briefly describes components of X-33 VHM architecture flown at Dryden and summarizes the integration and flight demonstration of these X-33 VHM components. Finally, it presents early results from the integration and flight efforts. Author X-33 Reusable Launch Vehicle; Systems Health Monitoring; Management Systems; Systems Integration; Flight Tests

15 20030066328 Lockheed Martin Michoud Space Systems, New Orleans, LA, USA Composite Development and Applications for RLV Tankage Wright, Richard J.; Achary, David C.; McBain, Michael C.; [2003]; In English; AIAA/ICAS International Air and Space Symposium and Exposition, 14-17 Jul. 2003, Dayton, OH, USA Contract(s)/Grant(s): NCC8-191; No Copyright; Avail: CASI; A03, Hardcopy The development of polymer composite cryogenic tanks is a critical step in creating the next generation of launch vehicles. Future launch vehicles need to minimize the gross liftoff weight (GLOW), which is possible due to the 28%-41% reduction in weight that composite materials can provide over current aluminum technology. The development of composite cryogenic tanks, feedlines, and unpressurized structures are key enabling technologies for performance and cost enhancements for Reusable Launch Vehicles (RLVs). The technology development of composite tanks has provided direct and applicable data for feedlines, unpressurized structures, material compatibility, and cryogenic ﬂuid containment for highly loaded complex structures and interfaces. All three types of structure have similar material systems, processing parameters, scaling issues, analysis methodologies, NDE development, damage tolerance, and repair scenarios. Composite cryogenic tankage is the most complex of the 3 areas and provides the largest breakthrough in technology. A building block approach has been employed to bring this family of difficult technologies to maturity. This approach has built up composite materials, processes, design, analysis and test methods technology through a series of composite test programs beginning with the NASP program to meet aggressive performance goals for reusable launch vehicles. In this paper, the development and application of advanced composites for RLV use is described. Author Reusable Launch Vehicles; Composite Materials; Tanks (Containers); Cryogenics

20030066309 NASA Marshall Space Flight Center, Huntsville, AL, USA Flight Demonstrations of Orbital Space Plane (OSP) Technologies Turner, Susan; [2003]; In English; AIAA/ICAS International Air and Space Symposium, 14-18 Jul. 2003, Dayton, OH, USA Report No.(s): AIAA Paper 2003-2710; No Copyright; Avail: CASI; A02, Hardcopy The Orbital Space Plane (OSP) Program embodies NASA s priority to transport Space Station crews safely, reliably, and affordably, while it empowers the Nation s greater strategies for scientiﬁc exploration and space leadership. As early in the development cycle as possible, the OSP will provide crew rescue capability, offering an emergency ride home from the Space Station, while accommodating astronauts who are deconditioned due to long- duration missions, or those that may be ill or injured. As the OSP Program develops a fully integrated system, it will use existing technologies and employ computer modeling and simulation. Select ﬂight demonstrator projects will provide valuable data on launch, orbital, reentry, and landing conditions to validate thermal protection systems, autonomous operations, and other advancements, especially those related to crew safety and survival. Author Aerospace Vehicles; X-37 Vehicle; Space Transportation System

20030066222 NASA Marshall Space Flight Center, Huntsville, AL, USA Design and Analysis of Turbines for Space Applications Griffin, Lisa W.; Dorney, Daniel J.; Huber, Frank W.; June 13, 2003; In English; 33rd AIAA Fluid Dynamics Conference, 23-26 Jun. 2003, Orlando, FL, USA; Copyright; Avail: CASI; A03, Hardcopy In order to mitigate the risk of rocket propulsion development, efficient, accurate, detailed fluid dynamics analysis of the turbomachinery is necessary. This analysis is used for component development, design parametrics, performance prediction, and environment definition. To support this requirement, a task was developed at NASAh4arshall Space Flight Center (MSFC) to improve turbine aerodynamic performance through the application of advanced design and analysis tools. There are four major objectives of this task: 1) to develop, enhance, and integrate advanced turbine aerodynamic design and analysis tools; 2) to develop the methodology for application of the analytical techniques; 3) to demonstrate the benefits of the advanced turbine design procedure through its application to a relevant turbine design point; and 4) to verify the optimized design and analysis with testing. The turbine chosen on which to demonstrate the procedure was a supersonic design suitable for a reusable launch vehicle (RLV). The hot gas path and blading were redesigned to obtain an increased efficiency. The redesign of the turbine was conducted with a consideration of system requirements, realizing that a highly efficient turbine that, for example, significantly increases engine weight, is of limited benefit. Both preliminary and detailed designs were considered.

16 To generate an improved design, one-dimensional (1D) design and analysis tools, computational ﬂuid dynamics (CFD), response surface methodology (RSM), and neural nets (NN) were used. Author Turbomachinery; Design Analysis; Computational Fluid Dynamics; Performance Prediction; Dynamic Response; Aerodynamic Characteristics

20030066115 NASA Marshall Space Flight Center, Huntsville, AL, USA Earth-to-Orbit Rocket Propulsion Beaurain, Andre; Souchier, Alain; Moravie, Michel; Sackheim, Robert L.; Cikanek, Harry A., III; March 31, 2003; In English; International Air and Space Symposium and Exposition: The Next 100 Years, 14-18 Jul. 2003, Dayton, OH, USA; Copyright; Avail: CASI; A03, Hardcopy The Earth-to-orbit (ETO) phase of access to space is and always will be the first and most critical phase of all space missions. This first phase of all space missions has unique characteristics that have driven space launcher propulsion requirements for more than half a century. For example, the need to overcome the force of the Earth s gravity in combination with high levels of atmospheric drag to achieve the initial orbital velocity; i.e., Earth parking orbit or =9 km/s, will always require high thrust- to-weight (TN) propulsion systems. These are necessary with a T/W ratio greater than one during the ascent phase. The only type of propulsion system that can achieve these high T/W ratios are those that convert thermal energy to kinetic energy. There are only two basic sources of onboard thermal energy: chemical combustion-based systems or nuclear thermal-based systems (fission, fusion, or antimatter). The likelihood of advanced open-cycle, nuclear thermal propulsion being developed for flight readiness or becoming environmentally acceptable during the next century is extremely low. This realization establishes that chemical propulsion for ET0 launchers will be the technology of choice for at least the next century, just as it has been for the last half century of rocket flight into space. The world s space transportation propulsion requirements have evolved through several phases over the history of the space program, as has been necessitated by missions and systems development, technological capabilities available, and the growth and evolution of the utilization of space for economic, security, and science benefit. Current projections for the continuing evolution of requirements and concepts may show how future space transportation system needs could be addressed. The evolution and projections will be described in detail in this manuscript. Author Earth Orbits; Reusable Rocket Engines; Reusable Spacecraft; Reusable Launch Vehicles

20030066112 NASA Marshall Space Flight Center, Huntsville, AL, USA Multi-Disciplinary Analysis for Future Launch Systems Using NASA’s Advanced Engineering Environment (AEE) Monell, Donald; Mathias, Donovan; Reuther, James; Garn, Michelle; [2003]; In English; 16th AIAA Computational Fluid Dynamics Conference, 23-26 Jun. 2003, Orlando, FL, USA Report No.(s): AIAA Paper 2003-3428; No Copyright; Avail: CASI; A03, Hardcopy A new engineering environment constructed for the purposes of analyzing and designing Reusable Launch Vehicles (RLVs) is presented. The new environment has been developed to allow NASA to perform independent analysis and design of emerging RLV architectures and technologies. The new Advanced Engineering Environment (AEE) is both collaborative and distributed. It facilitates integration of the analyses by both vehicle performance disciplines and life-cycle disciplines. Current performance disciplines supported include: weights and sizing, aerodynamics, trajectories, propulsion, structural loads, and CAD-based geometries. Current life-cycle disciplines supported include: DDT&E cost, production costs, operations costs, ﬂight rates, safety and reliability, and system economics. Involving six NASA centers (ARC, LaRC, MSFC, KSC, GRC and JSC), AEE has been tailored to serve as a web-accessed agency-wide source for all of NASA’s future launch vehicle systems engineering functions. Thus, it is conﬁgured to facilitate (a) data management, (b) automated tool/process integration and execution, and (c) data visualization and presentation. The core components of the integrated framework are a customized PTC Windchill product data management server, a set of RLV analysis and design tools integrated using Phoenix Integration’s Model Center, and an XML-based data capture and transfer protocol. The AEE system has seen production use during the Initial Architecture and Technology Review for the NASA 2nd Generation RLV program, and it continues to undergo development and enhancements in support of its current main customer, the NASA Next Generation Launch Technology (NGLT) program. Author Reusable Launch Vehicles; Design Analysis; Systems Engineering; Systems Analysis

17 20030066011 NASA Glenn Research Center, Cleveland, OH, USA Tantalum Addition to Zirconium Diboride for Improved Oxidation Resistance Levine, Stanley R.; Opila, Eliizabeth J.; July 2003; In English; 27th Annual Conference on Composites, Materials and Structures, 27-31 Jan. 2003, Cape Canaveral, FL, USA; Original contains color and black and white illustrations Contract(s)/Grant(s): WBS 22-706-85-08 Report No.(s): NASA/TM-2003-212483; E-14025; NAS 1.15:212483; Copyright; Avail: CASI; A03, Hardcopy Ultrahigh temperature ceramics have performed unreliably due to material flaws and attachment design. These deficiencies are brought to the fore by the low fracture toughness and thermal shock resistance of UHTCs. If these deficiencies are overcome, we are still faced with poor oxidation resistance as a limitation on UHTC applicability to reusable launch vehicles. We have been addressing the deficiencies of UHTCs with our focus on composite constructions and functional grading to address the mechanical issues, and on composition modification to address the oxidation issue. The approaches and progress toward the latter are reported. Author Ceramics; Refractory Materials; Tantalum; Zirconium; Borides; Oxidation Resistance; Mechanical Properties

20030065957 NASA Marshall Space Flight Center, Huntsville, AL, USA Design and Testing of Non-Toxic RCS Thrusters for Second Generation Reusable Launch Vehicle Calvignac, Jacky; Tramel, Terri; [2003]; In English; Joint Propulsion Conference, Jul. 2003, Huntsville, AL, USA Contract(s)/Grant(s): NAS8-01110; No Copyright; Avail: Other Sources; Abstract Only The current NASA Space Shuttle auxiliary propulsion system utilizes nitrogen tetroxide (NTO) and monomethylhydrazine (MMH), hypergolic propellants. This use of these propellants has resulted in high levels of maintenance and precautions that contribute to costly launch operations. By employing alternate propellant combinations, those less toxic to humans, the hazards and time required between missions can be significantly reduced. Use of alternate propellants can thereby increase the efficiency and lower the cost in launch operations. In support of NASA’s Space Launch Initiative (SLI), TRW proposed a three-phase project structured to significantly increase the technology readiness of a high-performance reaction control subsystem (RCS) thruster using non-toxic propellant for an operationally efficient and reusable auxiliary propulsion system (APS). The project enables the development of an integrated primary/vernier thruster capable of providing dual-thrust levels of both 1000-lbf class thrust and 25-lbf thrust. The intent of the project is to reduce the risk associated with the development of an improved RCS flight design that meets the primary NASA objectives of improved safety and reliability while reducing systems operations and maintenance costs. TRW proposed two non-toxic auxiliary propulsion engine designs, one using liquid oxygen and liquid hydrogen and the other using liquid oxygen and liquid ethanol, as candidates to meet the goals of reliability and affordability at the RCS level. Both of these propellant combinations offer the advantage of a safe environment for maintenance, while at the same time providing adequate to excellent performance for a conventional liquid propulsion systems. The key enabling technology incorporated in both TRW thrusters is the coaxial liquid on liquid pintle injector. This paper will concentrate on only the design and testing of one of the thrusters, the liquid oxygen (LOX) and liquid hydrogen (LH2) thruster. The LOX/LH2 thruster design includes a LOX-centered pintle injector, consisting of two rows of slots that create a radial spoke spray pattern in the combustion chamber. The main fuel injector creates a continuous sheet of LH2 originating upstream of the LOX pintle injector. The two propellants impinge at the pintle slots, where the resulting momentum ratio and spray pattern determines the combustion efficiency and thermal effects on the hardware. Another enabling technology used in the design of this thruster is fuel film cooling through a duct, lining the inner wall of the combustion chamber barrel section. The duct is also acts as a secondary fuel injection point. The variation in the amount of LH2 used for the duct allows for adjustments in the cooling capacity for the thruster. The Non-Toxic LOX-LH2 RCS Workhorse Thruster was tested at the NASA Marshall Space Flight Center’s Test Stand 500. Hot-fire tests were conducted between March 08, 2002 and April 05, 2002. All testing during the program base period were performed at sea-level conditions. During the test program, 7 configurations were tested, including 2 combustion chambers, 3 LOX injector pintle tips, and 4 LH2 injector stroke settings. The operating conditions that were surveyed varied thrust levels, mixture ratio and LH2 duct cooling flow. The copper heat sink chamber was used for 16 burns, each burn lasting from 0.4 to 10 seconds, totaling 51.4 seconds, followed by Haynes chamber testing ranging from 0.9 to 120 seconds, totaling 300.9 seconds. The total accumulated burn time for the test program is 352.3 seconds. C* efficiency was calculated and found to be within expectable limits for most operating conditions. The temperature on the Haynes combustion chamber remained below established material limits, with the exception of one localized hot spot. The test results demonstrate that both the coaxial liquid-on-liquid pintle injector design and fuel duct concepts are viable for the intended application. The thruster head-e design maintained cryogenic injection temperatures while firing, which validates the concept for minimal heat soak back. By injecting fuel into the duct, the throat temperatures were manageable, yet the split of fuel through the cooling duct does not compromise the overall combustion efficiency, which

18 indicates that, provided proper design reﬁnement, such a concept can be applied to a high-performance version of the thruster. These hot ﬁre tests demonstrate the robustness of the duct design concept and good capability to withstand off-nominal operating conditions without adversely impacting the thermal response of the engine, a key design feature for a cryogenic thruster. Author Liquid Rocket Propellants; Thrustors; Ethyl Alcohol; Engine Tests; Aerospace Safety; Hydrogen Fuels; Liquid Hydrogen; Liquid Oxygen

20030065923 NASA Marshall Space Flight Center, Huntsville, AL, USA, NASA Johnson Space Center, Houston, TX, USA NASA’s New Orbital Space Plane: A Bridge to the Future Davis, Stephan R.; Engler, Leah M.; Fisher, Mark F.; Dumbacher, Dan L.; Boswell, Barry E.; [2003]; In English; AIAA 10th Anniverary of Flight, 14-17 Jul. 2003, Dayton, OH, USA; No Copyright; Avail: CASI; A03, Hardcopy NASA is developing a new spacecraft system called the Orbital Space Plane (OSP). The OSP will be launched on an expendable launch vehicle and serve to augment the shuttle in support of the International Space Station by transporting astronauts to and from the International Space Station and by providing a crew rescue system. Author Aerospace Planes; Development; Project Management

20030065841 NASA Marshall Space Flight Center, Huntsville, AL, USA Rocket Engine Health Management: Early Definition of Critical Flight Measurements Christenson, Rick L.; Nelson, Michael A.; Butas, John P.; July 2003; In English; 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 20-23 Jul. 2003, Huntsville, AL, USA; Original contains black and white illustrations Report No.(s): AIAA Paper 2003-5251; No Copyright; Avail: CASI; A02, Hardcopy The NASA led Space Launch Initiative (SLI) program has established key requirements related to safety, reliability, launch availability and operations cost to be met by the next generation of reusable launch vehicles. Key to meeting these requirements will be an integrated vehicle health management ( M) system that includes sensors, harnesses, software, memory, and processors. Such a system must be integrated across all the vehicle subsystems and meet component, subsystem, and system requirements relative to fault detection, fault isolation, and false alarm rate. The purpose of this activity is to evolve techniques for defining critical flight engine system measurements-early within the definition of an engine health management system (EHMS). Two approaches, performance-based and failure mode-based, are integrated to provide a proposed set of measurements to be collected. This integrated approach is applied to MSFC s MC-1 engine. Early identification of measurements supports early identification of candidate sensor systems whose design and impacts to the engine components must be considered in engine design. Author Reusable Launch Vehicles; Rocket Engines; Failure Modes; Fault Detection; False Alarms; Reliability

20030065656 NASA Marshall Space Flight Center, Huntsville, AL, USA Analysis of Parallel Burn Without Crossfeed TSTO RLV Architectures and Comparison to Parallel Burn With Crossfeed and Series Burn Architectures Smith, Garrett; Phillips, Alan; November 4, 2002; In English; 39th Joint Propulsion Conference, 20-23 Jul. 2003, Huntsville, AL, USA Contract(s)/Grant(s): NAS8-02057; Copyright; Avail: Other Sources; Abstract Only There are currently three dominant TSTO class architectures. These are Series Burn (SB), Parallel Burn with crossfeed (PBw/cf), and Parallel Burn without crossfeed (PBncf). The goal of this study was to determine what factors uniquely affect PBncf architectures, how each of these factors interact, and to determine from a performance perspective whether a PBncf vehicle could be competitive with a PBw/cf or SB vehicle using equivalent technology and assumptions. In all cases, performance was evaluated on a relative basis for a ﬁxed payload and mission by comparing gross and dry vehicle masses of a closed vehicle. Propellant combinations studied were LOX: LH2 propelled orbiter and booster (HH) and LOX: Kerosene booster with LOX: LH2 orbiter (KH). The study conclusions were: 1) a PBncf orbiter should be throttled as deeply as possible after launch until the staging point. 2) a detailed structural model is essential to accurate architecture analysis and evaluation. 3) a PBncf TSTO architecture is feasible for systems that stage at mach 7. 3a) HH architectures can achieve a mass growth relative to PBw/cf of < 20%. 3b) KH architectures can achieve a mass growth relative to Series Burn of < 20%. 4) center of gravity (CG) control will be a major issue for a PBncf vehicle, due to the low orbiter speciﬁc thrust to weight ratio and to the

19 position of the orbiter required to align the nozzle heights at liftoff. 5 ) thrust to weight ratios of 1.3 at liftoff and between 1.0 and 0.9 when staging at mach 7 appear to be close to ideal for PBncf vehicles. 6) performance for all vehicles studied is better when staged at mach 7 instead of mach 5. The study showed that a Series Burn architecture has the lowest gross mass for HH cases, and has the lowest dry mass for KH cases. The potential disadvantages of SB are the required use of an air-start for the orbiter engines and potential CG control issues. A Parallel Burn with crossfeed architecture solves both these problems, but the mechanics of a large bipropellant crossfeed system pose significant technical difficulties. Parallel Burn without crossfeed vehicles start both booster and orbiter engines on the ground and thus avoid both the risk of orbiter air-start and the complexity of a crossfeed system. The drawback is that the orbiter must use 20% to 35% of its propellant before reaching the staging point. This induces a weight penalty in the orbiter in order to carry additional propellant, which causes a further weight penalty in the booster to achieve the same staging point. One way to reduce the orbiter propellant consumption during the first stage is to throttle down the orbiter engines as much as possible. Another possibility is to use smaller or fewer engines. Throttling the orbiter engines soon after liftoff minimizes CG control problems due to a low orbiter liftoff thrust, but may result in an unnecessarily high orbiter thrust after staging. Reducing the number or size of engines size may cause CG control problems and drift at launch. The study suggested possible methods to maximize performance of PBncf vehicle architectures in order to meet mission design requirements. Author Combustion; Propulsion System Configurations; Propulsion System Performance; Reusable Launch Vehicles

20030065617 Lockheed Martin Corp., Denver, CO, USA Lockheed Martin Response to the OSP Challenge Sullivan, Robert T.; Munkres, Randy; Megna, Thomas D.; Beckham, Joanne; June 25, 2003; In English; International Air and Space Symposium and Exposition, 14-17 Jul. 2003, Dayton, OH, USA Contract(s)/Grant(s): NAS8-01098 Report No.(s): AIAA Paper 2003-2708; No Copyright; Avail: CASI; A02, Hardcopy The Lockheed Martin Orbital Space Plane System provides crew transfer and rescue for the International Space Station more safely and affordably than current human space transportation systems. Through planned upgrades and spiral development, it is also capable of satisfying the Nation’s evolving space transportation requirements and enabling the national vision for human space flight. The OSP System, formulated through rigorous requirements definition and decomposition, consists of spacecraft and launch vehicle flight elements, ground processing facilities and existing transportation, launch complex, range, mission control, weather, navigation, communication and tracking infrastructure. The concept of operations, including procurement, mission planning, launch preparation, launch and mission operations and vehicle maintenance, repair and turnaround, is structured to maximize flexibility and mission availability and minimize program life cycle cost. The approach to human rating and crew safety utilizes simplicity, performance margin, redundancy, abort modes and escape modes to mitigate credible hazards that cannot be designed out of the system. Author Aerospace Planes; Launch Vehicles; Ground Support Systems; Spacecraft Launching; Navigation; Spacecraft Docking

20030064145 NASA Marshall Space Flight Center, Huntsville, AL, USA X-37 C-SiC CMC Control Surface Components Development: Status of the NASA/Boeing/USAF Orbital Vehicle and Related Efforts Valentine, Peter G.; Rivers, H. Kevin; Chen, Victor L.; 2003; In English; 27th Annual Conference on Composites, Materials, and Structures, 27-30 Jan. 2003, Cape Canaveral, FL, USA; Copyright; Avail: Other Sources; Abstract Only Carbon/Silicon-Carbide (C-SiC) ceramic matrix composite (CMC) ﬂaperon and ruddervator control surface components are being developed for the X-37 Orbital Vehicle (OV). The results of the prior NASA LaRC led work, aimed at developing C-SiC ﬂaperon and ruddervator components for the X-37, will be reviewed. The status of several on-going and/or planned NASA, USAF, and Boeing programs that will support the development of control surface components for the X-37 OV will also be reviewed. The overall design and development philosophy being employed to assemble a team(s) to develop both (a) C-SiC hot structure control surface components for the X-37 OV, and (b) carbon-carbon (C-C) hot structure components (a risk-reduction backup option for the OV), will be presented. Author Ceramic Matrix Composites; X-37 Vehicle; Silicon Carbides; Control Surfaces

20 20030064094 NASA Marshall Space Flight Center, Huntsville, AL, USA Permeability After Impact Testing of Composite Laminates Nettles, Alan T.; [2003]; In English; SAMPE 2003 Symposium and Exhibition, 11-15 May 2003, Long Beach, CA, USA; No Copyright; Avail: CASI; A03, Hardcopy Since composite laminates are beginning to be identified for use in reusable launch vehicle propulsion systems, an understanding of their permeance is needed. A foreign object impact event can cause a localized area of permeability (leakage) in a polymer matrix composite and it is the aim of this study to assess a method of quantifying permeability-after-impact results. A simple test apparatus is presented and variables that could affect the measured values of permeability-after-impact were assessed. Once it was determined that valid numbers were being measured, a fiber/resin system was impacted at various impact levels and the resulting permeability measured, first with a leak check solution (qualitative) then using the new apparatus (quantitative). The results showed that as the impact level increased, so did the measured leakage. As the pressure to the specimen was increased, the leak rate was seen to increase in a non-linear fashion for almost all of the specimens tested. Author Laminates; Permeability; Impact Tests; Leakage; Polymer Matrix Composites

20030064052 NASA Marshall Space Flight Center, Huntsville, AL, USA Ceramic Matrix Composite Cooled Nozzle Material Development Program Lawrence, Tim; Eckel, Andy; Porter, John; Pichon, T.; Patterson, B.; Paquette, T.; [2002]; In English; JANNAF 13th NDES Meeting, 27-29 Aug. 2002, San Antonio, TX, USA Contract(s)/Grant(s): NAS8-00141; NAS8-99112; NAS8-00140; NAS8-00124; No Copyright; Avail: Other Sources; Abstract Only The X-33 program initiated a risk reduction technology project to develop an actively cooled cermic matrix composite (CMC) nozzle ramp for the linear aerospike engine. The objective was to reduce the weight and increase the operating temperature capabilities of the nozzle ramp. A complement to this original project was subsequently supported by NASA’s Second Generation Reusable Launch Vehicle Program to develop a high risk high payoff cooled composite nozzle ramp. This project focused on less mature technologies and concepts having the potential to achieve a significant weight reduction beyond those systems which were being considered in the original X-33 project. The aerospike engine was not selected under the initial Space Launch Initiative (SLI) Program research announcement. However, in recognition of the tremendous application opportunities of such technology to other areas and systems in the rocket industry, the effort was continued and is in the process of being transferred to the 3rd Generation Reusable Launch Vehicle (RLV) Program where it will be combined with the current cooled CMC panel project. The objective of the refocused project is to advance the material, design, and analysis work on the ramp concepts to a point that would allow the selection of the most promising candidate(s) for continued development. The concept(s) carried on in the 3rd Generation RLV Program will be modified to address the goals of this program. Originally, four contracts with different design concepts were initiated. Each contractor has performed design and analysis of their concept and submitted a subscale component for testing in the Cell 22 test rig at Glenn Research Center. This paper will discuss the results to date of each design concept and the potential applications to future rocket nozzle systems. The engineering technology challenges for each concept were determined and addressed during this phase of the effort. These challenges and the success in addressing them will also be discussed. High temperature, noneroding nozzle materials have the potential to significantly improve nozzle performance. The work performed under this effort is at the forefront of CMC material development and will give the rocket nozzle community a good view into the status of cooled CMC materials. Author Ceramic Matrix Composites; X-33 Reusable Launch Vehicle; Rocket Nozzles; Ramps (Structures)

20030064017 Boeing Phantom Works, Huntington Beach, CA, USA Boeing 2nd Generation Reusable Launch Vehicle Architecture Bienhoff, Dallas; [2003]; In English; Space Technology and Applications International Forum (STAIF 2003), 2-5 Feb. 2003, Albuquerque, NM, USA Contract(s)/Grant(s): NAS8-01099; No Copyright; Avail: CASI; A03, Hardcopy This paper presents viewgraphs on the Space Launch Initiative (SLI) space shuttle safety upgrades. CASI Reusable Launch Vehicles; Spacecraft Design; Space Shuttles; Space Transportation System

21 20030063938 NASA Marshall Space Flight Center, Huntsville, AL, USA Selection and Evaluation of An Alloy for Nozzle Application Pandey, Awadh; Shah, Sandeep; Shadoan, Mike; May 22, 2003; In English; Aeromat Conference 2003, 9-12 Jun. 2003, Dayton, OH, USA; Copyright; Avail: CASI; A03, Hardcopy The goal was to select and characterize a material with superior weldability and hydrogen resistance for nozzle application under Second Generation Reusable Launch Vehicle (2GRLV) program funded by NASA_MSFC. Currently standard A-286 is used for SSME nozzle.Howewer, material with superior weldability and hydrogen resistance will be desired for (2GRLV). Derived from text Product Development; Alloys; Hydrogen; Weldability

20030063899 Lockheed Martin Engineering and Sciences Co., Hampton, VA, USA Characterization of Thin Film Polymers Through Dynamic Mechanical Analysis and Permeation Herring, Helen; June 2003; In English Contract(s)/Grant(s): NAS1-00135; 762-30-51-03 Report No.(s): NASA/CR-2003-212422; NAS 1.26:212422; No Copyright; Avail: CASI; A03, Hardcopy Thin polymer films are being considered, as candidate materials to augment the permeation resistance of cryogenic hydrogen fuel tanks such as would be required for future reusable launch vehicles. To evaluate performance of candidate films after environmental exposure, an experimental study was performed to measure the thermal/mechanical and permeation performance of six, commercial-grade materials. Dynamic storage modulus, as measured by Dynamic Mechanical Analysis, was found over a range of temperatures. Permeability, as measured by helium gas diffusion, was found at room temperature. Test data was correlated with respect to film type and pre-test exposure to moisture, elevated temperature, and cryogenic temperature. Results indicated that the six films were comparable in performance and their resistance to environmental degradation. Author Thin Films; Permeability; Polymeric Films; Cryogenic Fluid Storage; Hydrogen Fuels

20030063104 Clark-Atlanta Univ., GA, USA Characterization of Polyimide Foams for Ultra-Lightweight Space Structures Meador, Michael, Technical Monitor; Hillman, Keithan; Veazie, David R.; HBCUs/OMUs Research Conference Agenda and Abstracts; February 2003; In English; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document Ultra-lightweight materials have played a significant role in nearly every area of human activity ranging from magnetic tapes and artificial organs to atmospheric balloons and space inflatables. The application range of ultra-lightweight materials in past decades has expanded dramatically due to their unsurpassed efficiency in terms of low weight and high compliance properties. A new generation of ultra-lightweight materials involving advanced polymeric materials, such as TEEK (TM) polyimide foams, is beginning to emerge to produce novel performance from ultra-lightweight systems for space applications. As a result, they require that special conditions be fulfilled to ensure adequate structural performance, shape retention, and thermal stability. It is therefore important and essential to develop methodologies for predicting the complex properties of ultra-lightweight foams. To support NASA programs such as the Reusable Launch Vehicle (RLV), Clark Atlanta University, along with SORDAL, Inc., has initiated projects for commercial process development of polyimide foams for the proposed cryogenic tank integrated structure (see figure 1). Fabrication and characterization of high temperature, advanced aerospace-grade polyimide foams and filled foam sandwich composites for specified lifetimes in NASA space applications, as well as quantifying the lifetime of components, are immensely attractive goals. In order to improve the development, durability, safety, and life cycle performance of ultra-lightweight polymeric foams, test methods for the properties are constant concerns in terms of timeliness, reliability, and cost. A major challenge is to identify the mechanisms of failures (i.e., core failure, interfacial debonding, and crack development) that are reflected in the measured properties. The long-term goal of the this research is to develop the tools and capabilities necessary to successfully engineer ultra-lightweight polymeric foams. The desire is to reduce density at the material and structural levels, while at the same time maintaining or increasing mechanical and other properties. Author Foams; Polyimides; Technology Utilization; Costs; Fabrication

22 20030062263 Boeing Co., Huntington Beach, CA, USA Thermal-Mechanical Cyclic Test of a Composite Cryogenic Tank for Reusable Launch Vehicles Messinger, Ross; Pulley, John; [2003]; In English; 44th AIAA/ASME/ASCE/AHS/ASC Structure, Structural Dynamics, & Material Conference, 7-10 Apr. 2003, Norfolk, VA, USA Contract(s)/Grant(s): NCC8-39; No Copyright; Avail: CASI; A03, Hardcopy This viewgraph presentation provides an overview of thermal-mechanical cyclic tests conducted on a composite cryogenic tank designed for reusable launch vehicles. Topics covered include: a structural analysis of the composite cryogenic tank, a description of Marshall Space Flight Center’s Cryogenic Structure Test Facility, cyclic test plans and accomplishments, burst test and analysis and post-testing evaluation. CASI Cryogenic Fluid Storage; Storage Tanks; Composite Materials; Destructive Tests; Cryogenic Fluids; Thermal Cycling Tests; Mechanical Properties

20030062173 NASA Marshall Space Flight Center, Huntsville, AL, USA Vision for CFD-Based Combustion Instability Predictions Rocker, Marvin; West, Jeffrey S.; April 15, 2003; In English; NASA Spring Workshop on Fluids, 22-24 Apr. 2003, Birmingham, AL, USA; No Copyright; Avail: CASI; A03, Hardcopy This viewgraph representation presents an overview of the proposed development for computational ﬂuid dynamics (CFD) based combustion instability predictions. These predictions are discussed in regard to Next Generation Launch Technologies (NGLT). Topics covered include: relevance of combustion instability to NGLT, evolution of combustion instability prediction techniques, and development of CFD-based models to support combustion instability predictions to support NGLT and beyond. CASI Combustion Stability; Computational Fluid Dynamics; Prediction Analysis Techniques; Launch Vehicles; Spacecraft Propulsion; Mathematical Models

20030062123 NASA Langley Research Center, Hampton, VA, USA Comparison of Methods for Determining Boundary Layer Edge Conditions for Transition Correlations Liechty, Derek S.; Berry, Scott A.; Hollis, Brian R.; Horvath, Thomas J.; [2003]; In English; 33rd AIAA Fluid Dynamics Conference and Exhibit, 23-26 Jun. 2003, Orlando, FL, USA; Original contains color illustrations Report No.(s): AIAA Paper 2003-3590; No Copyright; Avail: CASI; A03, Hardcopy Data previously obtained for the X-33 in the NASA Langley Research Center 20-Inch Mach 6 Air Tunnel have been reanalyzed to compare methods for determining boundary layer edge conditions for use in transition correlations. The experimental results were previously obtained utilizing the phosphor thermography technique to monitor the status of the boundary layer downstream of discrete roughness elements via global heat transfer images of the X-33 windward surface. A boundary layer transition correlation was previously developed for this data set using boundary layer edge conditions calculated using an inviscid/integral boundary layer approach. An algorithm was written in the present study to extract boundary layer edge quantities from higher fidelity viscous computational fluid dynamic solutions to develop transition correlations that account for viscous effects on vehicles of arbitrary complexity. The boundary layer transition correlation developed for the X-33 from the viscous solutions are compared to the previous boundary layer transition correlations. It is shown that the boundary layer edge conditions calculated using an inviscid/integral boundary layer approach are significantly different than those extracted from viscous computational fluid dynamic solutions. The present results demonstrate the differences obtained in correlating transition data using different computational methods. Author Boundary Layer Transition; Inviscid Flow; Computational Fluid Dynamics; Surface Roughness; Viscous Flow

20030062069 NASA Marshall Space Flight Center, Huntsville, AL, USA Systems Engineering Approach to Technology Integration for NASA’s 2nd Generation Reusable Launch Vehicle Thomas, Dale; Smith, Charles; Thomas, Leann; Kittredge, Sheryl; [2002]; In English; Space Technology Applications International Forum, 2-6 Feb. 2003, Albuquerque, NM, USA; Copyright; Avail: Other Sources; Abstract Only The overall goal of the 2nd Generation RLV Program is to substantially reduce technical and business risks associated with developing a new class of reusable launch vehicles. NASA’s speciﬁc goals are to improve the safety of a 2nd-generation system by 2 orders of magnitude - equivalent to a crew risk of 1-in-10,000 missions - and decrease the cost tenfold, to

23 approximately $1,000 per pound of payload launched. Architecture definition is being conducted in parallel with the maturating of key technologies specifically identified to improve safety and reliability, while reducing operational costs. An architecture broadly includes an Earth-to-orbit reusable launch vehicle, on-orbit transfer vehicles and upper stages, mission planning, ground and flight operations, and support infrastructure, both on the ground and in orbit. The systems engineering approach ensures that the technologies developed - such as lightweight structures, long-life rocket engines, reliable crew escape, and robust thermal protection systems - will synergistically integrate into the optimum vehicle. To best direct technology development decisions, analytical models are employed to accurately predict the benefits of each technology toward potential space transportation architectures as well as the risks associated with each technology. Rigorous systems analysis provides the foundation for assessing progress toward safety and cost goals. The systems engineering review process factors in comprehensive budget estimates, detailed project schedules, and business and performance plans, against the goals of safety, reliability, and cost, in addition to overall technical feasibility. This approach forms the basis for investment decisions in the 2nd Generation RLV Program’s risk-reduction activities. Through this process, NASA will continually refine its specialized needs and identify where Defense and commercial requirements overlap those of civil missions. Author Reusable Launch Vehicles; Risk; Safety; Reliability; Flight Operations; Ground Operational Support System; Mission Planning

20030061202 NASA Marshall Space Flight Center, Huntsville, AL, USA Overview of MSFC’s Applied Fluid Dynamics Analysis Group Activities Garcia, Roberto; Griffin, Lisa; Williams, Robert; [2002]; In English; MSFC Fall Fluids Workshop, 19-21 Nov. 2002, Huntsville, AL, USA; No Copyright; Avail: CASI; A03, Hardcopy This viewgraph report presents an overview of activities and accomplishments of NASA’s Marshall Space Flight Center’s Applied Fluid Dynamics Analysis Group. Expertise in this group focuses on high-fidelity fluids design and analysis with application to space shuttle propulsion and next generation launch technologies. Topics covered include: computational fluid dynamics research and goals, turbomachinery research and activities, nozzle research and activities, combustion devices, engine systems, MDA development and CFD process improvements. CASI Computational Fluid Dynamics; NASA Programs; Research Facilities; Turbomachinery; Systems Engineering; Design Analysis; Spacecraft Propulsion

20030061159 NASA Marshall Space Flight Center, Huntsville, AL, USA Systems Engineering Approach to Technology Integration for NASA’s 2nd Generation Reusable Launch Vehicle Thomas, Dale; Smith, Charles; Thomas, Leann; Kittredge, Sheryl; [2002]; In English; Space Technology Applications International Forum, 2-6 Feb. 2003, Albuquerque, NM, USA; No Copyright; Avail: CASI; A03, Hardcopy The overall goal of the 2nd Generation RLV Program is to substantially reduce technical and business risks associated with developing a new class of reusable launch vehicles. NASA’s specific goals are to improve the safety of a 2nd generation system by 2 orders of magnitude - equivalent to a crew risk of 1-in-10,000 missions - and decrease the cost tenfold, to approximately $1,000 per pound of payload launched. Architecture definition is being conducted in parallel with the maturating of key technologies specifically identified to improve safety and reliability, while reducing operational costs. An architecture broadly includes an Earth-to-orbit reusable launch vehicle, on-orbit transfer vehicles and upper stages, mission planning, ground and flight operations, and support infrastructure, both on the ground and in orbit. The systems engineering approach ensures that the technologies developed - such as lightweight structures, long-life rocket engines, reliable crew escape, and robust thermal protection systems - will synergistically integrate into the optimum vehicle. To best direct technology development decisions, analytical models are employed to accurately predict the benefits of each technology toward potential space transportation architectures as well as the risks associated with each technology. Rigorous systems analysis provides the foundation for assessing progress toward safety and cost goals. The systems engineering review process factors in comprehensive budget estimates, detailed project schedules, and business and performance plans, against the goals of safety, reliability, and cost, in addition to overall technical feasibility. This approach forms the basis for investment decisions in the 2nd Generation RLV Program’s risk-reduction activities. Through this process, NASA will continually refine its specialized needs and identify where Defense and commercial requirements overlap those of civil missions. Author Reusable Launch Vehicles; Systems Engineering; Technology Assessment; Space Transportation; Safety Factors; Ground Operational Support System

24 20030060643 NASA Marshall Space Flight Center, Huntsville, AL, USA Base-Bleed Effect on X-33 Aerospike Plume Induced Base-Heating Environment During Power-Pack Out Wang, Tee-See; Droege, Alan; D’Agostino, Mark; Lee, Young-Ching; Williams, Robert; [2003]; In English; 36th AIAA Thermophysics Conference, 23-26 Jun. 2003, Orlando, FL, USA; No Copyright; Avail: CASI; A03, Hardcopy A computational heat transfer methodology was developed to study the dual-engine linear aerospike plume induced base-heating environment during one power-pack out, in ascent flight. One power-pack out results in reduction of power levels for both engines. That, in turn, reduces the amount of base-bleed and changes the distribution of base-bleed on the two pillows. Hence, the concern of increased base-heating during power-pack out. The thermo-flowfield of the entire vehicle was computed. The computational methodology for the convective heating is based on a three-dimensional, finite-volume, viscous, chemically reacting, and pressure-based computational fluid dynamics formulation. The computational methodology for the radiative heating is based on a three-dimensional, finite-volume, and spectral-line-based weighted-sum-of-gray-gases absorption computational radiation heat transfer formulation. A separate radiation model was used for diagnostic purposes. The computational methodology was systematically benchmarked. In this study, near-base radiative heat fluxes were computed and they compared well with those measured from an installed linear aerospike engine tests. The base-heating environment of 18 trajectory points selected from three power-pack out ascent scenarios was computed and is presented here. The power-pack out condition has the most impact on convective base-heating when it happens early in flight. The some of its impact comes from the asymmetric and reduced base-bleed. Author Aerospike Engines; X-33 Reusable Launch Vehicle; Base Heating; Plumes

20030060421 NASA Marshall Space Flight Center, Huntsville, AL, USA CFD-Based Design Optimization for Single Element Rocket Injector Vaidyanathan, Rajkumar; Tucker, Kevin; Papila, Nilay; Shyy, Wei; [2003]; In English; 41st Aerospace Sciences Meeting and Exhibit, 6-9 Jan. 2003, Reno, NV, USA; Copyright; Avail: CASI; A03, Hardcopy To develop future Reusable Launch Vehicle concepts, we have conducted design optimization for a single element rocket injector, with overall goals of improving reliability and performance while reducing cost. Computational solutions based on the Navier-Stokes equations, finite rate chemistry, and the k-E turbulence closure are generated with design of experiment techniques, and the response surface method is employed as the optimization tool. The design considerations are guided by four design objectives motivated by the consideration in both performance and life, namely, the maximum temperature on the oxidizer post tip, the maximum temperature on the injector face, the adiabatic wall temperature, and the length of the combustion zone. Four design variables are selected, namely, H2 flow angle, H2 and O2 flow areas with fixed flow rates, and O2 post tip thickness. In addition to establishing optimum designs by varying emphasis on the individual objectives, better insight into the interplay between design variables and their impact on the design objectives is gained. The investigation indicates that improvement in performance or life comes at the cost of the other. Best compromise is obtained when improvements in both performance and life are given equal importance. Author Computational Fluid Dynamics; Injectors; Reusable Launch Vehicles; Design Optimization; Rocket Engines; Space Shuttle Main Engine

20030059030 NASA Langley Research Center, Hampton, VA, USA X-33 Computational Aeroheating/Aerodynamic Predictions and Comparisons With Experimental Data Hollis, Brian R.; Thompson, Richard A.; Berry, Scott A.; Horvath, Thomas J.; Murphy, Kelly J.; Nowak, Robert J.; Alter, Stephen J.; January 05, 2003; In English; Original contains color and black and white illustrations Contract(s)/Grant(s): WU 762-30-51-30 Report No.(s): NASA/TP-2003-212160; L-18254; NAS 1.60:212160; No Copyright; Avail: CASI; A05, Hardcopy This report details a computational fluid dynamics study conducted in support of the phase II development of the X-33 vehicle. Aerodynamic and aeroheating predictions were generated for the X-33 vehicle at both flight and wind-tunnel test conditions using two finite-volume, Navier-Stokes solvers. Aerodynamic computations were performed at Mach 6 and Mach 10 wind-tunnel conditions for angles of attack from 10 to 50 with body-flap deflections of 0 to 20. Additional aerodynamic computations were performed over a parametric range of free-stream conditions at Mach numbers of 4 to 10 and angles of attack from 10 to 50. Laminar and turbulent wind-tunnel aeroheating computations were performed at Mach 6 for angles of attack of 20 to 40 with body-flap deflections of 0 to 20. Aeroheating computations were performed at four flight conditions with Mach numbers of 6.6 to 8.9 and angles of attack of 10 to 40. Surface heating and pressure distributions, surface

25 streamlines, ﬂow ﬁeld information, and aerodynamic coefficients from these computations are presented, and comparisons are made with wind-tunnel data. Author Aerodynamic Heating; Computational Fluid Dynamics; Flow Distribution; Navier-Stokes Equation; Pressure Distribution

20030059016 NASA Langley Research Center, Hampton, VA, USA Advanced Durability and Damage Tolerance Design and Analysis Methods for Composite Structures: Lessons Learned from NASA Technology Development Programs Harris, Charles E.; Starnes, James H., Jr.; Shuart, Mark J.; June 2003; In English Contract(s)/Grant(s): 706-85-12-01 Report No.(s): NASA/TM-2003-212420; L-18297; NAS 1.15:212420; No Copyright; Avail: CASI; A03, Hardcopy Aerospace vehicles are designed to be durable and damage tolerant. Durability is largely an economic life-cycle design consideration whereas damage tolerance directly addresses the structural airworthiness (safety) of the vehicle. However, both durability and damage tolerance design methodologies must address the deleterious effects of changes in material properties and the initiation and growth of microstructural damage that may occur during the service lifetime of the vehicle. Durability and damage tolerance design and certiﬁcation requirements are addressed for commercial transport aircraft and NASA manned spacecraft systems. The state-of-the-art in advanced design and analysis methods is illustrated by discussing the results of several recently completed NASA technology development programs. These programs include the NASA Advanced Subsonic Technology Program demonstrating technologies for large transport aircraft and the X-33 hypersonic test vehicle demonstrating technologies for a single-stage-to-orbit space launch vehicle. Author Aircraft Reliability; Damage; Design Analysis; Microstructure; Life (Durability)

20030059007 NASA Langley Research Center, Hampton, VA, USA X-33 Rev-F Turbulent Aeroheating Results From Test 6817 in NASA Langley 20-Inch Mach 6 Air Tunnel and Comparisons With Computations Hollis, Brian R.; Horvath, Thomas J.; Berry, Scott A.; June 2003; In English Contract(s)/Grant(s): WU 721-28-00-05 Report No.(s): NASA/TM-2003-211962; L-18293; NAS 1.15:211962; No Copyright; Avail: CASI; A04, Hardcopy Measurements and predictions of the X-33 turbulent aeroheating environment have been performed at Mach 6, perfect-gas air conditions. The purpose of this investigation was to compare measured turbulent aeroheating levels on smooth models, models with discrete trips, and models with arrays of bowed panels (which simulate bowed thermal protections system tiles) with each other and with predictions from two Navier-Stokes codes, LAURA and GASP. The wind tunnel testing was conducted at free stream Reynolds numbers based on length of 1.8 x 10(exp 6) to 6.1 x 10(exp 6) on 0.0132 scale X-33 models at a = 40-deg. Turbulent ﬂow was produced by the discrete trips and by the bowed panels at ill but the lowest Reynolds number, but turbulent ﬂow on the smooth model was produced only at the highest Reynolds number. Turbulent aeroheating levels on each of the three model types were measured using global phosphor thermography and were found to agree to within .he estimated uncertainty (plus or minus 15%) of the experiment. Computations were performed at the wind tunnel free stream conditions using both codes. Turbulent aeroheating levels predicted using the LAURA code were generally 5%-10% lower than those from GASP, although both sets of predictions fell within the experimental accuracy of the wind tunnel data. Author X-33 Reusable Launch Vehicle; Aerodynamic Heating; Hypersonic Speed; Turbulent Flow; Scale Models; Wind Tunnel Tests; Computational Fluid Dynamics

20030055618 NASA Glenn Research Center, Cleveland, OH, USA Affordable Flight Demonstration of the GTX Air-Breathing SSTO Vehicle Concept Krivanek, Thomas M.; Roche, Joseph M.; Riehl, John P.; Kosnareo, Daniel N.; April 2003; In English; Combustion, Airbreathing Propulsion, Propulsion Systems Hazards, and Modelling and Simulation Subcommittees Joint Meeting, 8-12 Apr. 2002, Destin, FL, USA; Original contains color illustrations Contract(s)/Grant(s): WBS 22-708-90-63 Report No.(s): NASA/TM-2003-212315; E-13316; NAS 1.15:212315; APS-III-22; No Copyright; Avail: CASI; A03, Hardcopy The rocket based combined cycle (RBCC) powered single-stage-to-orbit (SSTO) reusable launch vehicle has the potential

26 to significantly reduce the total cost per pound for orbital payload missions. To validate overall system performance, a flight demonstration must be performed. This paper presents an overview of the first phase of a flight demonstration program for the GTX SSTO vehicle concept. Phase 1 will validate the propulsion performance of the vehicle configuration over the supersonic and hypersonic airbreathing portions of the trajectory. The focus and goal of Phase 1 is to demonstrate the integration and performance of the propulsion system flowpath with the vehicle aerodynamics over the air-breathing trajectory. This demonstrator vehicle will have dual mode ramjetkcramjets, which include the inlet, combustor, and nozzle with geometrically scaled aerodynamic surface outer mold lines (OML) defining the forebody, boundary layer diverter, wings, and tail. The primary objective of this study is to demonstrate propulsion system performance and operability including the ram to scram transition, as well as to validate vehicle aerodynamics and propulsion airframe integration. To minimize overall risk and develop ment cost the effort will incorporate proven materials, use existing turbomachinery in the propellant delivery systems, launch from an existing unmanned remote launch facility, and use basic vehicle recovery techniques to minimize control and landing requirements. A second phase would demonstrate propulsion performance across all critical portions of a space launch trajectory (lift off through transition to all-rocket) integrated with flight-like vehicle systems. Author Rocket-Based Combined-Cycle Engines; Single Stage to Orbit Vehicles; Air Breathing Engines; Flight Tests; Reusable Launch Vehicles; Aircraft Design; Hypersonic Aircraft

20030054429 NASA Kennedy Space Center, Cocoa Beach, FL, USA STS-102 Photo-op/Suit-up/Depart O&C/Launch Discovery On Orbit/Landing/Crew Egress March 22, 2001; In English; No Copyright; Avail: CASI; V04, Videotape-VHS; B04, Videotape-Beta The spacecrews of STS-102 and the Expedition 1 and 2 crews of the International Space Station (ISS) are seen in this video, which presents an overview of their activities. The crew consists of Commander Jim Wetherbee, Pilot James Kelly, and Mission Specialists Andrew Thomas, and Paul Richards. The sections of the video include: Photo-op, Suit-up, Depart O&C, Ingress, Launch with Playbacks, On-orbit, Landing with Playbacks, and Crew Egress & Departs. The prelaunch activities are explained by two narrators, and the crew members are assisted in the White Room just before boarding the Space Shuttle Discovery. Isolated views of the shuttle’s launch include: VAB, PAD-B, DLTR-3, UCS-23 Tracker, PATRICK IGOR, UCS-10 Tracker, Grandstand, Tower-1, OTV-160, OTV-170, OTV-171, and On-board Camera. The video shows two extravehicular activities (EVAs) to perform work on the ISS, one by astronauts Helms and Voss from Expedition 2, and another by Richards and Thomas. The attachment of the Leonardo Multipurpose Logistics Module, a temporary resupply module, is shown in a series of still images. The on-orbit footage also includes a view of the Nile River, and a crew exhange ceremony between Expedition 1 (Commander Yuri Gidzenko, Flight Engineer Sergei Krikalev) and Expedition 2 (Commander Yury Usachev, Flight Engineers James Voss, Susan Helms). Isolated views of the landing at Kennedy Space Center include: North Runway Camera, VAB, Tower-1, Mid-field, Midfield IR, Tower-2, and UCS-12 IR. The Crew Transfer Vehicle (CTV) for unloading the astronauts is shown, administrators greet the crew upon landing, and Commander Wetherbee gives a briefing. CASI Discovery (Orbiter); International Space Station; Spacecrews; Liftoff (Launching); Horizontal Spacecraft Landing; Spacecrew Transfer

20030053149 Vanderbilt Univ., Nashville, TN, USA System Risk Assessment and Allocation in Conceptual Design Mahadevan, Sankaran; Smith, Natasha L.; Zang, Thomas A., Technical Monitor; May 2003; In English Contract(s)/Grant(s): NCC1-01031; RTOP 706-31-61-01 Report No.(s): NASA/CR-2003-212162; NAS 1.26:21262; No Copyright; Avail: CASI; A06, Hardcopy As aerospace systems continue to evolve in addressing newer challenges in air and space transportation, there exists a heightened priority for significant improvement in system performance, cost effectiveness, reliability, and safety. Tools, which synthesize multidisciplinary integration, probabilistic analysis, and optimization, are needed to facilitate design decisions allowing trade-offs between cost and reliability. This study investigates tools for probabilistic analysis and probabilistic optimization in the multidisciplinary design of aerospace systems. A probabilistic optimization methodology is demonstrated for the low-fidelity design of a reusable launch vehicle at two levels, a global geometry design and a local tank design. Probabilistic analysis is performed on a high fidelity analysis of a Navy missile system. Furthermore, decoupling strategies are introduced to reduce the computational effort required for multidisciplinary systems with feedback coupling. Author Multidisciplinary Design Optimization; Aerospace Systems; Probability Theory; Reliability Analysis; Design Analysis; Risk; Systems Engineering; Models

27 20030046938 Analysis of space concepts enabled by new transportation (ASCENT) study Webber, Derek; AIP Conference Proceedings; January 14, 2002; ISSN 0094-243X; Volume 608, no. 1; In English; SPACE TECHNOLOGY and APPLICATIONS INTERNATIONAL FORUM- STAIF 2002, 3-6 Feb 2002, Albuquerque, New Mexico, USA; Copyright The ASCENT Study is evaluating all those markets (both commercial and government missions) that may be served by a future Second Generation Reusable Launch Vehicle. The markets are being analyzed, and twenty-year forecasts are being generated, in order to assist NASA in deﬁning the attributes of the Second Generation RLV that will eventually replace the Space Shuttle. The forecasts will make possible an evaluation of alternative architectures for the Second Generation RLV, and also help assess the extent to which new commercial markets, enabled by the proposed pricing and reliability beneﬁts of the Second Generation RLV, can contribute to the funding of the vehicle. This paper provides a status of the project at the time of the conference, and describes the way in which the markets are being treated and data is being obtained. The sectors range from the relatively well understood telecommunications markets to the newer evolving market opportunities like public space travel. [copyright] 2002 American Institute of Physics. Author (AIP) Aerospace Industry; Commerce; Cost Analysis; Cost Effectiveness; Cost Estimates; Market Research; Reliability; Reusable Launch Vehicles; Space Transportation; Spacecraft; Telecommunication; Transportation

20030046831 Advanced spacecraft designs in support of human missions to earth’s neighborhood Fletcher, David; AIP Conference Proceedings; January 14, 2002; ISSN 0094-243X; Volume 608, no. 1; In English; SPACE TECHNOLOGY and APPLICATIONS INTERNATIONAL FORUM- STAIF 2002, 3-6 Feb 2002, Albuquerque, New Mexico, USA; Copyright NASA’s strategic planning for technology investment draws on engineering studies of potential future missions. A number of hypothetical mission architectures have been studied. A recent study completed by the NASA/JSC Advanced Design Team addresses one such possible architecture strategy for missions to the moon. This conceptual study presents an overview of each of the spacecraft elements that would enable such missions. These elements include an orbiting lunar outpost at lunar L1 called the Gateway, a crew transfer vehicle (CTV) which ferries a crew of four from the ISS to the Gateway, a lunar lander which ferries the crew from the Gateway to the lunar surface, and a one-way lunar habitat lander capable of supporting the crew for 30 days. Other supporting elements of this architecture discussed below include the CTV kickstage, a solar-electric propulsion (SEP) stage, and a logistics lander capable of re-supplying the 30-day habitat lander and bringing other payloads totaling 10. 3 mt in support of surface mission activities. Launch vehicle infrastructure to low-earth orbit includes the Space Shuttle, which brings up the CTV and crew, and the Delta-IV Heavy expendable launch vehicle which launches the landers, kickstage, and SEP. [copyright] 2002 American Institute of Physics. Author (AIP) Earth Orbits; Electric Propulsion; Lunar Surface; Management Planning; Spacecraft

20030046819 Estimate of avoidance maneuver rate for HASTOL tether boost facility Forward, Robert L.; AIP Conference Proceedings; January 14, 2002; ISSN 0094-243X; Volume 608, no. 1; In English; SPACE TECHNOLOGY and APPLICATIONS INTERNATIONAL FORUM- STAIF 2002, 3-6 Feb 2002, Albuquerque, New Mexico, USA; Copyright The Hypersonic Airplane Space Tether Orbital Launch (HASTOL) Architecture uses a hypersonic airplane (or reusable launch vehicle) to carry a payload from the surface of the Earth to 150 km altitude and a speed of Mach 17. The hypersonic airplane makes a rendezvous with the grapple at the tip of a long, rotating, orbiting space tether boost facility, which picks up the payload from the airplane. Release of the payload at the proper point in the tether rotation boosts the payload into a higher orbit, typically into a Geosynchronous Transfer Orbit (GTO), with lower orbits and Earth escape other options. The HASTOL Tether Boost Facility will have a length of 636 km. Its center of mass will be in a 604 km by 890 km equatorial orbit. It is estimated that by the time of the start of operations of the HASTOL Tether Boost facility in the year 2020, there will be 500 operational spacecraft using the same volume of space as the HASTOL facility. These operational spacecraft would likely be made inoperative by an impact with one of the lines in the multiline HASTOL Hoytether[trademark] and should be avoided. There will also be non-operational spacecraft and large pieces of orbital debris with effective size greater than ﬁve meters in diameter that could cut a number of lines in the HASTOL Hoytether[trademark], and should also be avoided. It is estimated, using two different methods and combining them, that the HASTOL facility will need to make avoidance

28 maneuvers about once every four days if the 500 operational spacecraft and large pieces of orbital debris greater than5min diameter, were each protected by a 2 km diameter miss distance protection sphere. If by 2020, the ability to know the positions of operational spacecraft and large pieces of orbital debris improved to allow a 600 m diameter miss distance protection sphere around each object, then the number of HASTOL facility maneuvers needed drops to one every two weeks. [copyright] 2002 American Institute of Physics. Author (AIP) Artiﬁcial Satellites; Earth Orbits; Hypersonic Vehicles; Mach Number; Reusable Launch Vehicles; Space Debris; Spacecraft Launching

20030032198 NASA Glenn Research Center, Cleveland, OH, USA Advanced Control Surface Seal Development at NASA GRC for 3rd Generation RLV Dunlap, Patrick H., Jr.; Steinetz, Bruce M.; DeMange, Jeffrey J.; Breen, Daniel P.; November 12, 2002; In English; RLV/SOV Airframe Technology Review, 19-22 Nov. 2002, Hampton, VA, USA; Original contains black and white illustrations; Copyright; Avail: CASI; A03, Hardcopy NASA’s Glenn Research Center (GRC) is developing advanced control surface seal technologies for the 3rd Generation RLV program. GRC has designed several new test rigs to simulate the temperatures, pressures, and scrubbing conditions that seals would have to endure during service. A hot compression test rig and hot scrub test rig are being developed to perform tests at temperatures up to 3000 F. Another new test rig allows simultaneous seal ﬂow and scrub tests at room temperature to evaluate changes in seal performance with scrubbing. High temperature ceramic preloading devices are being developed and evaluated as ways to improve seal resiliency. Author Reusable Launch Vehicles; Control Surfaces; Design Analysis; Temperature; Simulation; Pressure; Seals (Stoppers)

20030022661 NASA Glenn Research Center, Cleveland, OH, USA High Altitude Launch for a Practical SSTO Landis, Geoffrey A.; Denis, Vincent; Lyons, Valerie, Technical Monitor; February 2002; In English; Conference on Next Generation Space Transportation, Space Technology and Applications Forum, 2-6 Feb. 2002, Albuquerque, NM, USA; Copyright; Avail: CASI; A02, Hardcopy Existing engineering materials allow the construction of towers to heights of many kilometers. Orbital launch from a high altitude has signiﬁcant advantages over sea-level launch due to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. High-altitude launch sites are particularly advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2% of the initial launch mass. An earlier paper enumerated some of the advantages of high altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25 kilometer altitudes above sea level. The performance increase can be directly translated into increased payload capability to orbit, ranging from 5 to 20% increase in the mass to orbit. For a candidate vehicle with an initial payload fraction of 2% of gross lift-off weight, this corresponds to 31% increase in payload (for 5-km launch altitude) to 122% additional payload (for 25-km launch altitude). Author High Altitude; Towers; Spacecraft Launching; Materials Selection; Trajectory Analysis; Single Stage to Orbit Vehicles

20030020634 NASA Glenn Research Center, Cleveland, OH, USA Towers for Earth Launch Landis, Geoffrey A.; Lyons, Valerie J., Technical Monitor; September 2002; In English; 1st Workshop on Space Elevators, 12-13 Aug. 2002, Seattle, WA., USA Contract(s)/Grant(s): RTOP 755-A4-02; No Copyright; Avail: CASI; A03, Hardcopy This report lists some characteristics of a hypothetical 15 kilometer tower for launching spacecraft, the advantages of launching from high altitude, and some equations pertaining to launch from a 15 kilometer tower. CASI High Altitude; Spacecraft Launching; Towers

20030016552 NASA Marshall Space Flight Center, Huntsville, AL USA Analysis of X-33 Linear Aerospike Plume Induced Base-Heating Environment During Power-Pack Out Wang, Ten-See; Williams, Robert; Droege, Alan; Dagnostino, Mark; Lee, Young-Ching; Douglas, Stan; Turner, James E., Technical Monitor; [2001]; In English; NASA/MSFC Fluids Workshop, 4-5 Apr. 2001, Huntsville, AL, USA; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy

29 The objectives of this viewgraph presentation are to predict the following: (1) dual-engine base-heating at 57% PL at sea level, and (2) dual-engine base-heating during PPO at three ascent abort trajectories. A systematically anchored computational fluid dynamics and heat transfer three-dimensional transfer simulation is being used to study the effect of reduced power levels on base-heating environments during sea level testing and during PPO. Preliminary results show the following: (1) convective heating is higher for the 57% PL than for 100% PL on most of the pillows and flex seals during sea level testing; and (2) convective heating on pillows and flex seals on the ‘off’ engine side is higher than that on the ‘on’ engine side. Derived from text Computational Fluid Dynamics; Convective Heat Transfer; X-33 Reusable Launch Vehicle; Computerized Simulation; Three Dimensional Models

20030015407 NASA Ames Research Center, Moffett Field, CA USA Progress in Unsteady Turbopump Flow Simulations Using Overset Grid Systems Kiris, Cetin C.; Chan, William; Kwak, Dochan; [2002]; In English; NASA MSFC 2002 Workshop, 19-21 Nov. 2002, Huntsville, AL, USA; No Copyright; Avail: CASI; A03, Hardcopy This viewgraph presentation provides information on unsteady flow simulations for the Second Generation RLV (Reusable Launch Vehicle) baseline turbopump. Three impeller rotations were simulated by using a 34.3 million grid points model. MPI/OpenMP hybrid parallelism and MLP shared memory parallelism has been implemented and benchmarked in INS3D, an incompressible Navier-Stokes solver. For RLV turbopump simulations a speed up of more than 30 times has been obtained. Moving boundary capability is obtained by using the DCF module. Scripting capability from CAD geometry to solution is developed. Unsteady flow simulations for advanced consortium impeller/diffuser by using a 39 million grid points model are currently underway. 1.2 impeller rotations are completed. The fluid/structure coupling is initiated. Derived from text Turbine Pumps; Unsteady Flow; Computational Grids; Impellers; Simulation; Rotation

20030014810 NASA Marshall Space Flight Center, Huntsville, AL USA Advanced Guidance and Control Methods for Reusable Launch Vehicles: Test Results Hanson, John M.; Jones, Robert E.; Krupp, Don R.; Fogle, Frank R., Technical Monitor; [2002]; In English; AIAA Guidance, Navigation and Control Conference, 5-8 Aug. 2002, Monterey, CA, USA; Copyright; Avail: CASI; A03, Hardcopy; Distribution as joint owner in the copyright There are a number of approaches to advanced guidance and control (AG&C) that have the potential for achieving the goals of signiﬁcantly increasing reusable launch vehicle (RLV) safety/reliability and reducing the cost. In this paper, we examine some of these methods and compare the results. We brieﬂy introduce the various methods under test, list the test cases used to demonstrate that the desired results are achieved, show an automated test scoring method that greatly reduces the evaluation effort required, and display results of the tests. Results are shown for the algorithms that have entered testing so far. Author Reusable Launch Vehicles; Flight Control; Flight Tests; Entry Guidance (Sts); Control Systems Design; Methodology

20030014743 NASA Ames Research Center, Moffett Field, CA USA Calculation of Supersonic Combustion Using Implicit Schemes Yoon, Seokkwan; Kwak, Dochan, Technical Monitor; [2003]; In English; AIAA 16th Computational Fluid Dynamics Conference, 23-26 Jun. 2003, Orlando, FL, USA; No Copyright; Avail: CASI; A01, Hardcopy One of the technology goals of NASA for advanced space transportation is to develop highly efficient propulsion systems to reduce the cost of payload for space missions. Developments of rockets for the second generation Reusable Launch Vehicle (RLV) in the past several years have been focused on low-cost versions of conventional engines. However, recent changes in the Integrated Space Transportation Program to build a crew transportation vehicle to extend the life of the Space Shuttle ﬂeet might suggest that air-breathing rockets could reemerge as a possible propulsion system for the third generation RLV to replace the Space Shuttle after 2015. The weight of the oxygen tank exceeds thirty percent of the total weight of the Space Shuttle at launch while the payload is only one percent of the total weight. The air-breathing rocket propulsion systems, which consume oxygen in the air, offer clear advantages by making vehicles lighter and more efficient. Experience in the National Aerospace Plane Program in the late 1980s indicates that scramjet engines can achieve high speciﬁc impulse for low hypersonic vehicle speeds. Whether taking a form of Rocket Based Combined Cycle (RBCC) or Turbine Based Combined Cycle (TBCC), the scramjet is an essential mode of operation for air-breathing rockets. It is well known that fuel-air mixing

30 and rapid combustion are of crucial importance for the success of scramjet engines since the spreading rate of the supersonic mixing layer decreases as the Mach number increases. A factored form of the Gauss-Seidel relaxation method has been widely used in hypersonic flow research since its first application to non-equilibrium flows. However, difficulties in stability and convergence have been encountered when there is strong interaction between fluid motion and chemical reaction, such as multiple fuel injection problems. The present paper reports the results from investigation of the effect of modifications to the original algorithm on the performance for multiple injectors. Author Reusable Launch Vehicles; Supersonic Combustion Ramjet Engines; Rocket-Based Combined-Cycle Engines; Hypersonic Vehicles

20030014601 NASA Ames Research Center, Moffett Field, CA USA Thermal Protection System Evaluation Using Arc-jet Flows: Flight Simulation or Research Tool? Stewart, David A.; Venkatapathy, Ethiras, Technical Monitor; Nov. 20, 2002; In English; 105th Annual Meeting and Exposition of the American Ceramic Society, 27-30 Apr. 203, Nashville, TN, USA; No Copyright; Avail: Other Sources; Abstract Only The arc-jet has been used to evaluate thermal protection systems (TPS) and materials for the past forty years. Systems that have been studied in this environmerd include ablators, active, and passive TPS concepts designed for vehicles entering planetary and Earth atmospheres. The question of whether arc-jet flow can simulate a flight environment or is it a research tool that provides an aero-thermodynamic heating environment to obtain critical material properties will be addressed. Stagnation point tests in arc-jets are commonly used to obtain material properties such as mass loss rates, thermal chemical stability data, optical properties, and surface catalytic efficiency. These properties are required in computational fluid dynamic codes to accurately predict the performance of a TPS during flight. Special facilities have been developed at NASA Ames Research Center to approximate the flow environment over the mid-fuselage and body flap regions of proposed space-planes type vehicles. This paper compares flow environments generated in flight over a vehicle with those created over an arc-jet test articles in terms of scale, chemistry, and fluid dynamic properties. Flight experiments are essential in order to validate the material properties obtained from arc-jet tests and used to predict flight performance of any TPS being considered for use on a vehicle entering the Earth atmosphere at hypersonic speed. Author Thermal Protection; Flight Simulation; Arc Jet Engines; Jet Flow; Computational Fluid Dynamics; Aerothermodynamics

20030013452 NASA Marshall Space Flight Center, Huntsville, AL USA NASA 2nd Generation RLV Program Introduction, Status and Future Plans Dumbacher, Dan L.; Smith, Dennis E., Technical Monitor; [2002]; In English; 38th Joint Propulsion Conference, 7-10 Jul. 2002, Indianapolis, IN, USA Report No.(s): AIAA Paper 2002-4570; Copyright; Avail: CASI; A02, Hardcopy; Distribution under U.S. Government purpose rights The Space Launch Initiative (SLI), managed by the Second Generation Reusable Launch Vehicle (2ndGen RLV) Program, was established to examine the possibility of revolutionizing space launch capabilities, deﬁne conceptual architectures, and concurrently identify the advanced technologies required to support a next-generation system. Initial Program funds have been allocated to design, evaluate, and formulate realistic plans leading to a 2nd Gen RLV full-scale development (FSD) decision by 2006. Program goals are to reduce both risk and cost for accessing the limitless opportunities afforded outside Earth’s atmosphere fo civil, defense, and commercial enterprises. A 2nd Gen RLV architecture includes a reusable Earth-to-orbit launch vehicle, an on-orbit transport and return vehicle, ground and ﬂight operations, mission planning, and both on-orbit and on-the-ground support infrastructures All segments of the architecture must advance in step with development of the RLV if a next-generation system is to be fully operational early next decade. However, experience shows that propulsion is the single largest contributor to unreliability during ascent, requires the largest expenditure of time for maintenance, and takes a long time to develop; therefore, propulsion is the key to meeting safety, reliability, and cost goals. For these reasons, propulsion is SLI’s top technology investment area. Author Reusable Launch Vehicles; Spacecraft Launching; NASA Space Programs; Spacecraft Design; Technology Utilization

31 20030007706 NASA Langley Research Center, Hampton, VA USA X-33 Turbulent Aeroheating Measurements and Predictions Hollis, Brian R.; Berry, Scott A.; Horvath, Thomas J.; [2002]; In English; AIAA Atmospheric Flight Mechanics Conference and Exhibit, 5-8 Aug. 2002, Monterey, CA, USA; Original contains color illustrations Report No.(s): AIAA Paper 2002-4700; Copyright; Avail: CASI; A03, Hardcopy; Distribution under U.S. Government purpose rights Measurements and predictions of the X-33 turbulent aeroheating environment have been performed for Mach 6, perfect-gas air conditions. The purpose of this investigation was to compare turbulent aeroheating predictions from two Navier-Stokes codes, LAURA and GASP, with each other and with experimental data in which turbulent flow was produced through either natural transition or forced transition using roughness elements. The wind tunnel testing was conducted at free stream Reynolds numbers of 0.72 x 10(exp 7)/m to 2.4 x 10(exp 7)/m (2.2 x 10(exp 6)/ft to 7.3 x 10(exp 6)/ft) on 0.254 m (10.0-in.) X-33 models at alpha = 40 deg with smooth surfaces, smooth surfaces with discrete trips, and surfaces with simulated bowed thermal protection system panels. Turbulent flow was produced by the discrete trips and bowed panels for all but the lowest Reynolds number, while turbulent flow on the smooth model was produced only at the highest Reynolds number. Turbulent aeroheating levels on each of the three model types were measured using global phosphor thermography and agreed to within the experimental accuracy (+/= 15%) of the test technique. Computations were performed at the wind tunnel free stream conditions using both codes. Turbulent aeroheating levels predicted using the LAURA code were generally 5%-10% lower than those from GASP, although both sets of predictions fell within the experimental accuracy of the wind tunnel data. Author Aerodynamic Heating; Wind Tunnel Tests; Navier-Stokes Equation; Turbulent Flow; Reynolds Number; Thermography; Computer Programs; Computational Fluid Dynamics

20030006793 High Altitude Launch for a Practical SSTO Landis, Geoffrey A.; Denis, Vincent; AIP Conference Proceedings; January 28, 2003; ISSN 0094-243X; Volume 654, no. 1; In English; SPACE TECHNOLOGY and APPLICATIONS INT.FORUM-STAIF 2003: Conf.on Thermophysics in Microgravity; Commercial/Civil Next Generation Space Transportation; Human Space Exploration, 2-5 February 2003, Albuquerque, New Mexico, USA; Copyright Existing engineering materials allow the constuction of towers to heights of many kilometers. Orbital launch from a high altitude has signiﬁcant advantages over sea-level launch due to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. High-altitude launch sites are particularly advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2% of the initial launch mass. An earlier paper enumerated some of the advantages of high altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25 kilometer altitudes above sea level. The performance increase can be directly translated into increased payload capability to orbit, ranging from 5 to 20% increase in the mass to orbit. For a candidate vehicle with an initial payload fraction of 2% of gross lift-off weight, this corresponds to 31% increase in payload (for 5-km launch altitude) to 122% additional payload (for 25-km launch altitude). [copyright] 2003 American Institute of Physics Author (AIP) Cost Estimates; Density; Fractals; Ground Support Systems; High Altitude; Single Stage to Orbit Vehicles; Spacecraft; Structural Engineering; Towers

20030006266 NASA Marshall Space Flight Center, Huntsville, AL USA Propellant Densification for Shuttle: The SSME Perspective Greene, William D.; Boxx, Dayna L.; Tiller, Bruce K., Technical Monitor; May 20, 2002; In English; 38th AIAA/ASME/ SAE/ASEE Joint Propulsion Conference and Exhibit, 7-10 Jul. 2002, Indianapolis, IN, USA Report No.(s): AIAA Paper 2002-3602; Copyright; Avail: CASI; A03, Hardcopy; Distribution under U.S. Government purpose rights The subject of cryogenic propellant densification as a potential upgrade to the Space Shuttle is a subject that has been raised on several occasions over the last decade. Due to advancements in densification technology made as a part of and in parallel to the X-33 project, the subject was raised and studied once again in May 2001. Across the Space Shuttle program people from many disciplines converged to discuss issues and perform trade studies to determine whether densified propellants was worth pursuing. This paper discusses one of these areas, specifically the Space Shuttle Main Engine (SSME). The effects

32 of propellant densification on steady state performance are presented along with discussions of potential transient performance issues. Engine component redesign and retrofit issues are discussed as well the high level requirements to modify the ground test stands to accommodate propellant densification hardware and tanks. And finally, the matter of programmatic concerns enters the subject at hand as part of a discussion of SSME recertification requirements. In the end, potential benefits to SSME performance can be demonstrated and, subject to the densification scheme chosen, there does not appear to insurmountable technical obstacles. Author Space Shuttle Main Engine; Cryogenic Rocket Propellants

20030006264 NASA Langley Research Center, Hampton, VA USA Preliminary Thermal-Mechanical Sizing of Metallic TPS: Process Development and Sensitivity Studies Poteet, Carl C.; Abu-Khajeel, Hasan; Hsu, Su-Yuen; [2002]; In English; 40th AIAA Aerospace Sciences Meeting and Exhibit, 14-17 Jan. 2002, Reno, NV, USA Report No.(s): AIAA Paper 2002-0505; Copyright; Avail: CASI; A03, Hardcopy; Distribution under U.S. Government purpose rights The purpose of this research was to perform sensitivity studies and develop a process to perform thermal and structural analysis and sizing of the latest Metallic Thermal Protection System (TPS) developed at NASA LaRC (Langley Research Center). Metallic TPS is a key technology for reducing the cost of reusable launch vehicles (RLV), offering the combination of increased durability and competitive weights when compared to other systems. Accurate sizing of metallic TPS requires combined thermal and structural analysis. Initial sensitivity studies were conducted using transient one-dimensional finite element thermal analysis to determine the influence of various TPS and analysis parameters on TPS weight. The thermal analysis model was then used in combination with static deflection and failure mode analysis of the sandwich panel outer surface of the TPS to obtain minimum weight TPS configurations at three vehicle stations on the windward centerline of a representative RLV. The coupled nature of the analysis requires an iterative analysis process, which will be described herein. Findings from the sensitivity analysis are reported, along with TPS designs at the three RLV vehicle stations considered. Author Sensitivity Analysis; Thermodynamics; Thermal Analysis; Structural Analysis; Thermal Protection

20030005865 NASA Langley Research Center, Hampton, VA USA Investigation of Fundamental Modeling and Thermal Performance Issues for a Metallic Thermal Protection System Design Blosser, Max L.; [2002]; In English; 40th Aerospace Sciences Meeting and Exhibit, 14-17 Jan. 2002, Reno, NV, USA Report No.(s): AIAA Paper 2002-0503; No Copyright; Avail: CASI; A03, Hardcopy; Distribution under U.S. Government purpose rights A study was performed to develop an understanding of the key factors that govern the performance of metallic thermal protection systems for reusable launch vehicles. A current advanced metallic thermal protection system (TPS) concept was systematically analyzed to discover the most important factors governing the thermal performance of metallic TPS. A large number of relevant factors that influence the thermal analysis and thermal performance of metallic TPS were identified and quantified. Detailed finite element models were developed for predicting the thermal performance of design variations of the advanced metallic TPS concept mounted on a simple, unstiffened structure. The computational models were also used, in an automated iterative procedure, for sizing the metallic TPS to maintain the structure below a specified temperature limit. A statistical sensitivity analysis method, based on orthogonal matrix techniques used in robust design, was used to quantify and rank the relative importance of the various modeling and design factors considered in this study. Results of the study indicate that radiation, even in small gaps between panels, can reduce significantly the thermal performance of metallic TPS, so that gaps should be eliminated by design if possible. Thermal performance was also shown to be sensitive to several analytical assumptions that should be chosen carefully. One of the factors that was found to have the greatest effect on thermal performance is the heat capacity of the underlying structure. Therefore the structure and TPS should be designed concurrently. Author Mathematical Models; Matrices (Mathematics); Performance Prediction; Sensitivity Analysis; Statistical Analysis; Temperature Effects; Thermal Analysis; Thermal Protection

33 20030005683 Alabama Univ., Huntsville, AL USA Improved Re-Configurable Sliding Mode Controller for Reusable Launch Vehicle of Second Generation Addressing Aerodynamic Surface Failures and Thrust Deficiencies Shtessel, Yuri B.; 2002; In English Contract(s)/Grant(s): NAG8-1787; No Copyright; Avail: CASI; A03, Hardcopy In this report we present a time-varying sliding mode control (TV-SMC) technique for reusable launch vehicle (RLV) attitude control in ascent and entry flight phases. In ascent flight the guidance commands Euler roll, pitch and yaw angles, and in entry flight it commands the aerodynamic angles of bank, attack and sideslip. The controller employs a body rate inner loop and the attitude outer loop, which are separated in time-scale by the singular perturbation principle. The novelty of the TVSMC is that both the sliding surface and the boundary layer dynamics can be varied in real time using the PD-eigenvalue assignment technique. This salient feature is used to cope with control command saturation and integrator windup in the presence of severe disturbance or control effector failure, which enhances the robustness and fault tolerance of the controller. The TV-SMC is developed and tuned up for the X-33 sub-orbital technology demonstration vehicle in launch and re-entry modes. A variety of nominal, dispersion and failure scenarios have tested via high fidelity 6DOF simulations using MAVERIC/SLIM simulation software. Author Aerodynamics; Controllers; Reusable Launch Vehicles; Sliding; Thrust; Flight Control

20030003786 NASA Glenn Research Center, Cleveland, OH USA Experimental Evaluation of a Subscale Gaseous Hydrogen/Gaseous Oxygen Coaxial Rocket Injector Smith, Timothy D.; Klem, Mark D.; Breisacher, Kevin J.; Farhangi, Shahram; Sutton, Robert; November 2002; In English; Original contains color illustrations Contract(s)/Grant(s): RTOP 708-73-10 Report No.(s): NASA/TM-2002-211982; E-13652; NAS 1.15:211982; Copyright; Avail: CASI; A03, Hardcopy; Distribution as joint owner in the copyright The next generation reusable launch vehicle may utilize a Full-Flow Stage Combustion (FFSC) rocket engine cycle. One of the key technologies required is the development of an injector that uses gaseous oxygen and gaseous hydrogen as propellants. Gas-gas propellant injection provides an engine with increased stability margin over a range of throttle set points. This paper summarizes an injector design and testing effort that evaluated a coaxial rocket injector for use with gaseous oxygen and gaseous hydrogen propellants. A total of 19 hot-ﬁre tests were conducted up to a chamber pressure of 1030 psia, over a range of 3.3 to 6.7 for injector element mixture ratio. Post-test condition of the hardware was also used to assess injector face cooling. Results show that high combustion performance levels could be achieved with gas-gas propellants and there were no problems with excessive face heating for the conditions tested. Author Injectors; Test Facilities; Coaxial Flow; Hydrogen; Oxygen; Gaseous Rocket Propellants; Rocket Engine Design

20030002800 NASA Marshall Space Flight Center, Huntsville, AL USA, Boeing Co., Huntington Beach, CA USA Launch Vehicle Sizing Benefits Utilizing Main Propulsion System Crossfeed and Project Status Chandler, Frank; Scheiern, M.; Champion, R.; Mazurkivich, P.; Lyles, Garry, Technical Monitor; Jul. 09, 2002; In English; 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 7-10 Jul. 2002, Indianapolis, IN, USA Contract(s)/Grant(s): NAS8-01099 Report No.(s): AIAA Paper 2002-3900; Copyright; Avail: CASI; A02, Hardcopy; Distribution as joint owner in the copyright To meet the goals for a next generation Reusable Launch Vehicle (RLV), a unique propulsion feed system concept was identified using crossfeed between the booster and orbiter stages that could reduce the Two-Stage-to-Orbit (TSTO) vehicle weight and Design, Development, Test and Evaluation (DDT&E) costs by approximately 25%, while increasing safety and reliability. The Main Propulsion System (MPS) crossfeed water demonstration test program addresses all activities required to reduce the risks for the MPS crossfeed system from a Technology Readiness Level (TRL) of 2 to 4 by the completion of testing and analysis by June 2003. During the initial period, that ended in March 2002, a subscale water flow test article was defined. Procurement of a subscale crossfeed check valve was initiated and the specifications for the various components were developed. The fluid transient and pressurization analytical models were developed separately and successfully integrated. The test matrix for the water flow test was developed to correlate the integrated model. A computational fluid dynamics (CFD) model of the crossfeed check valve was developed to assess flow disturbances and internal flow dynamics. Based on the results, the passive crossfeed system concept was very feasible and offered a safe system to be used in an RLV architecture. A water flow test article was designed to accommodate a wide range of flows simulating a number of different types of

34 propellant systems. During the follow-on period, the crossfeed system model will be further refined, the test article will be completed, the water flow test will be performed, and finally the crossfeed system model will be correlated with the test data. This validated computer model will be used to predict the full-scale vehicle crossfeed system performance. Author Computational Fluid Dynamics; Feed Systems; Mathematical Models; Reusable Launch Vehicles; Computerized Simulation; Technology Assessment; Propulsion System Performance

20030002474 NASA Marshall Space Flight Center, Huntsville, AL USA Space Ops 2002: Bringing Space Operations into the 21st Century. Track 3: Operations, Mission Planning and Control. 2nd Generation Reusable Launch Vehicle-Concepts for Flight Operations Hagopian, Jeff; [2002]; In English; No Copyright; Avail: CASI; A02, Hardcopy With the successful implementation of the International Space Station (ISS), the National Aeronautics and Space Administration (NASA) enters a new era of opportunity for scientific research. The ISS provides a working laboratory in space, with tremendous capabilities for scientific research. Utilization of these capabilities requires a launch system capable of routinely transporting crew and logistics to/from the ISS, as well as supporting ISS assembly and maintenance tasks. The Space Shuttle serves as NASA’s launch system for performing these functions. The Space Shuttle also serves as NASA’s launch system for supporting other science and servicing missions that require a human presence in space. The Space Shuttle provides proof that reusable launch vehicles are technically and physically implementable. However, a couple of problems faced by NASA are the prohibitive cost of operating and maintaining the Space Shuttle and its relative inability to support high launch rates. The 2nd Generation Reusable Launch Vehicle (2nd Gen RLV) is NASA’s solution to this problem. The 2nd Gen RLV will provide a robust launch system with increased safety, improved reliability and performance, and less cost. The improved performance and reduced costs of the 2nd Gen RLV will free up resources currently spent on launch services. These resource savings can then be applied to scientific research, which in turn can be supported by the higher launch rate capability of the 2nd Gen RLV. The result is a win - win situation for science and NASA. While meeting NASA’s needs, the 2nd Gen RLV also provides the USA aerospace industry with a commercially viable launch capability. One of the keys to achieving the goals of the 2nd Gen RLV is to develop and implement new technologies and processes in the area of flight operations. NASA’s experience in operating the Space Shuttle and the ISS has brought to light several areas where automation can be used to augment or eliminate functions performed by crew and ground controllers. This experience has also identified the need for new approaches to staffing and training for both crew and ground controllers. This paper provides a brief overview of the mission capabilities provided by the 2nd Gen RLV, a description of NASA’s approach to developing the 2nd Gen RLV, a discussion of operations concepts, and a list of challenges to implementing those concepts. Author Flight Operations; Reusable Launch Vehicles; Space Shuttles; International Space Station; NASA Programs

20030002215 Ohio Univ., USA On-Board Generation of Three-Dimensional Constrained Entry Trajectories Shen, Zuojun; Lu, Ping; Jackson, Scott, Technical Monitor; Jan. 09, 2002; In English; 2002 AIAA GN&C Conference, 5-8 Aug. 2002, Monterey, CA, USA Contract(s)/Grant(s): NAS8-01105; No Copyright; Avail: CASI; A02, Hardcopy A methodology for very fast design of 3DOF entry trajectories subject to all common inequality and equality constraints is developed. The approach make novel use of the well known quasi-equilibrium glide phenomenon in lifting entry as a center piece for conveniently enforcing the inequality constraints which are otherwise difficulty to handle. The algorithm is able to generate a complete feasible 3DOF entry trajectory, given the entry conditions, values of constraint parameters, and ﬁnal conditions in about 2 seconds on a PC. Numerical simulations with the X-33 vehicle model for various entry missions to land at Kennedy Space Center will be presented. Author Trajectories; Algorithms; Three Dimensional Models; Atmospheric Entry

20030001848 NASA Glenn Research Center, Cleveland, OH USA NASA Overview Rusick, Jeff; Fifth Annual Workshop on the Application of Probabilistic Methods for Gas Turbine Engines; October 2002; In English; No Copyright; Avail: CASI; A03, Hardcopy This viewgraph presentation provides information on computer programs for the application of probabilistics in aerospace

35 design. The presentation deﬁnes risk, and lists probabilistic tools which can take risk into account. It then lists current and future examples of NASA aerospace technology program support which utilize probabilistic methods. Current examples include turbofan engines, combustor liners, life estimation, the analysis of rotors and blades, and shuttle ground operations. The 2nd Generation Reusable Launch Vehicle is one of the future areas of support. The presentation also suggests other future uses of probabilistic tools in aerospace design, and optimal methods for their implementation. CASI Computer Programs; Probability Theory; Risk; Aircraft Design; Spacecraft Design; Engine Design

20030001139 NASA Marshall Space Flight Center, Huntsville, AL USA Space Launch Initiative: New Capabilities ... New Horizons Dumbacher, Daniel L.; Nov. 15, 2002; In English; 4th European Conference on Hot Structures and Thermal Protection Systems for Space Vehicles, 25-29 Nov. 2002, Palermo, Italy; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy This paper presents NASA’s Space Launch Initiative (SLI) with new capabilities and new horizons. The topics include: 1) Integrated Space Transportation Plan; 2) SLI: The Work of an Nation; 3) SLI Goals and Status; 4) Composites and Materials; and 5) SLI & DoD/USAF Collaboration. This paper is presented in viewgraph form. CASI Spacecraft Launching; NASA Space Programs; Space Transportation System; Aerospace Engineering; Reusable Launch Vehicles

20030001115 NASA Ames Research Center, Moffett Field, CA USA Automated CFD Parameter Studies on Distributed Parallel Computers Rogers, Stuart E.; Aftosmis, Michael; Pandya, Shishir; Tejnil, Edward; Ahmad, Jasim; Kwak, Dochan, Technical Monitor; [2002]; In English; 16th AIAA Computational Fluid Dynamics Conference, 23-26 Jun. 2003, Orlando, FL, USA; No Copyright; Avail: CASI; A01, Hardcopy The objective of the current work is to build a prototype software system which will automated the process of running CFD jobs on Information Power Grid (IPG) resources. This system should remove the need for user monitoring and intervention of every single CFD job. It should enable the use of many different computers to populate a massive run matrix in the shortest time possible. Such a software system has been developed, and is known as the AeroDB script system. The approach taken for the development of AeroDB was to build several discrete modules. These include a database, a job-launcher module, a run-manager module to monitor each individual job, and a web-based user portal for monitoring of the progress of the parameter study. The details of the design of AeroDB are presented in the following section. The following section provides the results of a parameter study which was performed using AeroDB for the analysis of a reusable launch vehicle (RLV). The paper concludes with a section on the lessons learned in this effort, and ideas for future work in this area. Derived from text Computational Fluid Dynamics; Parallel Computers; Automatic Control; Information Management; Distributed Processing

20030000980 NASA Marshall Space Flight Center, Huntsville, AL USA Manufacturing Process Simulation of Large-Scale Cryotanks Babai, Majid; Phillips, Steven; Griffin, Brian; Munafo, Paul M., Technical Monitor; [2002]; In English; AMPET, 16-18 Sep. 2002, Huntsville, AL, USA; No Copyright; Avail: Other Sources; Abstract Only NASA’s Space Launch Initiative (SLI) is an effort to research and develop the technologies needed to build a second-generation reusable launch vehicle. It is required that this new launch vehicle be 100 times safer and 10 times cheaper to operate than current launch vehicles. Part of the SLI includes the development of reusable composite and metallic cryotanks. The size of these reusable tanks is far greater than anything ever developed and exceeds the design limits of current manufacturing tools. Several design and manufacturing approaches have been formulated, but many factors must be weighed during the selection process. Among these factors are tooling reachability, cycle times, feasibility, and facility impacts. The manufacturing process simulation capabilities available at NASA’s Marshall Space Flight Center have played a key role in down selecting between the various manufacturing approaches. By creating 3-D manufacturing process simulations, the varying approaches can be analyzed in a virtual world before any hardware or infrastructure is built. This analysis can detect and eliminate costly ﬂaws in the various manufacturing approaches. The simulations check for collisions between devices, verify that design limits on joints are not exceeded, and provide cycle times which aid in the development of an optimized process ﬂow. In addition, new ideas and concerns are often raised after seeing the visual representation of a manufacturing

36 process ﬂow. The output of the manufacturing process simulations allows for cost and safety comparisons to be performed between the various manufacturing approaches. This output helps determine which manufacturing process options reach the safety and cost goals of the SLI. Author Reusable Launch Vehicles; Manufacturing; Feasibility; Cryogenic Fluid Storage; Fuel Tanks; Simulation

20030000832 NASA Langley Research Center, Hampton, VA USA Managing MDO Software Development Projects Townsend, J. C.; Salas, A. O.; [2002]; In English; 9th AIAA/ISSMO Symposium on Multidisciplinary Analysis and Optimization, 4-6 Sep. 2002, Atlanta, GA, USA Report No.(s): AIAA Paper 2002-5442; Copyright; Avail: CASI; A03, Hardcopy; Distribution under U.S. Government purpose rights Over the past decade, the NASA Langley Research Center developed a series of ‘grand challenge’ applications demonstrating the use of parallel and distributed computation and multidisciplinary design optimization. All but the last of these applications were focused on the high-speed civil transport vehicle; the final application focused on reusable launch vehicles. Teams of discipline experts developed these multidisciplinary applications by integrating legacy engineering analysis codes. As teams became larger and the application development became more complex with increasing levels of fidelity and numbers of disciplines, the need for applying software engineering practices became evident. This paper briefly introduces the application projects and then describes the approaches taken in project management and software engineering for each project; lessons learned are highlighted. Author Multidisciplinary Design Optimization; Distributed Processing; Computer Networks; Software Engineering; Computer Aided Design; Applications Programs (Computers); Project Management; Parallel Computers

20030000831 NASA Langley Research Center, Hampton, VA USA Discrete Roughness Effects on Shuttle Orbiter at Mach 6 Berry, Scott A.; Hamilton, H. Harris, II; [2002]; In English; 32nd AIAA Fluid Dynamics Conference, 24-27 Jun. 2002, Saint Louis, MO, USA; Original contains color illustrations Report No.(s): AIAA Paper 2002-2744; Copyright; Avail: CASI; A03, Hardcopy; Distribution under U.S. Government purpose rights Discrete roughness boundary layer transition results on a Shuttle Orbiter model in the NASA Langley Research Center 20-Inch Mach 6 Air Tunnel have been reanalyzed with new boundary layer calculations to provide consistency for comparison to other published results. The experimental results were previously obtained utilizing the phosphor thermography system to monitor the status of the boundary layer via global heat transfer images of the Orbiter windward surface. The size and location of discrete roughness elements were systematically varied along the centerline of the 0.0075-scale model at an angle of attack of 40 deg and the boundary layer response recorded. Various correlative approaches were attempted, with the roughness transition correlations based on edge properties providing the most reliable results. When a consistent computational method is used to compute edge conditions, transition datasets for different conﬁgurations at several angles of attack have been shown to collapse to a well-behaved correlation. Author Hypersonic Speed; Scale Models; Space Shuttle Orbiters; Reusable Launch Vehicles; Wind Tunnel Tests

20030000749 NASA Marshall Space Flight Center, Huntsville, AL USA Thermal/Fluid Analysis of a Composite Heat Exchanger for Use on the RLV Rocket Engine Nguyen, Dalton; Twelfth Thermal and Fluids Analysis Workshop; July 2002; In English; Original contains color illustrations; No Copyright; Avail: CASI; A02, Hardcopy As part of efforts to design a regeneratively cooled composite nozzle ramp for use on the reusable vehicle (RLV) rocket engine, an C-SiC composites heat exchanger concept was proposed for thermal performance evaluation. To test the feasibility of the concept, sample heat exchanger panels were made to fit the Glenn Research Center’s cell 22 for testing. Operation of the heat exchanger was demonstrated in a combustion environment with high heat fluxes similar to the RLV Aerospike Ramp. Test measurements were reviewed and found to be valuable for the on going fluid and thermal analysis of the actual RLV composite ramp. Since the cooling fluid for the heat exchanger is water while the RLV Ramp cooling fluid is LH2, fluid and thermal models were constructed to correlate to the specific test set-up. The knowledge gained from this work will be helpful

37 for analyzing the thermal response of the actual RLV Composite Ramp. The coolant thermal properties for the models are taken from test data. The heat exchanger’s cooling performance was analyzed using the Generalized Fluid System Simulation Program (GFSSP). Temperatures of the heat exchanger’s structure were predicted in ﬁnite element models using Patran and Sinda. Results from the analytical models and the tests show that RSC’s heat exchanger satisﬁed the combustion environments in a series of 16 tests. Author Thermal Analysis; Heat Exchangers; Reusable Launch Vehicles; Rocket Engines

20030000734 NASA Ames Research Center, Moffett Field, CA USA Time-Dependent Simulations of Turbopump Flows Kiris, Cetin; Kwak, Dochan; Chan, William; Williams, Robert; Twelfth Thermal and Fluids Analysis Workshop; July 2002; In English; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy Unsteady ﬂow simulations for RLV (Reusable Launch Vehicles) 2nd Generation baseline turbopump for one and half impeller rotations have been completed by using a 34.3 Million grid points model. MLP (Multi-Level Parallelism) shared memory parallelism has been implemented in INS3D, and benchmarked. Code optimization for cash based platforms will be completed by the end of September 2001. Moving boundary capability is obtained by using DCF module. Scripting capability from CAD (computer aided design) geometry to solution has been developed. Data compression is applied to reduce data size in post processing. Fluid/Structure coupling has been initiated. Author Turbine Pumps; Unsteady Flow; Computerized Simulation; Computational Grids; Computational Fluid Dynamics

20030000733 NASA Marshall Space Flight Center, Huntsville, AL USA Stage Separation CFD Tool Development and Evaluation Droege, Alan; Gomez, Reynaldo; Wang, Ten-See; Twelfth Thermal and Fluids Analysis Workshop; July 2002; In English; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy This viewgraph presentation evaluates CFD (Computational Fluid Dynamics) tools for solving stage separation problems. The demonstration and validation of the tools is for a second generation RLV (Reusable Launch Vehicle) stage separation. The flow solvers are: Cart3D; Overflow/Overflow-D; Unic. CASI Computational Fluid Dynamics; Stage Separation; Computer Programs

20030000479 NASA Marshall Space Flight Center, Huntsville, AL USA Space Launch Initiative: New Capabilities - New Horizons Dumbacher, Daniel; Smith, Dennis E., Technical Monitor; Sep. 10, 2002; In English; 10th Japan-US Conference on Composites Materials, 16-18 Sep. 2002, Stanford, CA, USA; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy This paper presents NASA’s Space Launch Initiative (SLI) with new capabilities and new horizons. The topics include: 1) Integrated Space Transportation Plan; 2) SLI: The Work of a Nation; 3) SLI Goals and Status; 4) Composites and Materials; and 5) SLI and DOD/USAF Collaboration. This paper is in viewgraph form. CASI Spacecraft Launching; Space Transportation System; NASA Programs; Reusable Launch Vehicles; Technology Utilization

20030000444 NASA Langley Research Center, Hampton, VA USA Technology Roadmap for Dual-Mode Scramjet Propulsion to Support Space-Access Vision Vehicle Development Cockrell, Charles E., Jr.; Auslender, Aaron H.; Guy, R. Wayne; McClinton, Charles R.; Welch, Sharon S.; [2002]; In English; 11th AIAA/AAAF International Space Planes and Hypersonic Systems, 24 Sep. - 4 Oct. 2002, Orleans, France Report No.(s): AIAA Paper 2002-5188; No Copyright; Avail: CASI; A03, Hardcopy Third-generation reusable launch vehicle (RLV) systems are envisioned that utilize airbreathing and combined-cycle propulsion to take advantage of potential performance beneﬁts over conventional rocket propulsion and address goals of reducing the cost and enhancing the safety of systems to reach earth orbit. The dual-mode scramjet (DMSJ) forms the core of combined-cycle or combination-cycle propulsion systems for single-stage-to-orbit (SSTO) vehicles and provides most of the orbital ascent energy. These concepts are also relevant to two-stage-to-orbit (TSTO) systems with an airbreathing ﬁrst or second stage. Foundation technology investments in scramjet propulsion are driven by the goal to develop efficient Mach 3-15

38 concepts with sufficient performance and operability to meet operational system goals. A brief historical review of NASA scramjet development is presented along with a summary of current technology efforts and a proposed roadmap. The technology addresses hydrogen-fueled combustor development, hypervelocity scramjets, multi-speed ﬂowpath performance and operability, propulsion-airframe integration, and analysis and diagnostic tools. Author Reusable Launch Vehicles; Supersonic Combustion Ramjet Engines; Ascent Propulsion Systems; Systems Analysis; Propulsive Effıciency; Research and Development; Propulsion System Performance

20020094306 NASA Marshall Space Flight Center, Huntsville, AL USA An Automated Method to Compute Orbital Re-entry Trajectories with Heating Constraints Zimmerman, Curtis; Dukeman, Greg; Hanson, John; Fogle, Frank R., Technical Monitor; [2002]; In English; AIAA GNC Conference, 5-8 Aug. 2002, Monterey, CA, USA; No Copyright; Avail: CASI; A01, Hardcopy Determining how to properly manipulate the controls of a re-entering re-usable launch vehicle (RLV) so that it is able to safely return to Earth and land involves the solution of a two-point boundary value problem (TPBVP). This problem, which can be quite difficult, is traditionally solved on the ground prior to flight. If necessary, a nearly unlimited amount of time is available to find the ‘best’ solution using a variety of trajectory design and optimization tools. The role of entry guidance during flight is to follow the pre- determined reference solution while correcting for any errors encountered along the way. This guidance method is both highly reliable and very efficient in terms of onboard computer resources. There is a growing interest in a style of entry guidance that places the responsibility of solving the TPBVP in the actual entry guidance flight software. Here there is very limited computer time. The powerful, but finicky, mathematical tools used by trajectory designers on the ground cannot in general be converted to do the job. Non-convergence or slow convergence can result in disaster. The challenges of designing such an algorithm are numerous and difficult. Yet the payoff (in the form of decreased operational costs and increased safety) can be substantiaL This paper presents an algorithm that incorporates features of both types of guidance strategies. It takes an initial RLV orbital re-entry state and finds a trajectory that will safely transport the vehicle to Earth. During actual flight, the computed trajectory is used as the reference to be flown by a more traditional guidance method. Author Reentry Trajectories; Trajectory Optimization; Reusable Launch Vehicles

20020092099 NASA Marshall Space Flight Center, Huntsville, AL USA Life Cycle Systems Engineering Approach to NASA’s 2nd Generation Reusable Launch Vehicle Thomas, Dale; Smith, Charles; Safie, Fayssal; Kittredge, Sheryl; [2002]; In English; Workshop on Life Cycle Systems Engineering, 6-7 Nov. 2002, Redstone Arsenal, AL, USA; No Copyright; Avail: CASI; A03, Hardcopy The overall goal of the 2nd Generation RLV Program is to substantially reduce technical and business risks associated with developing a new class of reusable launch vehicles. NASA’s specific goals are to improve the safety of a 2nd- generation system by 2 orders of magnitude - equivalent to a crew risk of 1 -in- 10,000 missions - and decrease the cost tenfold, to approximately $1,000 per pound of payload launched. Architecture definition is being conducted in parallel with the maturating of key technologies specifically identified to improve safety and reliability, while reducing operational costs. An architecture broadly includes an Earth-to-orbit reusable launch vehicle, on-orbit transfer vehicles and upper stages, mission planning, ground and flight operations, and support infrastructure, both on the ground and in orbit. The systems engineering approach ensures that the technologies developed - such as lightweight structures, long-life rocket engines, reliable crew escape, and robust thermal protection systems - will synergistically integrate into the optimum vehicle. Given a candidate architecture that possesses credible physical processes and realistic technology assumptions, the next set of analyses address the system’s functionality across the spread of operational scenarios characterized by the design reference missions. The safety/reliability and cost/economics associated with operating the system will also be modeled and analyzed to answer the questions ‘How safe is it?’ and ‘How much will it cost to acquire and operate?’ The systems engineering review process factors in comprehensive budget estimates, detailed project schedules, and business and performance plans, against the goals of safety, reliability, and cost, in addition to overall technical feasibility. This approach forms the basis for investment decisions in the 2nd Generation RLV Program’s risk-reduction activities. Through this process, NASA will continually refine its specialized needs and identify where Defense and commercial requirements overlap those of civil missions. Author NASA Programs; Reusable Launch Vehicles; Systems Engineering; Life Cycle Costs; Technology Utilization

39 20020092015 Ohio Univ., OH USA Direct Fault Tolerant RLV Altitude Control: A Singular Perturbation Approach Zhu, J. J.; Lawrence, D. A.; Fisher, J.; Shtessel, Y. B.; Hodel, A. S.; Lu, P.; Jackson, Scott, Technical Monitor; Jan. 09, 2002; In English; 2002 AIAA GN and C Conference, 5-8 Aug. 2002, Monterey, CA, USA Contract(s)/Grant(s): NAS8-01105; No Copyright; Avail: CASI; A03, Hardcopy In this paper, we present a direct fault tolerant control (DFTC) technique, where by ‘direct’ we mean that no explicit fault identification is used. The technique will be presented for the attitude controller (autopilot) for a reusable launch vehicle (RLV), although in principle it can be applied to many other applications. Any partial or complete failure of control actuators and effectors will be inferred from saturation of one or more commanded control signals generated by the controller. The saturation causes a reduction in the effective gain, or bandwidth of the feedback loop, which can be modeled as an increase in singular perturbation in the loop. In order to maintain stability, the bandwidth of the nominal (reduced-order) system will be reduced proportionally according to the singular perturbation theory. The presented DFTC technique automatically handles momentary saturations and integrator windup caused by excessive disturbances, guidance command or dispersions under normal vehicle conditions. For multi-input, multi-output (MIMO) systems with redundant control effectors, such as the RLV attitude control system, an algorithm is presented for determining the direction of bandwidth cutback using the method of minimum-time optimal control with constrained control in order to maintain the best performance that is possible with the reduced control authority. Other bandwidth cutback logic, such as one that preserves the commanded direction of the bandwidth or favors a preferred direction when the commanded direction cannot be achieved, is also discussed. In this extended abstract, a simplistic example is proved to demonstrate the idea. In the final paper, test results on the high fidelity 6-DOF X-33 model with severe dispersions will be presented. Author Actuators; Altitude Control; Attitude Control; Control Equipment; Failure; Fault Tolerance

20020092014 Ohio Univ., OH USA Autocommander: A Supervisory Controller for Integrated Guidance and Control for the 2nd Generation Reusable Launch Vehicle Fisher, J. E.; Lawrence, D. A.; Zhu, J. J.; Jackson, Scott, Technical Monitor; Jan. 09, 2002; In English; 2002 AIAA GN and C Conference, 5-8 Aug. 2002, Monterey, CA, USA Contract(s)/Grant(s): NAS8-01105; No Copyright; Avail: CASI; A03, Hardcopy This paper presents a hierarchical architecture for integrated guidance and control that achieves risk and cost reduction for NASA’s 2d generation reusable launch vehicle (RLV). Guidance, attitude control, and control allocation subsystems that heretofore operated independently will now work cooperatively under the coordination of a top-level autocommander. In addition to delivering improved performance from a flight mechanics perspective, the autocommander is intended to provide an autonomous supervisory control capability for traditional mission management under nominal conditions, G&C reconfiguration in response to effector saturation, and abort mode decision-making upon vehicle malfunction. This high-level functionality is to be implemented through the development of a relational database that is populated with the broad range of vehicle and mission specific data and translated into a discrete event system model for analysis, simulation, and onboard implementation. A Stateflow Autocoder software tool that translates the database into the Stateflow component of a Matlab/Simulink simulation is also presented. Author Attitude Control; Autocoders; Automatic Control; Mathematical Models; Systems Analysis; Architecture (Computers)

20020092011 Ohio Univ., Athens, OH USA Reusable Launch Vehicle Attitude Control Using a Time-Varying Sliding Mode Control Technique Shtessel, Yuri B.; Zhu, J. Jim; Daniels, Dan; Jackson, Scott, Technical Monitor; Jan. 09, 2002; In English; 2002 AIAA GN and C Conference, 5-8 Aug. 2002, Monterey, CA, USA Contract(s)/Grant(s): NAS8-01105; No Copyright; Avail: CASI; A03, Hardcopy In this paper we present a time-varying sliding mode control (TVSMC) technique for reusable launch vehicle (RLV) attitude control in ascent and entry flight phases. In ascent flight the guidance commands Euler roll, pitch and yaw angles, and in entry flight it commands the aerodynamic angles of bank, attack and sideslip. The controller employs a body rate inner loop and the attitude outer loop, which are separated in time-scale by the singular perturbation principle. The novelty of the TVSMC is that both the sliding surface and the boundary layer dynamics can be varied in real time using the PD-eigenvalue assignment technique. This salient feature is used to cope with control command saturation and integrator windup in the presence of severe disturbance or control effector failure, which enhances the robustness and fault tolerance of the controller.

40 The TV-SMC ascent and descent designs are currently being tested with high ﬁdelity, 6-DOF dispersion simulations. The test results will be presented in the ﬁnal version of this paper. Author Attitude Control; Control Equipment; Fault Tolerance; Reusable Launch Vehicles; Robustness (Mathematics)

20020092010 NASA Marshall Space Flight Center, Huntsville, AL USA A Quantitative Reliability, Maintainability and Supportability Approach for NASA’s Second Generation Reusable Launch Vehicle Safie, Fayssal M.; Daniel, Charles; Kalia, Prince; Smith, Charles A., Technical Monitor; [2002]; In English; Workshop on Life Cycle System Engineering, 6-7 Nov. 2002, Redstone Arsenal, AL, USA; No Copyright; Avail: CASI; A03, Hardcopy The USA National Aeronautics and Space Administration (NASA) is in the midst of a 10-year Second Generation Reusable Launch Vehicle (RLV) program to improve its space transportation capabilities for both cargo and crewed missions. The objectives of the program are to: significantly increase safety and reliability, reduce the cost of accessing low-earth orbit, attempt to leverage commercial launch capabilities, and provide a growth path for manned space exploration. The safety, reliability and life cycle cost of the next generation vehicles are major concerns, and NASA aims to achieve orders of magnitude improvement in these areas. To get these significant improvements, requires a rigorous process that addresses Reliability, Maintainability and Supportability (RMS) and safety through all the phases of the life cycle of the program. This paper discusses the RMS process being implemented for the Second Generation RLV program. Author Reliability; Reusable Launch Vehicles; Safety; Maintainability; Life Cycle Costs; Life (Durability)

20020091983 NASA Glenn Research Center, Cleveland, OH USA Overview of GRC’s Advanced Sensor and Instrumentation Development Mercer, Carolyn; 2001 NASA Seal/Secondary Air System Workshop; October 2002; Volume 1; In English; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy Glenn Research Center develops advanced diagnostic techniques to measure surface and ﬂow properties in research facilities. We support a variety of aerospace propulsion applications: Shuttle, X-33, X-43, ISS, and research engine components: inlets, compressors, combustors, nozzles. We are developing a suite of instrumentation speciﬁcally for 3rd Generation Reusable Launch Vehicle testing. Derived from text Sensors; Reusable Launch Vehicles; Measuring Instruments; Flow Characteristics; General Overviews; Test Facilities

20020091980 NASA Glenn Research Center, Cleveland, OH USA Development and Capabilities of Unique Structural Seal Test Rigs DeMange, Jeffrey J.; Dunlap, Patrick H., Jr.; Steinetz, Bruce M.; Breen, Daniel P.; Robbie, Malcolm G.; 2001 NASA Seal/Secondary Air System Workshop; October 2002; Volume 1; In English; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy High temperature structural seals are necessary in many aerospace and aeronautical applications to minimize any detrimental effects originating from undesired leakage. The NASA Glenn Research Center has been and continues to be a pioneer in the development and evaluation of these types of seals. The current focus for the development of structural seals is for the 3rd Generation Reusable Launch Vehicle (RLV), which is scheduled to replace the current space shuttle system around 2025. Speciﬁc areas of development under this program include seals for propulsion systems (such as the hypersonic air-breathing ISTAR engine concept based upon Rocket Based Combined Cycle technology) and control surface seals for spacecraft including the autonomous rescue X-38 Crew Return Vehicle and the X-37 Space Maneuver Vehicle. Derived from text Seals (Stoppers); Engine Design; Air Breathing Engines; Spacecraft Propulsion; Aerospace Engineering; Reusable Launch Vehicles

20020091979 NASA Glenn Research Center, Cleveland, OH USA Third Generation RLV Structural Seal Development Programs at NASA GRC Dunlap, Patrick H., Jr.; Steinetz, Bruce M.; DeMange, Jeffrey J.; 2001 NASA Seal/Secondary Air System Workshop; October 2002; Volume 1; In English; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy NASA GRC’s work on high temperature structural seal development began in the late 1980’s and early 1990’s under the

41 NASP (National Aero-Space Plane) project. Bruce Steinetz led the in-house propulsion system seal development program and oversaw industry efforts for propulsion system and airframe seal development for this vehicle. a propulsion system seal location in the NASP engine is shown. The seals were located along the edge of a movable panel in the engine to seal the gap between the panel and adjacent engine sidewalls. More recently, we worked with Rocketdyne on high temperature seals for the linear aerospike engine ramps. In applications such as the former X-33 program, multiple aerospike engine modules would be installed side by side on the vehicle. Seals are required in between adjacent engine modules along the edges and base of the engines. The seals have to withstand the extreme temperatures produced byt he thrusters at the top of the ramps while accommodating large deﬂections between adjacent ramps. We came up with several promising seal concepts for this application and shared them with Rocketdyne. Derived from text National Aerospace Plane Program; X-33 Reusable Launch Vehicle; Seals (Stoppers); Systems Engineering; NASA Programs

20020091968 NASA Glenn Research Center, Cleveland, OH USA NASA’s Advanced Space Transportation Program: RTA Project Summary Bartolotta, Paul; McNelis, Nancy; 2001 NASA Seal/Secondary Air System Workshop; October 2002; Volume 1; In English; Original contains color illustrations; No Copyright; Avail: CASI; A02, Hardcopy This paper presents a general overview of NASA’s Revolutionary Turbine Accelerator (RTA) Project. The contents include: 1) Why Turbine Accelerators; 2) Turbine Base Combination Cycle (TBCC) Interrelationships; 3) RTA Bridges the Gap Between Mach 3 and Mach 5; 4) TBCC and RTA; and 6) RTA Project Areas of Emphasis. A short summary of the TBCC and RTA project is also given. This paper is in viewgraph form. CASI Space Transportation; NASA Programs; Turbine Engines; Accelerators; Spacecraft Propulsion; Technology Utilization

20020091933 NASA Marshall Space Flight Center, Huntsville, AL USA, Science Applications International Corp., Huntsville, AL USA Integrated Technology Assessment Center (ITAC) Update Taylor, J. L.; Neely, M. A.; Curran, F. M.; Christensen, E. R.; Escher, D.; Lovell, N.; Morris, Charles, Technical Monitor; Jul. 10, 2002; In English; 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 7-10 Jul. 2002, Indianapolis, IN, USA Contract(s)/Grant(s): NASA Order H-32738-D Report No.(s): AIAA Paper 2002-3550; Copyright; Avail: CASI; A02, Hardcopy; Distribution as joint owner in the copyright The Integrated Technology Assessment Center (ITAC) has developed a flexible systems analysis framework to identify long-term technology needs, quantify payoffs for technology investments, and assess the progress of ASTP-sponsored technology programs in the hypersonics area. For this, ITAC has assembled an experienced team representing a broad sector of the aerospace community and developed a systematic assessment process complete with supporting tools. Concepts for transportation systems are selected based on relevance to the ASTP and integrated concept models (ICM) of these concepts are developed. Key technologies of interest are identified and projections are made of their characteristics with respect to their impacts on key aspects of the specific concepts of interest. Both the models and technology projections are then fed into the ITAC’s probabilistic systems analysis framework in ModelCenter. This framework permits rapid sensitivity analysis, single point design assessment, and a full probabilistic assessment of each concept with respect to both embedded and enhancing technologies. Probabilistic outputs are weighed against metrics of interest to ASTP using a multivariate decision making process to provide inputs for technology prioritization within the ASTP. ITAC program is currently finishing the assessment of a two-stage-to-orbit (TSTO), rocket-based combined cycle (RBCC) concept and a TSTO turbine-based combined cycle (TBCC) concept developed by the team with inputs from NASA. A baseline all rocket TSTO concept is also being developed for comparison. Boeing has recently submitted a performance model for their Flexible Aerospace System Solution for Tomorrow (FASST) concept and the ISAT program will provide inputs for a single-stage-to-orbit (SSTO) TBCC based concept in the near-term. Both of these latter concepts will be analyzed within the ITAC framework over the summer. This paper provides a status update of the ITAC program. Author Technology Assessment; Systems Analysis; Systems Integration; Aerospace Systems; Hypersonics; Spacecraft Propulsion

20020091884 Ohio Univ., OH USA On-Line Computation of a Local Attainable Moment Set for Reusable Launch Vehicles Hodel, A. S.; Shtessel, Yuri B.; Jackson, Scott, Technical Monitor; Jan. 09, 2002; In English; 2002 AIAA GN and C Conference, 5-8 Aug. 2002, Monterey, CA, USA Contract(s)/Grant(s): NAS8-01105; No Copyright; Avail: CASI; A01, Hardcopy

42 Traditional attitude control design of reusable launch vehicles involves independent design of autopilot and control allocation modules [SBB99], [SK97], [SHJ00], [PS00], [Hod00]. Unfortunately, this results in the potential for overly aggressive commands in the autopilot resulting in a loss of performance due to actuator saturation, particularly if the autopilot may suffer from integrator wind-up [HH01]. Author Reusable Launch Vehicles; On-Line Systems; Modules; Attitude Control; Automatic Pilots

20020091874 NASA Marshall Space Flight Center, Huntsville, AL USA Operations Analysis of the 2nd Generation Reusable Launch Vehicle Noneman, Steven R.; Smith, C. A., Technical Monitor; [2002]; In English; Space Ops 2002, 9-12 Oct. 2002, Houston, TX, USA; No Copyright; Avail: CASI; A03, Hardcopy The Space Launch Initiative (SLI) program is developing a second-generation reusable launch vehicle. The program goals include lowering the risk of loss of crew to 1 in 10,000 and reducing annual operations cost to one third of the cost of the Space Shuttle. The SLI missions include NASA, military and commercial satellite launches and crew and cargo launches to the space station. The SLI operations analyses provide an assessment of the operational support and infrastructure needed to operate candidate system architectures. Measures of the operability are estimated (i.e. system dependability, responsiveness, and efficiency). Operations analysis is used to determine the impact of speciﬁc technologies on operations. A conceptual path to reducing annual operations costs by two thirds is based on key design characteristics, such as reusability, and improved processes lowering labor costs. New operations risks can be expected to emerge. They can be mitigated with effective risk management with careful identiﬁcation, assignment, tracking, and closure. SLI design characteristics such as nearly full reusability, high reliability, advanced automation, and lowered maintenance and servicing coupled with improved processes are contributors to operability and large operating cost reductions. Author Reusable Launch Vehicles; Spacecraft Launching; Design Analysis; Maintenance

20020090799 Boeing Phantom Works, Seattle, WA USA Advanced High Temperature Structural Seals Newquist, Charles W.; Verzemnieks, Juris; Keller, Peter C.; Rorabaugh, Michael; Shorey, Mark; October 2002; In English; Original contains color illustrations Contract(s)/Grant(s): F33615-95-D-3203; RTOP 706-85-33; AF Proj. 0014 Report No.(s): NASA/CR-2002-211973; E-13617; NAS 1.26:211973; No Copyright; Avail: CASI; A06, Hardcopy This program addresses the development of high temperature structural seals for control surfaces for a new generation of small reusable launch vehicles. Successful development will contribute signiﬁcantly to the mission goal of reducing launch cost for small, 200 to 300 pound payloads. Development of high temperature seals is mission enabling. For instance, ineffective control surface seals can result in high temperature (3100 F) ﬂows in the elevon area exceeding structural material limits. Longer sealing life will allow use for many missions before replacement, contributing to the reduction of hardware, operation and launch costs. Author Reusable Launch Vehicles; Seals (Stoppers); Research and Development; Elevons; Life (Durability); Product Development; Design Analysis; Thermal Analysis

20020085363 NASA Marshall Space Flight Center, Huntsville, AL USA Main Engine Prototype Development for 2nd Generation RLV RS-83 Vilja, John; Fisher, Mark; Lyles, Garry M., Technical Monitor; [2002]; In English; 1st AIAA/IAF Symposium on Future Reusable Launch Vehicles, 12 Apr. 2002, Huntsville, AL, USA; No Copyright; Avail: CASI; A03, Hardcopy This presentation reports on the NASA project to develop a prototype for RS-83 engine designed for use on reusable launch vehicles (RLV). Topics covered include: program objectives, overview schedule, organizational chart, integrated systems engineering processes, requirement analysis, catastrophic engine loss, maintainability analysis tools, and prototype design analysis. CASI Design Analysis; Engine Design; Reusable Launch Vehicles; Prototypes; Research and Development

43 20020085342 NASA Marshall Space Flight Center, Huntsville, AL USA On-Orbit Propulsion System Project Overview Champion, Robert H.; Lyles, Garry M., Technical Monitor; [2002]; In English; 1st AIAA/IAF Symposium on Future Reusable Launch Vehicles, 11-12 Apr. 2002, Huntsville, AL, USA; No Copyright; Avail: CASI; A03, Hardcopy This conference presentation reports on the progress on NASA’s On-Orbit Propulsion System Project which aims to support the development of second generation reusable launch vehicles (RLV) through advanced research and development and risk reduction activities. Topics covered include: project goals, project accomplishments, risk reduction activities, thruster design and development initiatives, and Aerojet LOX/Ethanol engine development and testing. CASI Reusable Launch Vehicles; Spacecraft Propulsion; Research and Development; Risk

20020082904 NASA Glenn Research Center, Cleveland, OH USA GTX Reference Vehicle Structural Verification Methods and Weight Summary Hunter, J. E.; McCurdy, D. R.; Dunn, P. W.; September 2002; In English; Combustion, Airbreathing Propulsion, Propulsion Systems Hazards and Modelling and Simulation Joint Meeting, 8-12 Apr. 2002, Destin, FL, USA Contract(s)/Grant(s): RTOP 708-90-63 Report No.(s): NASA/TM-2002-211884; E-13200-1; NAS 1.15:211884; No Copyright; Avail: CASI; A03, Hardcopy The design of a single-stage-to-orbit air breathing propulsion system requires the simultaneous development of a reference launch vehicle in order to achieve the optimal mission performance. Accordingly, for the GTX study a 300-lb payload reference vehicle was preliminarily sized to a gross liftoff weight (GLOW) of 238,000 lb. A finite element model of the integrated vehicle/propulsion system was subjected to the trajectory environment and subsequently optimized for structural efficiency. This study involved the development of aerodynamic loads mapped to finite element models of the integrated system in order to assess vehicle margins of safety. Commercially available analysis codes were used in the process along with some internally developed spreadsheets and FORTRAN codes specific to the GTX geometry for mapping of thermal and pressure loads. A mass fraction of 0.20 for the integrated system dry weight has been the driver for a vehicle design consisting of state-of-the-art composite materials in order to meet the rigid weight requirements. This paper summarizes the methodology used for preliminary analyses and presents the current status of the weight optimization for the structural components of the integrated system. Author Hypersonic Aircraft; Reusable Launch Vehicles; Structural Design; Design Analysis; Air Breathing Engines

20020080893 Ohio Aerospace Inst., Brook Park, OH USA Design and Analysis of UHTC Leading Edge Attachment Thomas, David J.; Nemeth, Noel N., Technical Monitor; July 2002; In English; Original contains color illustrations Contract(s)/Grant(s): NCC3-756; RTOP 706-85-31 Report No.(s): NASA/CR-2002-211505; E-13271; NAS 1.26:211505; No Copyright; Avail: CASI; A04, Hardcopy NASA Glenn Research Center was contacted to provide technical support to NASA Ames Research Center in the design and analysis of an ultra high temperature ceramic (UHTC) leading edge. UHTC materials are being considered for reusable launch vehicles because their high temperature capability may allow for un-cooled sharp leading edge designs. While ceramic materials have the design benefit of allowing subcomponents to run hot, they also provide a design challenge in that they invariably must be in contact with cooler subcomponents elsewhere in the structure. NASA Glenn Research Center proposed a modification to an existing attachment design. Thermal and structural analyses of the leading edge assembly were carried out using ABAQUS finite element software. Final results showed that the proposed modifications aided in thermally isolating hot and cold subcomponents and reducing bearing stresses at the attachment location. Author Ceramics; Design Analysis; High Temperature; Mathematical Models; Sharp Leading Edges

20020077952 NASA Langley Research Center, Hampton, VA USA A Procedure for Structural Weight Estimation of Single Stage to Orbit Launch Vehicles (Interim User’s Manual) Martinovic, Zoran N.; Cerro, Jeffrey A.; September 2002; In English Contract(s)/Grant(s): RTOP 706-88-21-03 Report No.(s): NASA/TM-2002-211931; L-18169; NAS 1.15:211931; No Copyright; Avail: CASI; A04, Hardcopy This is an interim user’s manual for current procedures used in the Vehicle Analysis Branch at NASA Langley Research

44 Center, Hampton, Virginia, for launch vehicle structural subsystem weight estimation based on ﬁnite element modeling and structural analysis. The process is intended to complement traditional methods of conceptual and early preliminary structural design such as the application of empirical weight estimation or application of classical engineering design equations and criteria on one dimensional ‘line’ models. Functions of two commercially available software codes are coupled together. Vehicle modeling and analysis are done using SDRC/I-DEAS, and structural sizing is performed with the Collier Research Corp. HyperSizer program. Author Structural Weight; Single Stage to Orbit Vehicles; Mathematical Models; User Manuals (Computer Programs); Launch Vehicles

20020074785 NASA Glenn Research Center, Cleveland, OH USA Characterization of C/SiC Ceramic Matrix Composites (CMCs) with Novel Interface Fiber Coatings Petko, Jeanne F.; Kiser, J. Douglas; Gray, Hugh R., Technical Monitor; [2002]; In English; National Space and Missile Materials Symposium (NSMMS), 24-28 Jun. 2002, Colorado Springs, CO, USA; Original contains color illustrations Contract(s)/Grant(s): RTOP 708-73-25; No Copyright; Avail: CASI; A03, Hardcopy Ceramic Matrix Composites (CMCs) are attractive candidate aerospace materials due to their high specific strength, low density and high temperature capabilities. The National Aeronautics and Space Administration (NASA) is pursuing the use of CMC components in advanced Reusable Launch Vehicle (RLV) propulsion applications. Carbon fiber-reinforced silicon carbide (C/SiC) is the primary material of interest for a variety of RLV propulsion applications. These composites consist of high-strength carbon fibers and a high modulus, oxidation resistant matrix. For RLV propulsion applications, environmental durability will be critical. Two types of carbon fibers were processed with both standard (pyrolytic carbon) and novel (multilayer and pseudoporous) types of interface coatings as part of a study investigating various combinations of constituents. The benefit of protecting the composites with a surface sealant was also investigated. The strengths, durability in oxidizing environments, and microstructures of these developmental composite materials are presented. The novel interface coatings and the surface sealant show promise for protecting the carbon fibers from the oxidizing environment. Author Carbon Fibers; Ceramic Matrix Composites; Silicon Carbides; Coatings; Mechanical Properties; Microstructure

20020073804 NASA Ames Research Center, Moffett Field, CA USA A Geometry-Centered Approach For Two-Stage-To-Orbit Vehicle Synthesis Papadopoulos, P. E.; Reuther, J.; Haimes, B.; Harris, S.; 2nd JANNAF Modeling and Simulation Subcommittee Meeting; April 2002; Volume 1; In English Contract(s)/Grant(s): NAS2-99092; Copyright; Avail: CPIA, 10630 Little Patuxent Pkwy., Suite 202, Columbia, MD 21044-3320 A geometry-centered approach to vehicle analysis and design is demonstrated for classes of two-stage-to-orbit (TSTO) launch configurations. The core capability unique to the new environment is its reliance on a CAD-neutral geometry application program interface. The environment permits a flexible vehicle closure process that includes (a) geometry manipulation, (b) propulsion system selection, (c) weights and sizing, (d) aerodynamic and aero-thermodynamic analysis, and (e) trajectory optimization. Conceptual level analysis tools are integrated into an automated framework using the Phoenix Integration ModelCenter software. The actual vehicle closure is highly iterative with several levels of interface between disciplines. Fundamental to these loops is the geometric manipulation. To provide robust yet accurate and automated geometry, software libraries have been created that allow bidirectional access to and control of a choice of user-chosen CAD kernels. The geometry libraries contain utilities that assemble a vehicle configuration according to prescribed parameter specifications, extract needed geometry parameters, and provide user-controllable water-tight surface tessellations of selected vehicle components. These operations are performed via commands given to the CAD kernel API. The CAPRI programming interface used to broker all the CAD API calls allows the higher-level vehicle assembly and parameter extraction utilities to remain CAD-neutral. The CAPRI interface also provides a consistent (CAD system independent) algorithm for surface tessellations. The geometry-centered philosophy in this analysis and design environment bypasses issues currently associated with CAD output translation, geometry manipulation and grid generation. Author Computer Aided Design; Computer Programs; Aerodynamic Configurations

45 20020073802 SpaceWorks Engineering, Inc., Atlanta, GA USA System Level Uncertainty Assessment for Collaborative RLV Design Charania, A. C.; Bradford, John E.; Olds, John R.; Graham, Matthew; 2nd JANNAF Modeling and Simulation Subcommittee Meeting; April 2002; Volume 1; In English; Copyright; Avail: CPIA, 10630 Little Patuxent Pkwy., Suite 202, Columbia, MD 21044-3320 A collaborative design process utilizing Probabilistic Data Assessment (PDA) is showcased. Given the limitation of financial resources by both the government and industry, strategic decision makers need more than just traditional point designs, they need to be aware of the likelihood of these future designs to meet their objectives. This uncertainty, an ever-present character in the design process, can be embraced through a probabilistic design environment. A conceptual design process is presented that encapsulates the major engineering disciplines for a Third Generation Reusable Launch Vehicle (RLV). Toolsets consist of aerospace industry standard tools in disciplines such as trajectory, propulsion, mass properties, cost, operations, safety, and economics. Variations of the design process are presented that use different fidelities of tools. The disciplinary engineering models are used in a collaborative engineering framework utilizing Phoenix Integration’s ModelCenter and AnalysisServer environment. These tools allow the designer to join disparate models and simulations together in a unified environment wherein each discipline can interact with any other discipline. The design process also uses probabilistic methods to generate the system level output metrics of interest for a RLV conceptual design. The specific system being examined is the Advanced Concept Rocket Engine 92 (ACRE-92) RLV. Previous experience and knowledge (in terms of input uncertainty distributions from experts and modeling and simulation codes) can be coupled with Monte Carlo processes to best predict the chances of program success. Author Computerized Simulation; Probability Theory; Assessments; Design Analysis

20020073535 NASA Ames Research Center, Moffett Field, CA USA Active Wireless Temperature Sensors for Aerospace Thermal Protection Systems Milos, Frank S.; Karunaratne, K.; Arnold, Jim, Technical Monitor; Jul. 10, 2002; In English; SPIE 8th Annual International Symposium on NDE for Health Monitoring and Diagnostics, 2-6 Mar. 2003, San Diego, CA, USA; No Copyright; Avail: Other Sources; Abstract Only Health diagnostics is an area where major improvements have been identiﬁed for potential implementation into the design of new reusable launch vehicles in order to reduce life-cycle costs, to increase safety margins, and to improve mission reliability. NASA Ames is leading the effort to advance inspection and health management technologies for thermal protection systems. This paper summarizes a joint project between NASA Ames and Korteks to develop active wireless sensors that can be embedded in the thermal protection system to monitor sub-surface temperature histories. These devices are thermocouples integrated with radio-frequency identiﬁcation circuitry to enable acquisition and non-contact communication of temperature data through aerospace thermal protection materials. Two generations of prototype sensors are discussed. The advanced prototype collects data from three type-k thermocouples attached to a 2.54-cm square integrated circuit. Author Systems Health Monitoring; Temperature Sensors; Temperature Measurement

20020073119 NASA Glenn Research Center, Cleveland, OH USA Affordable Flight Demonstration of the GTX Air-Breathing SSTO Vehicle Concept Krivanek, Thomas M.; Roche, Joseph M.; Riehl, John P.; Kosareo, Daniel N.; 26th JANNAF Airbreathing Propulsion Subcommittee Meeting; April 2002; Volume 1; In English; No Copyright; Avail: CPIA, 10630 Little Patuxent Pkwy., Suite 202, Columbia, MD 21044-3320 The rocket based combined cycle (RBCC) powered single-stage-to-orbit (SSTO) reusable launch vehicle has the potential to significantly reduce the total cost per pound for orbital payload missions. To validate overall system performance, a flight demonstration must be performed. This paper presents an overview of the first phase of a flight demonstration program for the GTX SSTO vehicle concept. Phase 1 will validate the propulsion performance of the vehicle configuration over the supersonic and hypersonic airbreathing portions of the trajectory. The focus and goal of Phase 1 is to demonstrate the integration and performance of the propulsion system flowpath with the vehicle aerodynamics over the air-breathing trajectory. This demonstrator vehicle will have dual mode ramjet/scramjets, which include the inlet, combustor, and nozzle with geometrically scaled aerodynamic surface outer mold lines (OML) defining the forebody, boundary layer diverter, wings, and tail. The primary objective of this study is to demonstrate propulsion system performance and operability including the ram to scram transition, as well as to validate vehicle aerodynamics and propulsion airframe integration. To minimize overall risk and development cost the effort will incorporate proven materials, use existing turbomachinery in the propellant delivery

46 systems, launch from an existing unmanned remote launch facility, and use basic vehicle recovery techniques to minimize control and landing requirements. A second phase would demonstrate propulsion performance across all critical portions of a space launch trajectory (lift off through transition to all-rocket) integrated with ﬂight-like vehicle systems. Author Air Breathing Engines; Cost Reduction; Engine Airframe Integration; Propulsion System Conﬁgurations; Propulsion System Performance; Reusable Launch Vehicles

20020073118 NASA Glenn Research Center, Cleveland, OH USA SRM-Assisted Trajectory for the GTX Reference Vehicle Riehl, John; Trefny, Charles; Kosareo, Daniel; 26th JANNAF Airbreathing Propulsion Subcommittee Meeting; April 2002; Volume 1; In English; No Copyright; Avail: CPIA, 10630 Little Patuxent Pkwy., Suite 202, Columbia, MD 21044-3320 A goal of the GTX effort has been to demonstrate the feasibility of a single stage- to- orbit (SSTO) vehicle that delivers a small payload to low earth orbit. The small payload class was chosen in order to minimize the risk and cost of development of this revolutionary system. A preliminary design study by the GTX team has resulted in the current conﬁguration that offers considerable promise for meeting the stated goal. The size and gross lift-off weight resulting from scaling the current design to closure however may be considered impractical for the small payload. In lieu of evolving the project’s reference vehicle to a large-payload class, this paper offers the alternative of using solid-rocket motors in order to close the vehicle at a practical scale. This approach offers a near-term, quasi-reusable system that easily evolves to reusable SSTO following subsequent development and optimization. This paper presents an overview of the impact of the addition of SRM’s to the GTX reference vehicle’s performance and trajectory. The overall methods of vehicle modeling and trajectory optimization will also be presented. A key element in the trajectory optimization is the use of the program OTIS 3.10 that provides rapid convergence and a great deal of ﬂexibility to the user. This paper will also present the methods used to implement GTX requirements into OTIS modeling. Author Solid Propellant Rocket Engines; Single Stage to Orbit Vehicles; Low Earth Orbits

20020073112 NASA Ames Research Center, Moffett Field, CA USA Cost Per Pound From Orbit Merriam, M. L.; 26th JANNAF Airbreathing Propulsion Subcommittee Meeting; April 2002; Volume 1; In English; No Copyright; Avail: CPIA, 10630 Little Patuxent Pkwy., Suite 202, Columbia, MD 21044-3320 Traditional studies of Reusable Launch Vehicle (RLV) designs have focused on designs that are completely reusable except for the fuel. This may not be realistic with current technology . An alternate approach is to look at partially reusable launch vehicles. This raises the question of which parts should be reused and which parts should be expendable. One approach is to consider the cost/pound of returning these parts from orbit. With the shuttle, this cost is about three times the cost/pound of launching payload into orbit. A subtle corollary is that RLVs are much less practical for higher orbits, such as the one on which the International Space Station resides, than they are for low earth orbits. Author Reusable Launch Vehicles; Costs; Design Optimization

20020073101 NASA Marshall Space Flight Center, Huntsville, AL USA NASA’s Hypersonic Investment Area Hueter, Uwe; Hutt, John; McClinton, Charles; 26th JANNAF Airbreathing Propulsion Subcommittee Meeting; April 2002; Volume 1; In English; No Copyright; Avail: CPIA, 10630 Little Patuxent Pkwy., Suite 202, Columbia, MD 21044-3320 NASA has established long term goals for access to space. The third generation launch systems are to be fully reusable and operational around 2025. The goal for third-generation launch systems represents significant reduction in cost and improved safety over the current first generation system. The Advanced Space Transportation Office (ASTP) at NASA s Marshall Space Flight Center (MSFC) has the agency lead to develop space transportation technologies. Within ASTP, under the Hypersonic Investment Area (HIA), third generation technologies are being pursued in the areas of propulsion, airframe, integrated vehicle health management (IVHM), avionics, power, operations and system analysis. These technologies are being matured through research and both ground and flight-testing. This paper provides an overview of the HIA program plans and recent accomplishments. Author Flight Tests; Hypersonics; Systems Analysis; Space Transportation; Safety; Reusable Launch Vehicles

47 20020072989 NASA Glenn Research Center, Cleveland, OH USA Oxidation Behavior of Copper Alloy Candidates for Rocket Engine Applications (Technical Poster) Ogbuji, Linus U. J.; Humphrey, Donald H.; Barrett, Charles A.; Greenbauer-Seng, Leslie, Technical Monitor; Gray, Hugh R., Technical Monitor; [2002]; In English; TMS IV Meeting, 6-10 Oct. 2002, Columbus, OH, USA; Original contains color illustrations Contract(s)/Grant(s): NAS3-98008; RTOP 721-20-XX; No Copyright; Avail: CASI; A01, Hardcopy A rocket engine’s combustion chamber is lined with material that is highly conductive to heat in order to dissipate the huge thermal load (evident in a white-hot exhaust plume). Because of its thermal conductivity copper is the best choice of liner material. However, the mechanical properties of pure copper are inadequate to withstand the high stresses, hence, copper alloys are needed in this application. But copper and its alloys are prone to oxidation and related damage, especially ‘blanching’ (an oxidation-reduction mode of degradation). The space shuttle main engine combustion chamber is lined with a Cu-Ag-Zr alloy, ‘NARloy-Z’, which exhibits blanching. A superior liner is being sought for the next generation of RLVs (Reusable Launch Vehicles) It should have improved mechanical properties and higher resistance to oxidation and blanching, but without substantial penalty in thermal conductivity. GRCop84, a Cu-8Cr-4Nb alloy (Cr2Nb in Cu matrix), developed by NASA Glenn Research Center (GRC) and Case Western Reserve University, is a prime contender for RLV liner material. In this study, the oxidation resistance of GRCop-84 and other related/candidate copper alloys are investigated and compared Author Rocket Engines; Oxidation; Copper Alloys; Oxidation Resistance

20020072848 NASA Glenn Research Center, Cleveland, OH USA Characterization of C/SiC Ceramic Matrix Composites (CMCs) with Novel Interface Fiber Coatings Petko, Jeanne F.; Kiser, James Douglas; Verilli, Mike; McCue, Terry; Gray, Hugh R., Technical Monitor; [2002]; In English; 26th Annual Conference on Materials and Structures, 28-31 Jan. 2002, Cape Canaveral, FL, USA Contract(s)/Grant(s): RTOP 708-73-25; No Copyright; Avail: CASI; A03, Hardcopy Ceramic Matrix Composites (CMCs) are attractive candidate materials in the aerospace industry due to their high specific strength, low density and higher temperature capabilities. The National Aeronautics and Space Administration (NASA) is pursuing the use of CMC components in advanced Reusable Launch Vehicle (RLV) propulsion applications. Carbon fiber-reinforced silicon carbide (C/SiQ is the primary material of interest for a variety of RLV propulsion applications. These composites offer high-strength carbon fibers and a high modulus, oxidation-resistant matrix. For comparison, two types of carbon fibers were processed with novel types of interface coatings (multilayer and pseudoporous). For RLV propulsion applications, environmental durability will be critical. The coatings show promise of protecting the carbon fibers from the oxidizing environment. The strengths and microstructures of these composite materials are presented. Author Ceramic Matrix Composites; Fiber-Matrix Interfaces; Carbon Fibers; Silicon Carbides; Reusable Launch Vehicles; High Strength

20020070918 NASA Glenn Research Center, Cleveland, OH USA SRM-Assisted Trajectory for the GTX Reference Vehicle Riehl, John; Trefny, Charles; Kosareo, Daniel, Technical Monitor; June 2002; In e; Combustion, Airbreathing Propulsion, Propulsion Systems Hazards, and Modelling and Simulation Subcommittees Joint Meeting, 8-12 Apr. 2002, Destin, FL, USA Contract(s)/Grant(s): RTOP 708-90-63 Report No.(s): NASA/TM-2002-211599; E-13386; NAS 1.15:211599; No Copyright; Avail: CASI; A03, Hardcopy A goal of the GTX effort has been to demonstrate the feasibility of a single stage-to-orbit (SSTO) vehicle that delivers a small payload to low earth orbit. The small payload class was chosen in order to minimize the risk and cost of development of this revolutionary system. A preliminary design study by the GTX team has resulted in the current conﬁguration that offers considerable promise for meeting the stated goal. The size and gross lift-off weight resulting from scaling the current design to closure however may be considered impractical for the small payload. In lieu of evolving the project’ reference vehicle to a large-payload class, this paper offers the alternative of using solid-rocket motors in order to close the vehicle at a practical scale. This approach offers a near-term, quasi-reusable system that easily evolves to reusable SSTO following subsequent development and optimization. This paper presents an overview of the impact of the addition of SRM’s to the GTX reference vehicle#s performance and trajectory. The overall methods of vehicle modeling and trajectory optimization will also be presented. A key element in the trajectory optimization is the use of the program OTIS 3.10 that provides rapid convergence

48 and a great deal of ﬂexibility to the user. This paper will also present the methods used to implement GTX requirements into OTIS modeling. Author Solid Propellant Rocket Engines; Launch Vehicles; Hypersonic Combustion; Trajectory Optimization

20020070891 NASA Glenn Research Center, Cleveland, OH USA Fuel Cell Activities at the NASA Glenn Research Center Kohout, Lisa L.; Lyons, Valerie, Technical Monitor; [2002]; In English; Ohio Technical College, 10 Jul. 2002, Cleveland, OH, USA; No Copyright; Avail: CASI; A03, Hardcopy Fuel cells have a long history in space applications and may have potential application in aeronautics as well. A fuel cell is an electrochemical energy conversion device that directly transforms the chemical energy of a fuel and oxidant into electrical energy. Alkaline fuel cells have been the mainstay of the U.S. space program, providing power for the Apollo missions and the Space Shuttle. However, Proton Exchange Membrane (PEM) fuel cells offer potential beneﬁts over alkaline systems and are currently under development for the next generation Reusable Launch Vehicle (RLV). Furthermore, primary and regenerative systems utilizing PEM technology are also being considered for future space applications such as surface power and planetary aircraft. In addition to these applications, the NASA Glenn Research Center is currently studying the feasibility of the use of both PEM and solid oxide fuel cells for low- or zero-emission electric aircraft propulsion. These types of systems have potential applications for high altitude environmental aircraft, general aviation and commercial aircraft, and high attitude airships. NASA Glenn has a unique set of capabilities and expertise essential to the successful development of advanced fuel cell power systems for space and aeronautics applications. NASA Glenn’s role in past fuel cell development programs as well as current activities to meet these new challenges will be presented Author Aircraft Engines; Fuel Cells; Research and Development; Spacecraft Power Supplies; Energy Conversion Effıciency; Aerospace Systems

20020070797 Mississippi Univ., University, MS USA NASA EPSCoR Preparation Grant Sukanek, Peter C.; [2002]; In English; American Society of Photogrammetry and Remote Sensing, May 2002, Washington, DC, USA Contract(s)/Grant(s): NCC5-405; No Copyright; Avail: CASI; A02, Hardcopy The NASA EPSCoR project in Mississippi involved investigations into three areas of interest to NASA by researchers at the four comprehensive universities in the state. These areas involved: (1) Noninvasive Flow Measurement Techniques, (2) Spectroscopic Exhaust Plume Measurements of Hydrocarbon Fueled Rocket Engines and (3) Integration of Remote Sensing and GIS data for Flood Forecasting on the Mississippi Gulf Coast. Each study supported a need at the Stennis Space Center in Mississippi. The ﬁrst two addressed needs in rocket testing, and the third, in commercial remote sensing. Students from three of the institutions worked with researchers at Stennis Space Center on the projects. Author Flow Measurement; Computational Fluid Dynamics; Reusable Launch Vehicles; Systems Health Monitoring; Exhaust Flow Simulation; Flame Spectroscopy; Flood Predictions; Hydrology Models

20020070788 Georgia Inst. of Tech., Atlanta, GA USA Permeability of Impacted Coated Composite Laminates Johnson, W. S.; Findley, Benjamin; August 2002; In English Contract(s)/Grant(s): NAG8-1819; GIT Proj. 1806675 Report No.(s): E-18-675; No Copyright; Avail: CASI; A07, Hardcopy Composite materials are being considered for use on future generations of Reusable Launch Vehicles (RLVs) for both fuel tanks and fuel feedlines. Through the use of composite materials NASA can reduce the overall weight of the vehicle dramatically. This weight savings can then be translated into an increase in the weight of payload sent into orbit, reducing the cost per pound of payload. It is estimated that by switching to composite materials for fuel tanks the weight of the tanks can be reduced by 40 percent, which translates to a total vehicle weight savings of 14 percent. In this research, carbon/epoxy composites were studied for fuel feedline applications. There are concerns about using composite materials for feedlines and fuel tanks because these materials are extremely vulnerable to impact in the form of inadvertent bumping or dropped tools both during installation and maintenance. Additionally, it has been found that some of the sample feedlines constructed have had

49 leaks, and thus there may be a need to seal preexisting leaks in the composite prior to usage. Derived from text Composite Materials; Carbon Fibers; Coatings; Epoxy Matrix Composites; Fiber Composites

20020070521 Wyoming Univ., Laramie, WY USA Reduction of Base Drag on Launch Vehicles Naughton, Jonathan W.; [2002]; In English Contract(s)/Grant(s): NAG4-208 Report No.(s): UWAL-2002-02; No Copyright; Avail: CASI; A03, Hardcopy Current reentry vehicle designs exhibit large amounts of base drag due to large base areas. These large base areas can arise from the integration of the propulsion system (X-33) or control surface placement (X-38). Large base drag limits the vehicle’s cross-range capability and causes a large glide-slope angle. Fortunately, there appears to be a possible means of lowering the base drag on these vehicles. Based on early work on the subsonic aerodynamics of lifting bodies, it appears that the addition of small amounts of viscous fore-body drag can produce a significant reduction in base drag. Recent work suggests that this phenomenon also occurs in the transonic and supersonic flight regimes. This report summarizes a study designed to demonstrate the reduction of base drag through the addition of fore-body viscous drag. The present study has focused on the measurement of viscous fore-body drag and the demonstration of the relationship between fore-body viscous drag and base drag at Reynolds Numbers up to 2.5 x 10(exp 6). The results of the present work do not conclusively demonstrate that viscous fore-body drag reduces base drag. The apparent contradictory results of the present study are attributed to the different geometry used in the present study. However, the results suggest that the increased boundary layer thickness at separation caused by larger fore-body viscous drag somehow affects the vortex structure in the wake thereby reducing the base drag. More research is required to confirm this postulated mechanism. Author Base Flow; Drag Reduction; Reentry Vehicles; Viscous Drag; Aerodynamic Drag; Forebodies

20020068832 Louisiana State Univ., Shreveport, LA USA Operations Analysis of Space Shuttle System Sarker, Bhaba R.; Research Reports: 2001 NASA/ASEE Summer Faculty Fellowship Program; July 2002; In English Contract(s)/Grant(s): NAG8-1786; No Copyright; Avail: CASI; A01, Hardcopy The space science program at NASA since 1950’s has gone through stages of development and implementation in rocketry and scientific advancement. It has become the nation’s largest scientific institution of research and innovation for space exploration and military research, and overall a pride of the nation at an exorbitant cost. After placing man on moon, space shuttle program at NASA is an on-going project with a high success rate at an average cost of about $450m per flight. A future endeavor from both government and private sectors needs to be undertaken for commercialization of this expensive mission. In order to attract private enterprises, it needs to boost up its operations with better technology at lower cost to cope with rapid changes in scientific advancement and economic competition. Thus, a second-generation reusable launch vehicle (2GRLV) will play a major role in future operation of NASA centers. Envisioning this potential way of saving the program by reducing the cost, NASA is currently managing an innovative program called the Space Launch Initiative (SLI) to develop key technologies that will support the development of second-generation reusable launch vehicles (RLV) which will be more economical and safer and reliable than the existing space shuttle system. The selection of which technologies to fund for further development is being based on their likelihood to contribute to providing cost reduction or safety improvements. It is envisioned that in the 2006 timeframe, NASA will make a decision as to whether or not to commit to the replacement of the current space shuttle system with a new RLV. The decision to proceed with a new RLV will be partly based on the likelihood that the new system will be better than the existing space shuttle. Government and private entrepreneurs are currently considering four different types of RLV projects: commercial programs, government programs, international concepts, and X Prize competitors. NASA has already commissioned a series of X-programs to study the future RLV program. Today, NASA decision makers need analytical tools to help determine which technologies to fund the development of this technology. In the 2006 timeframe, these same decision makers will need analytical tools to evaluate and compare various RLV architectures, including the existing space shuttle so as to make the best decision for whether or not to proceed with the development of a new RLV, and if so, then which one. A study is conducted here to establish ground level knowledge from the historical data and expertise experiences of the field personnel. Such information is compiled in the form of mission statement, goals, space shuttle operations, payloads and cargo constraints, resource constraints, and bottlenecks to the enhancement vector. Derived from text Reusable Launch Vehicles; Research and Development; Space Shuttles; Assessments; Project Planning; Operations Research

50 20020068823 College of New Jersey, Pomona, NJ USA An Overview of PRA with Applications to Aerospace Systems Navard, Sharon E.; Research Reports: 2001 NASA/ASEE Summer Faculty Fellowship Program; July 2002; In English; No Copyright; Avail: CASI; A01, Hardcopy Probabilistic Risk Assessment (PRA) is a systematic process for evaluating the probabilities and consequences of undesirable events that can occur in a process or system along with providing a measure of the uncertainty associated with these probability estimates. In the past it was looked at with suspicion by many at NASA, perhaps because of bad experiences with unsuccessful quantitative methods during the Apollo era, but since the Challenger accident NASA has mandated that it be used, and it has been very successful. With NASA’s new ‘faster, better, cheaper’ philosophy, it is vital that a tool be in place that can help to achieve these goals in the reliability area. This paper describes the history of PRA, gives examples of its aerospace applications to date, and gives suggestions for how it can be used in the future, both for space shuttle upgrades and for totally new technologies such as the Second Generation Reusable Launch Vehicle. Author Assessments; Probability Theory; Risk

20020068806 Alabama Agricultural and Mechanical Univ., Normal, AL USA Second Generation Reusable Launch Vehicle Development and Global Competitiveness of US Space Transportation Industry: Critical Success Factors Assessment Enyinda, Chris I.; Research Reports: 2001 NASA/ASEE Summer Faculty Fellowship Program; July 2002; In English; No Copyright; Avail: CASI; A01, Hardcopy In response to the unrelenting call in both public and private sectors fora to reduce the high cost associated with space transportation, many innovative partially or fully RLV (Reusable Launch Vehicles) designs (X-34-37) were initiated. This call is directed at all levels of space missions including scientific, military, and commercial and all aspects of the missions such as nonrecurring development, manufacture, launch, and operations. According to Wertz, tbr over thirty years, the cost of space access has remained exceedingly high. The consensus in the popular press is that to decrease the current astronomical cost of access to space, more safer, reliable, and economically viable second generation RLVs (SGRLV) must be developed. Countries such as Brazil, India, Japan, and Israel are now gearing up to enter the global launch market with their own commercial space launch vehicles. NASA and the US space launch industry cannot afford to lag behind. Developing SGRLVs will immeasurably improve the US’s space transportation capabilities by helping the US to regain the global commercial space markets while supporting the transportation capabilities of NASA’s space missions, Developing the SGRLVs will provide affordable commercial space transportation that will assure the competitiveness of the US commercial space transportation industry in the 21st century. Commercial space launch systems are having difficulty obtaining financing because of the high cost and risk involved. Access to key financial markets is necessary for commercial space ventures. However, public sector programs in the form of tax incentives and credits, as well as loan guarantees are not yet available. The purpose of this paper is to stimulate discussion and assess the critical success factors germane for RLVs development and US global competitiveness. Derived from text Cost Reduction; Reusable Launch Vehicles; Space Commercialization; Research and Development; Commercial Spacecraft; NASA Programs

20020068708 NASA Ames Research Center, Moffett Field, CA USA Progress in Unsteady Turbopump Flow Simulations Kiris, Cetin C.; Chan, William; Kwak, Dochan; Williams, Robert; [2002]; In English; JANNAF 2002 Meeting, 8-12 Apr. 2002, Destin, FL, USA; No Copyright; Avail: CASI; A03, Hardcopy This viewgraph presentation discusses unsteady ﬂow simulations for a turbopump intended for a reusable launch vehicle (RLV). The simulation process makes use of computational grids and parallel processing. The architecture of the parallel computers used is discussed, as is the scripting of turbopump simulations. CASI Turbine Pumps; Unsteady Flow; Computational Grids; Simulation; Systems Analysis

20020068450 NASA Marshall Space Flight Center, Huntsville, AL USA Estimating the Cost of NASA’s Space Launch Initiative: How SLI Cost Stack Up Against the Shuttle Hamaker, Joseph H.; Roth, Axel, Technical Monitor; [2002]; In English; SSCAG Meeting, 11-13 Jun. 2002, Frascati, Italy; No Copyright; Avail: Other Sources; Abstract Only

51 NASA is planning to replace the Space Shuttle with a new completely reusable Second Generation Launch System by approximately 2012. Numerous contracted and NASA in-house Space Transportation Architecture Studies and various technology maturation activities are proceeding and have resulted in scores of competing architecture conﬁgurations being proposed. Life cycle cost is a key discriminator between all these various concepts. However, the one obvious analogy for costing purposes remains the current Shuttle system. Are there credible reasons to believe that a second generation reusable launch system can be accomplished at less cost than the Shuttle? The need for a credible answer to this question is critical. This paper reviews the cost estimating approaches being used by the contractors and the government estimators to address this issue and explores the rationale behind the numbers. Author Space Shuttles; Life Cycle Costs; Cost Estimates; Reusable Launch Vehicles

20020067630 NASA Marshall Space Flight Center, Huntsville, AL USA Magnetohydrodynamic Augmented Propulsion Experiment Litchford, Ron J.; Cole, John; Lineberry, John; Chapman, Jim; Schmidt, Harold; Cook, Stephen, Technical Monitor; [2002]; In English; 33rd Plasmadynamics and Lasers Conference, 20-23 May 2002, Maui, HI, USA; No Copyright; Avail: Other Sources; Abstract Only A fundamental obstacle to routine space access is the specific energy limitations associated with chemical fuels. In the case of vertical take-off, the high thrust needed for vertical liftoff and acceleration to orbit translates into power levels in the 10 GW range. Furthermore, useful payload mass fractions are possible only if the exhaust particle energy (i.e., exhaust velocity) is much greater than that available with traditional chemical propulsion. The electronic binding energy released by the best chemical reactions (e.g., LOX/LH2 for example, is less than 2 eV per product molecule (approx. 1.8 eV per H2O molecule), which translates into particle velocities less than 5 km/s. Useful payload fractions, however, will require exhaust velocities exceeding 15 km/s (i.e., particle energies greater than 20 eV). As an added challenge, the envisioned hypothetical RLV (reusable launch vehicle) should accomplish these amazing performance feats while providing relatively low acceleration levels to orbit (2-3g maximum). From such fundamental considerations, it is painfully obvious that planned and current RLV solutions based on chemical fuels alone represent only a temporary solution and can only result in minor gains, at best. What is truly needed is a revolutionary approach that will dramatically reduce the amount of fuel and size of the launch vehicle. This implies the need for new compact high-power energy sources as well as advanced accelerator technologies for increasing engine exhaust velocity. Electromagnetic acceleration techniques are of immense interest since they can be used to circumvent the thermal limits associated with conventional propulsion systems. This paper describes the Magnetohydrodynamic Augmented Propulsion Experiment (MAPX) being undertaken at NASA Marshall Space Flight Center (MSFC). In this experiment, a 1-MW arc heater is being used as a feeder for a 1-MW magnetohydrodynamic (MHD) accelerator. The purpose of the experiment is to demonstrate that an MHD accelerator can be an effective augmentation system for increasing engine exhaust velocity. More specifically, the experiment is intended to show that electromagnetic effects are effective at producing flow acceleration whereas electrothermal effects do not cause unacceptable heating of the working fluid. The MHD accelerator was designed as an externally diagonalized segmented Faraday channel, which will be inserted into an existing 2-tesla electromagnet. This allows the external power to be connected through two terminals thereby minimizing the complexity and cost associated with powering each segment independently. The design of the accelerator and other components in the flow path has been completed and fabrication activities are underway. This paper provides a full description of MAPX including performance analysis, design, and test plans, and current status. Author Magnetohydrodynamics; Electromagnetic Acceleration; Vertical Takeoff; Arc Heating; Magnetohydrodynamic Generators; Thrust Augmentation; Propulsion System Configurations

20020066164 NASA Marshall Space Flight Center, Huntsville, AL USA A Plan for Advanced Guidance and Control Technology for 2nd Generation Reusable Launch Vehicles Hanson, John M.; Fogle, Frank, Technical Monitor; [2002]; In English; No Copyright; Avail: CASI; A02, Hardcopy Advanced guidance and control (AG&C) technologies are critical for meeting safety/reliability and cost requirements for the next generation of reusable launch vehicle (RLV). This becomes clear upon examining the number of expendable launch vehicle failures in the recent past where AG&C technologies would have saved a RLV with the same failure mode, the additional vehicle problems where this technology applies, and the costs associated with mission design with or without all these failure issues. The state-of-the-art in guidance and control technology, as well as in computing technology, is at the point where we can took to the possibility of being able to safely return a RLV in any situation where it can physically be recovered.

52 This paper outlines reasons for AG&C, current technology efforts, and the additional work needed for making this goal a reality. Author Reusable Launch Vehicles; Flight Control; Failure Modes; Mission Planning

20020065568 NASA Marshall Space Flight Center, Huntsville, AL USA An Automated Method to Compute Orbital Re-Entry Trajectories with Heating Constraints Zimmerman, Curtis; Dukeman, Greg; Hanson, John; Fogle, Frank R., Technical Monitor; [2002]; In English; No Copyright; Avail: CASI; A03, Hardcopy Determining how to properly manipulate the controls of a re-entering re-usable launch vehicle (RLV) so that it is able to safely return to Earth and land involves the solution of a two-point boundary value problem (TPBVP). This problem, which can be quite difficult, is traditionally solved on the ground prior to flight. If necessary, a nearly unlimited amount of time is available to find the ‘best’ solution using a variety of trajectory design and optimization tools. The role of entry guidance during flight is to follow the pre-determined reference solution while correcting for any errors encountered along the way. This guidance method is both highly reliable and very efficient in terms of onboard computer resources. There is a growing interest in a style of entry guidance that places the responsibility of solving the TPBVP in the actual entry guidance flight software. Here there is very limited computer time. The powerful, but finicky, mathematical tools used by trajectory designers on the ground cannot in general be made to do the job. Nonconvergence or slow convergence can result in disaster. The challenges of designing such an algorithm are numerous and difficult. Yet the payoff (in the form of decreased operational costs and increased safety) can be substantial. This paper presents an algorithm that incorporates features of both types of guidance strategies. It takes an initial RLV orbital re-entry state and finds a trajectory that will safely transport the vehicle to a Terminal Area Energy Management (TAEM) region. During actual flight, the computed trajectory is used as the reference to be flown by a more traditional guidance method. Author Reusable Launch Vehicles; Reentry Trajectories; Trajectory Optimization; Design Optimization; Boundary Value Problems; Flight Control

20020065525 Science Applications International Corp., Torrance, CA USA A Thermal Management Systems Model for the NASA GTX RBCC Concept Traci, Richard M.; Farr, John L., Jr.; Laganelli, Tony; Walker, James, Technical Monitor; June 2002; In English; Original contains color illustrations Contract(s)/Grant(s): NAS3-99147; RTOP 708-90-63 Report No.(s): NASA/CR-2002-211587; NAS 1.26:211587; E-13372; SAIC-284-002-017; No Copyright; Avail: CASI; A05, Hardcopy The Vehicle Integrated Thermal Management Analysis Code (VITMAC) was further developed to aid the analysis, design, and optimization of propellant and thermal management concepts for advanced propulsion systems. The computational tool is based on engineering level principles and models. A graphical user interface (GUI) provides a simple and straightforward method to assess and evaluate multiple concepts before undertaking more rigorous analysis of candidate systems. The tool incorporates the Chemical Equilibrium and Applications (CEA) program and the RJPA code to permit heat transfer analysis of both rocket and air breathing propulsion systems. Key parts of the code have been validated with experimental data. The tool was speciﬁcally tailored to analyze rocket-based combined-cycle (RBCC) propulsion systems being considered for space transportation applications. This report describes the computational tool and its development and veriﬁcation for NASA GTX RBCC propulsion system applications. Author Aerothermodynamics; Rocket-Based Combined-Cycle Engines; Temperature Control; Aerospace Planes; Mathematical Models; Applications Programs (Computers)

20020063487 NASA Ames Research Center, Moffett Field, CA USA The NASA Integrated Vehicle Health Management Technology Experiment for X-37 Schwabacher, Mark; Samuels, Jeff; Brownston, Lee; Clancy, Daniel, Technical Monitor; [2002]; In English; Conference on Component and Systems Diagnosis, Prognosis and Health Management II, SPIE AeroSense Meeting, 3 Apr. 2002, Orlando, FL, USA; No Copyright; Avail: CASI; A03, Hardcopy The NASA Integrated Vehicle Health Management (IVHM) Technology Experiment for X-37 was intended to run IVHM

53 software on-board the X-37 spacecraft. The X-37 is intended to be an unpiloted vehicle that would orbit the Earth for up to 21 days before landing on a runway. The objectives of the experiment were to demonstrate the benefits of in-flight IVHM to the operation of a Reusable Launch Vehicle, to advance the Technology Readiness Level of this IVHM technology within a flight environment, and to demonstrate that the IVHM software could operate on the Vehicle Management Computer. The scope of the experiment was to perform real-time fault detection and isolation for X-37’s electrical power system and electro-mechanical actuators. The experiment used Livingstone, a software system that performs diagnosis using a qualitative, model-based reasoning approach that searches system-wide interactions to detect and isolate failures. Two of the challenges we faced were to make this research software more efficient so that it would fit within the limited computational resources that were available to us on the X-37 spacecraft, and to modify it so that it satisfied the X-37’s software safety requirements. Although the experiment is currently unfunded, the development effort had value in that it resulted in major improvements in Livingstone’s efficiency and safety. This paper reviews some of the details of the modeling and integration efforts, and some of the lessons that were learned. Author Systems Health Monitoring; Management Systems; Technology Assessment; X-37 Vehicle; Applications Programs (Computers)

20020063408 NASA Glenn Research Center, Cleveland, OH USA Status of the RBCC Direct-Connect Mixer Combustor Experiment Walker, James F.; Kamhawi, Hani; Krivanek, Thomas M.; Thomas, Scott R.; Smith, Timothy D.; May 2002; In English; Combustion, Airbreathing Propulsion, Propulsion Systems Hazards, and Modelling and Simulation Subcommittes Joint Meeting, 8-12 Apr. 2002, Destin, FL, USA; Original contains color illustrations Contract(s)/Grant(s): RTOP 708-90-63 Report No.(s): NASA/TM-2002-211555; NAS 1.15:211555; E-13334; No Copyright; Avail: CASI; A03, Hardcopy The NASA Glenn Research Center is developing hydrogen based combined cycle propulsion technology for a single-stage-to-orbit launch vehicle application under a project called GTX. Rocket Based Combined Cycle (RBCC) propulsion systems incorporate one or more rocket engines into an airbreathing flow path to increase specific impulse as compared to an all rocket-powered vehicle. In support of this effort, an RBCC direct-connect test capability was established at the Engine Components Research Laboratory to investigate low speed, ejector ramjet, and initial ramjet operations and performance. The facility and test article enables the evaluation of two candidate low speed operating schemes; the simultaneous mixing and combustion (SMC) and independent ramjet stream (IRS). The SMC operating scheme is based on the fuel rich operations of the rocket where performance depends upon mixing between the rocket plume and airstream. In contrast, the IRS scheme fuels the airstream separately and uses the rocket plume to ignite the fuel-air mixture. This paper describes the test hardware and facility upgrades installed to support the RBCC tests. It also defines and discusses low speed technical challenges being addressed by the experiments. Finally, preliminary test results, including rocket risk mitigating tests, unfueled airflow tests, and the integrated system hot fire test will be presented. Author Rocket-Based Combined-Cycle Engines; Propulsion System Performance; Propulsion System Configurations; Engine Tests; Pressure Distribution; Combustion Chambers

20020060131 NASA Langley Research Center, Hampton, VA USA Opportunities for Breakthroughs in Large-Scale Computational Simulation and Design Alexandrov, Natalia; Alter, Stephen J.; Atkins, Harold L.; Bey, Kim S.; Bibb, Karen L.; Biedron, Robert T.; Carpenter, Mark H.; Cheatwood, F. McNeil; Drummond, Philip J.; Gnoffo, Peter A., et al.; June 2002; In English Contract(s)/Grant(s): RTOP 706-31-21-80 Report No.(s): NASA/TM-2002-211747; L-18196; NAS 1.15:211747; No Copyright; Avail: CASI; A03, Hardcopy Opportunities for breakthroughs in the large-scale computational simulation and design of aerospace vehicles are presented. Computational ﬂuid dynamics tools to be used within multidisciplinary analysis and design methods are emphasized. The opportunities stem from speedups and robustness improvements in the underlying unit operations associated with simulation (geometry modeling, grid generation, physical modeling, analysis, etc.). Further, an improved programming environment can synergistically integrate these unit operations to leverage the gains. The speedups result from reducing the problem setup time through geometry modeling and grid generation operations, and reducing the solution time through the operation counts associated with solving the discretized equations to a sufficient accuracy. The opportunities are addressed only at a general level here, but an extensive list of references containing further details is included. The opportunities discussed are being addressed through the Fast Adaptive Aerospace Tools (FAAST) element of the Advanced Systems Concept

54 to Test (ASCoT) and the third Generation Reusable Launch Vehicles (RLV) projects at NASA Langley Research Center. The overall goal is to enable greater inroads into the design process with large-scale simulations. Author Aerospace Vehicles; Computational Fluid Dynamics; Simulation; Computer Aided Design

20020053646 NASA Ames Research Center, Moffett Field, CA USA Time Dependent Simulation of Turbopump Flows Kwak, Dochan; Kiris, Cetin C.; Chan, William; Williams, Robert; [2001]; In English; Thermal and Fluids Analysis Workshop (TFAWS-2001), 10-14 Sep. 2001, Huntsville, AL, USA; No Copyright; Avail: CASI; A03, Hardcopy Viewgraphs detailing time-dependent simulations of Turbopump ﬂows are presented. The topics include: 1) Major Drivers of Current Work; 2) Solution Methods; and 3) Unsteady Turbopump Flow. CASI Time Dependence; Computerized Simulation; Unsteady Flow; Turbine Pumps

20020051506 Hampton Univ., VA USA Statistical Methods for Rapid Aerothermal Analysis and Design Technology Morgan, Carolyn; DePriest, Douglas; Thompson, Richard, Technical Monitor; Jan. 31, 2002; In English Contract(s)/Grant(s): NAG1-01075; HU Proj. 5-23845; No Copyright; Avail: CASI; A03, Hardcopy The cost and safety goals for NASA’s next generation of reusable launch vehicle (RLV) will require that rapid high-ﬁdelity aerothermodynamic design tools be used early in the design cycle. To meet these requirements, it is desirable to establish statistical models that quantify and improve the accuracy, extend the applicability, and enable combined analyses using existing prediction tools. The research work was focused on establishing the suitable mathematical/statistical models for these purposes. It is anticipated that the resulting models can be incorporated into a software tool to provide rapid, variable-ﬁdelity, aerothermal environments to predict heating along an arbitrary trajectory. This work will support development of an integrated design tool to perform automated thermal protection system (TPS) sizing and material selection. Author Reusable Launch Vehicles; Thermal Analysis; Statistical Analysis; Costs; Safety; Aerothermodynamics

20020050420 NASA Marshall Space Flight Center, Huntsville, AL USA Water Flow Performance of a Superscale Model of the Fastrac Liquid Oxygen Pump Skelley, Stephen; Zoladz, Thomas; The Tenth Thermal and Fluids Analysis Workshop; July 2001; In English; No Copyright; Avail: CASI; A02, Hardcopy As part of the National Aeronautics and Space Administration’s ongoing effort to lower the cost of access to space, the Marshall Space Flight Center has developed a rocket engine with 60,000 pounds of thrust for use on the Reusable Launch Vehicle technology demonstrator slated for launch in 2000. This gas generator cycle engine, known as the Fastrac engine, uses liquid oxygen and RP-1 for propellants and includes single stage liquid oxygen and RP-1 pumps and a single stage supersonic turbine on a common shaft. The turbopump design effort included the first use and application of new suction capability prediction codes and three-dimensional blade generation codes in an attempt to reduce the turbomachinery design and certification costs typically associated with rocket engine development. To verify the pump’s predicted cavitation performance, a water flow test of a superscale model of the Fastrac liquid oxygen pump was conducted to experimentally evaluate the liquid oxygen pump’s performance at and around the design point. The water flow test article replicated the flow path of the Fastrac liquid oxygen pump in a 1.582x scale model, including scaled seal clearances for correct leakage flow at a model operating speed of 5000 revolutions per minute. Flow entered the 3-blade axial-flow inducer, transitioned to a shrouded, 6- blade radial impeller, and discharged into a vaneless radial diffuser and collection volute. The test article included approximately 50 total and static pressure measurement locations as well as flush-mounted, high frequency pressure transducers for complete mapping of the pressure environment. The primary objectives of the water flow test were to measure the steady-state and dynamic pressure environment of the liquid oxygen pump versus flow coefficient, suction specific speed, and back face leakage flow rate. Initial results showed acceptable correlation between the predicted and experimentally measured pump head rise at low suction specific speeds. Likewise, only small circumferential variations in steady-state were observed from 80% to 120% of the design flow coefficient, matching the computational predictions and confirming that the integrated design approach has minimized any exit volute-induced distortions. The test article exhibited suction performance trends typically observed in inducer designs with virtually constant head rise with decreasing inlet pressure until complete pump head breakdown. Unfortunately, the net positive suction head at 3% head fall-off occurred far below that predicted at

55 all tested flow coefficients, resulting in a negative net positive suction head margin at the design point in water. Additional testing to map the unsteady pressure environment was conducted and cavitation-induced flow disturbances at the inducer inlet were observed. Two distinct disturbances were identified, one rotating and one stationary relative to the fixed frame of reference, while the transition from one regime to the next produced significant effects on the steady state pump performance. The impact of the unsteady phenomena and the corresponding energy losses on the unexpectedly poor pump performance is also discussed. Author Water Flow; Scale Models; Liquid Oxygen; Pumps; Rocket Engines; Fabrication

20020050402 NASA Marshall Space Flight Center, Huntsville, AL USA Overview of Fluid Dynamic Activities at the Marshall Space Flight Center Garcia, Roberto; Griffin, Lisa; Wang, Ten-See; The Tenth Thermal and Fluids Analysis Workshop; July 2001; In English; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy Contents include following: (1) Organizational Changes at MSFC. (2) Recent Program Support & Technology Development: analysis & cold ﬂow testing: Fastrac, X-34, X-33, RLV, LFBB. (3) Ongoing Activities: RLV focused technology, RBCC concepts development, methodology & code development. (4) Future Activities and Direction: hardware design and development; tools development. (5) Concluding remarks: constraints, cooperation, opportunities. CASI Support Systems; Research Facilities; Research Management

20020050399 NASA Marshall Space Flight Center, Huntsville, AL USA A Collaborative Analysis Tool for Thermal Protection Systems for Single Stage to Orbit Launch Vehicles Alexander, Reginald; Stanley, Thomas Troy; The Tenth Thermal and Fluids Analysis Workshop; July 2001; In English; Original contains color illustrations; No Copyright; Avail: CASI; A02, Hardcopy Presented is a design tool and process that connects several disciplines which are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to all other systems, as is the case with SSTO vehicles with air breathing propulsion, which is currently being studied by the National Aeronautics and Space Administration (NASA). In particular, the thermal protection system (TPS) is linked directly to almost every major system. The propulsion system pushes the vehicle to velocities on the order of 15 times the speed of sound in the atmosphere before pulling up to go to orbit which results in high temperatures on the external surfaces of the vehicle. Thermal protection systems to maintain the structural integrity of the vehicle must be able to mitigate the heat transfer to the structure and be lightweight. Herein lies the interdependency, in that as the vehicle’s speed increases, the TPS requirements are increased. And as TPS masses increase the effect on the propulsion system and all other systems is compounded. To adequately calculate the TPS mass of this type of vehicle several engineering disciplines and analytical tools must be used preferably in an environment that data is easily transferred and multiple iterations are easily facilitated. Author Thermal Protection; Structural Failure; Single Stage to Orbit Vehicles

20020048698 NASA Marshall Space Flight Center, Huntsville, AL USA, Societe Nationale d’Etude et de Construction de Moteurs d’Aviation, Bordeaux France Demonstration of Advanced C/SiC Cooled Ramp Bouquet, Clement; Laithier, Frederic; Lawrence, Timothy; Eckel, Andrew; Munafo, Paul M., Technical Monitor; [2002]; In English; 6th International Symposium on Propulsion for Space Transportation of the 21st Century, 14-16 May 2002, Versailles, France; Original contains color illustrations Contract(s)/Grant(s): NAS8-99112; No Copyright; Avail: CASI; A02, Hardcopy Under a NASA contract, SPS is evaluating its C/SiC to metal brazing technique for the development of light, composite, actively cooled panels. The program ﬁrst consisted of deﬁning a system applicable to the X-33 nozzle ramp. SPS then performed evaluation tests for tube, composite, and braze material selection, and for the adaptation of braze process parameters to the parts geometry. SPS is presently manufacturing a 250x60 millimeter squared specimen, including 10 metallic tubes, which will be cycled in the NASA/GRC-CELL-22 test bed under engine representative conditions. Author Silicon Carbides; Carbon; Composite Materials; Manufacturing; Cooling

56 20020048696 ATK-Thiokol Propulsion, USA Automation of NDE on RSRM Metal Components Hartman, John; Kirby, Mark; McCool, Alex, Technical Monitor; [2002]; In English; ASNT Spring Conference, 18-22 Mar. 2002, Portland, OR, USA; Original contains color illustrations Contract(s)/Grant(s): NAS8-97238; No Copyright; Avail: CASI; A03, Hardcopy An automated eddy current system has been designed and built, and is being implemented to inspect RSRM (Space Shuttle) metal components. The system provides a signiﬁcant increase in inspection reliability, as well as other beneﬁts such as data storage, chemical waste reduction and reduction in overall process time. This paper is in viewgraph form. Derived from text Nondestructive Tests; Reusable Launch Vehicles; Automation; Metals; Spacecraft Components

20020048273 NASA Ames Research Center, Moffett Field, CA USA Fractographic Analysis of HfB2-SiC and ZrB2-SiC Composites Mecholsky, J.J., Jr.; Ellerby, D. T.; Johnson, S. M.; Stackpoole, M. M.; Loehman, R. E.; Arnold, Jim, Technical Monitor; [2001]; In English; 26th Annual International Conference on Advanced Ceramics and Composites, 13-18 Jan. 2002; Copyright; Avail: Other Sources; Abstract Only Hafnium diboride-silicon carbide and zirconium diboride-silicon carbide composites are potential materials for high temperature leading edge applications on reusable launch vehicles. In order to establish material constants necessary for evaluation of in-situ fracture, bars fractured in four point ﬂexure were examined using fractographic principles. The fracture toughness was determined from measurements of the critical crack sizes and the strength values, and the crack branching constants were established to use in forensic fractography of materials for future ﬂight applications. The fracture toughnesses range from about 13 MPam (sup 1/2) at room temperature to about 6 MPam (sup 1/2) at 1400 C for ZrB2-SiC composites and from about 11 MPam (sup 1/2) at room temperature to about 4 MPam (sup 1/2) at 1400 C for HfB2-SiC composites. Author Fracture Strength; Leading Edges; Reusable Launch Vehicles; Refractory Materials

20020047561 NASA Marshall Space Flight Center, Huntsville, AL USA Permeability After Impact Testing of Composite Laminates Nettles, A.T.; Munafo, Paul, Technical Monitor; 2002; In English; ASTM Symposia on Composites, 10-13 Mar. 2002, Pittsburgh, PA, USA; No Copyright; Avail: CASI; A03, Hardcopy Since composite laminates are beginning to be identified for use in reusable launch vehicle propulsion systems, an understanding of their permeance is needed. A foreign object impact event can cause a localized area of permeability (leakage) in a polymer matrix composite and it is the aim of this study to assess a method of quantifying permeability-after-impact results. A simple test apparatus is presented and variables that could affect the measured values of permeability-after-impact were assessed. Once it was determined that valid numbers were being measured, a fiber/resin system was impacted at various impact levels and the resulting permeability measured, first with a leak check solution (qualitative) then using the new apparatus (quantitative). The results showed that as the impact level increased, so did the measured leakage. As the pressure to the specimen was increased, the leak rate was seen to increase in a non-linear fashion for almost all of the specimens tested. Author Impact Tests; Laminates; Permeability; Polymer Matrix Composites; Leakage

20020045392 NASA Marshall Space Flight Center, Huntsville, AL USA Natural Atmospheric Environment Model Development for the National Aeronautics and Space Administration’s Second Generation Reusable Launch Vehicle Roberts, Barry C.; Leahy, Frank; Overbey, Glenn; Batts, Glen W.; Parker, Nelson, Technical Monitor; [2002]; In English; 9th Conference on Aviation, Range and Aerospace Meteorology, 13-16 May 2002, Portland, OR, USA; No Copyright; Avail: CASI; A01, Hardcopy The National Aeronautics and Space Administration (NASA) recently began development of a new reusable launch vehicle. The program office is located at Marshall Space Flight Center (MSFC) and is called the Second Generation Reusable Launch Vehicle (2GRLV). The purpose of the program is to improve upon the safety and reliability of the ﬁrst generation reusable launch vehicle, the Space Shuttle. Speciﬁcally, the goals are to reduce the risk of crew loss to less than 1-in-10,000 missions and decreased costs by a factor of 10 to approximately $1,000 per pound of payload launched to low Earth orbit. The program is currently in the very early stages of development and many two-stage vehicle concepts will be evaluated. Risk

57 reduction activities are also taking place. These activities include developing new technologies and advancing current technologies to be used by the vehicle. The Environments Group at MSFC is tasked by the 2GRLV Program to develop and maintain an extensive series of analytical tools and environmental databases which enable it to provide detailed atmospheric studies in support of structural, guidance, navigation and control, and operation of the 2GRLV. Author Atmospheric Models; Reusable Launch Vehicles; Technology Utilization

20020045346 Research Inst. for Advanced Computer Science, Moffett Field, CA USA New V and V Tools for Diagnostic Modeling Environment (DME) Pecheur, Charles; Nelson, Stacy; Merriam, Marshall, Technical Monitor; April 2002; In English Contract(s)/Grant(s): NRA-8-30/TA-5-Northrup-Grumman Report No.(s): NASA/CR-2002-211403; NAS 1.26:211403; No Copyright; Avail: CASI; A03, Hardcopy The purpose of this report is to provide correctness and reliability criteria for veriﬁcation and validation (V&V) of Second Generation Reusable Launch Vehicle (RLV) Diagnostic Modeling Environment, describe current NASA Ames Research Center tools for V&V of Model Based Reasoning systems, and discuss the applicability of Advanced V&V to DME. This report is divided into the following three sections: (1) correctness and reliability criteria; (2) tools for V&V of Model Based Reasoning; and (3) advanced V&V applicable to DME. The Executive Summary includes an overview of the main points from each section. Supporting details, diagrams, ﬁgures, and other information are included in subsequent sections. A glossary, acronym list, appendices, and references are included at the end of this report. Author Reliability Analysis; Reusable Launch Vehicles; Software Development Tools

20020045344 Research Inst. for Advanced Computer Science, Moffett Field, CA USA V and V of Advanced Systems at NASA Pecheur, Charles; Nelson, Stacy; April 2002; In English Contract(s)/Grant(s): NRA-8-30/TA-5/Northrop-Grumman Report No.(s): NASA/CR-2002-211401; NAS 1.26:211401; No Copyright; Avail: CASI; A05, Hardcopy The purpose of this report is to provide the following: 1) Overview of Formal Methods beneficial to Verification and Validation (V&V) of 2nd Generation Re-usable Launch Vehicle Integrated Vehicle Health Management (RLV IVHM); 2) Description of current use of Formal Methods at NASA and their applicability to 2nd Generation RLV IVHM software; and 3) Guide for incorporation Formal Methods into NASA V&V Standards for certification of airborne software like 2nd Generation RLV IVHM. Author NASA Programs; Reusable Launch Vehicles; Program Verification (Computers); Software Engineering

20020044823 NASA Marshall Space Flight Center, Huntsville, AL USA X-33 LH2 Tank Failure Investigation Findings Niedermeyer, Melinda; Munafo, Paul, Technical Monitor; [2002]; In English; Manufacturing Problem Prevention Workshop, 26-27 Feb. 2002, El Segundo, CA, USA; No Copyright; Avail: CASI; A03, Hardcopy This viewgraph presentation gives an overview of the X-33 LH2 tank failure investigation ﬁndings. The conclusions of the investigation include the following: (1) the inner skin microcracked and hydrogen inﬁltrated; (2) the cracks grew larger under pressure; (3) when pressure was removed, the cracks closed slightly; (4) when the tank was drained and warmed, the cracks closed and blocked the leak path; (5) FOD and debond areas provided an opportunity for a leak path; and (6) there is still hydrogen in the the other three lobes today. Derived from text Failure Analysis; Leakage; Microcracks; X-33 Reusable Launch Vehicle

20020044103 Research Inst. for Advanced Computer Science, Moffett Field, CA USA, Computer Sciences Corp., Moffett Field, CA USA Survey of NASA V and V Processes/Methods Pecheur, Charles; Nelson, Stacy; April 2002; In English Contract(s)/Grant(s): NRA-8-30/TA-5/Northrop-Grumman Report No.(s): NASA/CR-2002-211401; NAS 1.26:211401; No Copyright; Avail: CASI; A05, Hardcopy

58 The purpose of this report is to describe current NASA Veriﬁcation and Validation (V&V) techniques and to explain how these techniques are applicable to 2nd Generation RLV Integrated Vehicle Health Management (IVHM) software. It also contains recommendations for special V&V requirements for IVHM. This report is divided into the following three sections: 1) Survey - Current NASA V&V Processes/Methods; 2) Applicability of NASA V&V to 2nd Generation RLV IVHM; and 3) Special 2nd Generation RLV IVHM V&V Requirements. Author Systems Integration; Program Veriﬁcation (Computers); Software Engineering; NASA Programs; Surveys

20020043393 NASA Ames Research Center, Moffett Field, CA USA Thermal Protection Materials for Reentry Applications Johnson, Sylvia M.; Stackpoole, Mairead; Gusman, Mike; Loehman, Ron; Kotula, Paul; Ellerby, Donald; Arnold, James; Wercinski, Paul; Reuthers, James; Kontinos, Dean, et al.; [2001]; In English; Structural Ceramics and Ceramic Composites for High-Temperature Applications Conference, 11 Oct. 2001, Seville, Spain; No Copyright; Avail: CASI; A03, Hardcopy Thermal protection materials and systems (IRS) are used to protect spacecraft during reentry into Earth’s atmosphere or entry into planetary atmospheres. As such, these materials are subject to severe environments with high heat fluxes and rapid heating. Catalytic effects can increase the temperatures substantially. These materials are also subject to impact damage from micrometeorites or other debris during ascent, orbit, and descent, and thus must be able to withstand damage and to function following damage. Thermal protection materials and coatings used in reusable launch vehicles will be reviewed, including the needs and directions for new materials to enable new missions that require faster turnaround and much greater reusability. The role of ablative materials for use in high heat flux environments, especially for non-reusable applications and upcoming planetary missions, will be discussed. New thermal protection system materials may enable the use of sharp nose caps and leading edges on future reusable space transportation vehicles. Vehicles employing this new technology would have significant increases in maneuverability and out-of-orbit cross range compared to current vehicles, leading to increased mission safety in the event of the need to abort during ascent or from orbit. Ultrahigh temperature ceramics, a family of materials based on HfB2 and ZrB2 with SiC, will be discussed. The development, mechanical and thermal properties, and uses of these materials will be reviewed. Author Thermal Protection; Spacecraft Reentry; Reusable Launch Vehicles; Planetary Atmospheres; Ablative Materials

20020043183 NASA Ames Research Center, Moffett Field, CA USA NASA Exhibits Deardorff, Glenn; Djomehri, M. Jahed; Freeman, Ken; Gambrel, Dave; Green, Bryan; Henze, Chris; Hinke, Thomas; Hood, Robert; Kiris, Cetin; Moran, Patrick; Biegel, Bryan, Technical Monitor, et al.; [2001]; In English; Supercomputing 2001, 10-16 Nov. 2001, Denver, CO, USA Contract(s)/Grant(s): RTOP 704-40-12; RTOP 725-10-11; RTOP 725-10-31; RTOP 727-01-83; No Copyright; Avail: CASI; A03, Hardcopy A series of NASA presentations for the Supercomputing 2001 conference are summarized. The topics include: (1) Mars Surveyor Landing Sites ‘Collaboratory’; (2) Parallel and Distributed CFD for Unsteady Flows with Moving Overset Grids; (3) IP Multicast for Seamless Support of Remote Science; (4) Consolidated Supercomputing Management Office; (5) Growler: A Component-Based Framework for Distributed/Collaborative Scientiﬁc Visualization and Computational Steering; (6) Data Mining on the Information Power Grid (IPG); (7) Debugging on the IPG; (8) Debakey Heart Assist Device: (9) Unsteady Turbopump for Reusable Launch Vehicle; (10) Exploratory Computing Environments Component Framework; (11) OVERSET Computational Fluid Dynamics Tools; (12) Control and Observation in Distributed Environments; (13) Multi-Level Parallelism Scaling on NASA’s Origin 1024 CPU System; (14) Computing, Information, & Communications Technology; (15) NAS Grid Benchmarks; (16) IPG: A Large-Scale Distributed Computing and Data Management System; and (17) ILab: Parameter Study Creation and Submission on the IPG. Derived from text Supercomputers; Parallel Processing (Computers); Multiprocessing (Computers); Distributed Processing; Conferences; Computation

20020041448 Amain Electronics Co., Inc., Reno, NV USA Mosad and Stream Vision For A Telerobotic, Flying Camera System Mandl, William; Apr. 03, 2002; In English; Diskette: 1 3.5-Inch DSHD diskette containing full text document in PDF format Contract(s)/Grant(s): NAS3-00031 Report No.(s): Amain-Rept-02002; NONP-NASA-DK-2002067346; No Copyright; Avail: CASI; A03, Hardcopy

59 Two full custom camera systems using the Multiplexed OverSample Analog to Digital (MOSAD) conversion technology for visible light sensing were built and demonstrated. They include a photo gate sensor and a photo diode sensor. The system includes the camera assembly, driver interface assembly, a frame stabler board with integrated decimeter and Windows 2000 compatible software for real time image display. An array size of 320X240 with 16 micron pixel pitch was developed for compatibility with 0.3 inch CCTV optics. With 1.2 micron technology, a 73% ﬁll factor was achieved. Noise measurements indicated 9 to 11 bits operating with 13.7 bits best case. Power measured under 10 milliwatts at 400 samples per second. Nonuniformity variation was below noise ﬂoor. Pictures were taken with different cameras during the characterization study to demonstrate the operable range. The successful conclusion of this program demonstrates the utility of the MOSAD for NASA missions, providing superior performance over CMOS and lower cost and power consumption over CCD. The MOSAD approach also provides a path to radiation hardening for space based applications. Author Cameras; Telerobotics; Analog to Digital Converters; Fabrication; Sensors; Light (Visible Radiation); Charge Coupled Devices

20020036009 Lee and Associates, LLC, USA Support to 2nd Generation RLV Propulsion Project Office Lee, Thomas J.; Mar. 07, 2002; In English Contract(s)/Grant(s): NASA Order H-34351-D; No Copyright; Avail: CASI; A03, Hardcopy In this ﬁnal report regarding support to the second generation RLV (Reusable Launch Vehicle) propulsion project office, a list of tasks accomplished is presented. During this period, Lee & Associates, LLC participated in numerous Systems Requirements Reviews (SRR) related to the Cobra development program. CASI Propulsion; Reusable Launch Vehicles

20020034965 NASA Marshall Space Flight Center, Huntsville, AL USA X-33 Tank Failure During Autoclave Fabrication Nettles, Alan T.; Munafo, Paul, Technical Monitor; [2001]; In English; US Air Force Manufacturing Problem Prevention Program, 5-6 Jun. 2001, El Segundo, CA, USA; No Copyright; Avail: Other Sources; Abstract Only During a repair cure cycle on tank #1 of the X-33 liquid hydrogen tanks, a skin to core disbond occurred. Both the inner skin and outer skin of the lobe #1 sandwich panel was noted to have been disbonded and cracked- An investigation was undertaken to determine the cause of this failure. The investigation consisted of reviewing all of the processing data and performing testing on the failed lobe #1, as well as the other lobes, which did not fail during the cure cycle. The tests consisted of residual stress measurements in one of the intact lobes and ‘plug-pulls’ to assess skin to core strength on all of the remaining lobes. Results showed an extremely low bondline strength due to lack of proper ﬁlleting of the adhesive, in addition, tests showed a very rapid decrease in strength with increasing temperature, as well as a further decrease in strength with a larger number of cycles. Also, the honeycomb used was not vented so pressure could build up within the cells. All of these factors appeared to be contributors to the failure. Author X-33 Reusable Launch Vehicle; Propellant Tanks; Liquid Hydrogen; Failure

20020034460 NASA Marshall Space Flight Center, Huntsville, AL USA Findings from the X-33 Hydrogen Tank Failure Investigation Niedermeyer, Melinda; Munafo, Paul M., Technical Monitor; [2001]; In English; US Air Force Manufacturing Problem Prevention Program, 5-6 Jun. 2001, El Segundo, CA, USA; No Copyright; Avail: Other Sources; Abstract Only The X-33 Hydrogen tank failed during test in November of 1999 at MSFC. The tank completed the structural loading phase of the test successfully and was drained of hydrogen prior to the failure. The failure initiated in the acreage of Lobe 1 and was instantaneous, peeling the outer skin and core away from the inner skin. It was determined there were several factors that provided the opportunity for the tank to fail in this way. The factor giving life to these opportunistic circumstances was hydrogen inﬁltration into the core of the tank. The mechanism for this phenomenon will be discussed in this presentation. Author X-33 Reusable Launch Vehicle; Propellant Tanks; Liquid Hydrogen; Failure

60 20020032739 Georgia Inst. of Tech., Atlanta, GA USA Improvements in Thermal Protection Sizing Capabilities for TCAT: Conceptual Design for Advanced Space Transportation Systems Olds, John R.; Izon, Stephen James; February 2002; In English Contract(s)/Grant(s): NAG8-1597 Report No.(s): E-16-P51; No Copyright; Avail: CASI; A05, Hardcopy The Thermal Calculation Analysis Tool (TCAT), originally developed for the Space Systems Design Lab at the Georgia Institute of Technology, is a conceptual design tool capable of integrating aeroheating analysis into conceptual reusable launch vehicle design. It provides Thermal Protection System (TPS) unit thicknesses and acreage percentages based on the geometry of the vehicle and a reference trajectory to be used in calculation of the total cost and weight of the vehicle design. TCAT has proven to be reasonably accurate at calculating the TPS unit weights for in-ﬂight trajectories; however, it does not have the capability of sizing TPS materials above cryogenic fuel tanks for ground hold operations. During ground hold operations, the vehicle is held for a brief period (generally about two hours) during which heat transfer from the TPS materials to the cryogenic fuel occurs. If too much heat is extracted from the TPS material, the surface temperature may fall below the freezing point of water, thereby freezing any condensation that may be present at the surface of the TPS. Condensation or ice on the surface of the vehicle is potentially hazardous to the mission and can also damage the TPS. It is questionable whether or not the TPS thicknesses provided by the aeroheating analysis would be sufficiently thick to insulate the surface of the TPS from the heat transfer to the fuel. Therefore, a design tool has been developed that is capable of sizing TPS materials at these cryogenic fuel tank locations to augment TCAT’s TPS sizing capabilities. Derived from text Aerodynamic Heat Transfer; Surface Temperature; Systems Analysis; Systems Engineering; Thermal Protection; Computerized Simulation

20020024753 NASA Marshall Space Flight Center, Huntsville, AL USA An Advanced Video Sensor for Automated Docking Howard, Richard T.; Bryan, Thomas C.; Book, Michael L.; Roe, Fred, Technical Monitor; [2001]; In English; 20th Digital Avionics Systems Conference, 14-18 Oct. 2001, Daytona Beach, FL, USA; No Copyright; Avail: Other Sources; Abstract Only This paper describes the current developments in video-based sensors at the Marshall Space Flight Center. The Advanced Video Guidance Sensor is the latest in a line of video-based sensors designed for use in automated docking systems. The X-33, X-34, X-38, and X-40 are all designed to be unpiloted vehicles; such vehicles will require a sensor system that will provide adequate data for the vehicle to accomplish its mission. One of the primary tasks planned for re-usable launch vehicles is to resupply the space station. In order to approach the space station in a self-guided manner, the vehicle must have a reliable and accurate sensor system to provide relative position and attitude information between the vehicle and the space station. The Advanced Video Guidance Sensor is being designed and built to meet this requirement, as well as requirements for other vehicles docking to a variety of target spacecraft. The Advanced Video Guidance Sensor is being designed to allow range and bearing information to be measured at ranges up to 2 km. The sensor will measure 6-degree-of-freedom information (relative positions and attitudes) from approximately 40 meters all the way in to ﬁnal contact (approximately 1 meter range). The sensor will have a data output rate of 20 Hz during tracking mode, and will be able to acquire a target within one half of a second. The prototype of the sensor will be near completion at the time of the conference. Author Automatic Control; Spacecraft Docking; Reusable Launch Vehicles; Guidance Sensors

20020024007 NASA Glenn Research Center, Cleveland, OH USA High Voltage Design Guidelines: A Timely Update Hillard, G. Barry; Kirkici, H.; Ensworth, Clint, Technical Monitor; [2001]; In English; 2001 Conference on Electrical Insulation and Dielectric Phenomena, 14-17 Oct. 2001, Kitchener, Ontario, Canada Contract(s)/Grant(s): RTOP 376-80-21; No Copyright; Avail: CASI; A01, Hardcopy The evolving state of high voltage systems and their increasing use in the space program have called for a revision of the High Voltage Design Guidelines, Marshall Space Flight Center technical document MSFC-STD-531, originally issued September 1978 (previously 50 M05189b, October 1972). These guidelines deal in depth with issues relating to the specification of materials, particularly electrical insulation, as well as design practices and test methods. Emphasis is on corona and Paschen breakdown as well as plasma effects for Low Earth Orbiting systems. We will briefly review the history of these guidelines as well as their immediate predecessors and discuss their range of applicability. In addition, this document has served as the basis for several derived works that became focused, program-specific HV guidelines. We will briefly review

61 two examples, guidelines prepared for the X-33 program and for the Space Shuttle Electric Auxiliary Power Unit (EAPU) upgrade. Author High Voltages; Space Shuttles; Auxiliary Power Sources; Design Analysis; Space Programs

20020023442 Boeing Phantom Works, USA Dynamics and Stability and Control Characteristics of the X-37 Chaudhary, Ashwani; Nguyen, Viet; Tran, Hoi; Poladian, David; Falangas, Eric; Turner, Susan G., Technical Monitor; [2001]; In English; AIAA GN&C Conference, 9 Aug. 2001, Montreal, Canada Contract(s)/Grant(s): NCC8-190; Copyright; Avail: CASI; A02, Hardcopy This paper presents the stability and control analysis and the control design results for the Boeing/NASA/AFRL X-37. The X-37 is a ﬂight demonstrator vehicle that will go into space and after its mission, autonomously reenter and land on a conventional runway. This paper studies the dynamics and control of the X-37 from atmospheric reentry through landing. A nominal trajectory that lands on the Edwards Air Force Base Lakebed is considered for all the analysis and design. The X-37’s longitudinal and lateral/directional bare-airframe characteristics are presented. The level of maneuvering control power is assessed. Vehicle trim with multiple surfaces is discussed. Special challenges where the wings loose roll effectiveness are discussed and solutions are presented. Aerodynamic uncertainties and ﬂexibility modeling issues are presented. Control design results and robustness analysis methods are presented. Results are provided for the Entry, Terminal Area Energy Management (TAEM), and Approach and Land phases. Author Stability; Terminal Area Energy Management; X-37 Vehicle; Control Systems Design; Dynamic Characteristics

20020022529 NASA Glenn Research Center, Cleveland, OH USA 2000 NASA Seal/Secondary Air System Workshop, Volume 1 Steinetz, Bruce M., Editor; Hendricks, Robert C., Editor; October 2001; In English, 25-26 Oct. 2000, Cleveland, OH, USA Contract(s)/Grant(s): RTOP 706-85-37 Report No.(s): NASA/CP-2001-211208/VOL1; E-13062/VOL1; NAS 1.55:211208/VOL1; No Copyright; Avail: CASI; A21, Hardcopy The 2000 NASA Seal/Secondary Air System Workshop covered four main areas: (1) overviews of NASA-sponsored Ultra-Efficient Engine Technology (UEET) and Access to Space Programs, with emphasis on program goals and seal needs; (2) review of turbine engine seal issues from the perspective of end users such as United Airlines; (3) reviews of sealing concepts, test results, experimental facilities, and numerical predictions; and (4) reviews of material development programs relevant to advanced seals development. The NASA UEET overview illustrates for the reader the importance of advanced technologies, including seals, in meeting future engine system efficiency and emission goals. GE, Pratt & Whitney, and Honeywell presented advanced seal development work being performed within their organizations. The NASA-funded GE/Stein Seal team has successfully demonstrated a large (3-ft. diam) aspirating seal that can withstand all anticipated pressures, speeds, and rotor runouts anticipated for a GE90 L.P. turbine balance piston location. GE/Stein Seal are fabricating a full-scale seal to be tested in a GE-90 ground test engine in early 2002. Pratt & Whitney and Stein Seal are investigating carbon seals to accommodate large radial movements anticipated in future geared-fan gearbox locations. Honeywell presented a ﬁnger seal design being considered for a high-temperature static combustor location incorporating ceramic ﬁnger elements. Successful demonstration of the braided carbon rope thermal barriers to extreme temperatures (5500 F) for short durations provide a new form of very high temperature thermal barrier for future Shuttle solid rocket motor nozzle joints. The X-37, X-38, and future highly reusable launch vehicles pose challenging control surface seal demands that require new seal concepts made from emerging high temperature ceramics and other materials. Author Aerospace Engineering; Seals (Stoppers); Conferences; Product Development

20020022190 Georgia Inst. of Tech., Atlanta, GA USA Evaluation of Advanced Thermal Protection Techniques for Future Reusable Launch Vehicles Olds, John R.; Cowart, Kris; June 2001; In English Contract(s)/Grant(s): NCC2-5332; No Copyright; Avail: CASI; A09, Hardcopy A method for integrating Aeroheating analysis into conceptual reusable launch vehicle RLV design is presented in this thesis. This process allows for faster turn-around time to converge a RLV design through the advent of designing an optimized

62 thermal protection system (TPS). It consists of the coupling and automation of four computer software packages: MINIVER, TPSX, TCAT and ADS. MINIVER is an Aeroheating code that produces centerline radiation equilibrium temperatures, convective heating rates, and heat loads over simplified vehicle geometries. These include flat plates and swept cylinders that model wings and leading edges, respectively. TPSX is a NASA Ames material properties database that is available on the World Wide Web. The newly developed Thermal Calculation Analysis Tool (TCAT) uses finite difference methods to carry out a transient in-depth I-D conduction analysis over the center mold line of the vehicle. This is used along with the Automated Design Synthesis (ADS) code to correctly size the vehicle’s thermal protection system JPS). The numerical optimizer ADS uses algorithms that solve constrained and unconstrained design problems. The resulting outputs for this process are TPS material types, unit thicknesses, and acreage percentages. TCAT was developed for several purposes. First, it provides a means to calculate the transient in-depth conduction seen by the surface of the TPS material that protects a vehicle during ascent and reentry. Along with the in-depth conduction, radiation from the surface of the material is calculated along with the temperatures at the backface and interior parts of the TPS material. Secondly, TCAT contributes added speed and automation to the overall design process. Another motivation in the development of TCAT is optimization. Derived from text Aerodynamic Heating; Algorithms; Computer Programs; Convective Heat Transfer; Design Analysis; Thermal Analysis

20020022187 NASA Marshall Space Flight Center, Huntsville, AL USA Thermal/Fluid Analysis of a Composite Heat Exchanger for Use on the RLV Rocket Engine Nguyen, Dalton; Turner, Larry D., Technical Monitor; [2001]; In English; Thermal Fluid Analysis Workshop, 10-14 Sep. 2001, Huntsville, AL, USA; Original contains color illustrations; No Copyright; Avail: CASI; A02, Hardcopy As part of efforts to design a regeneratively cooled composite nozzle ramp for use on the reusable vehicle (RLV) rocket engine, a C-SiC composite heat exchanger concept was proposed for thermal performance evaluation. To test the feasibility of the concept, sample heat exchanger panels were made to fit the Glenn Research Center’s cell 22 for testing. Operation of the heat exchanger was demonstrated in a combustion environment with high heat fluxes similar to the RLV Aerospike Ramp. Test measurements were reviewed and found to be valuable for the on-going fluid and thermal analysis of the actual RLV composite ramp. Since the cooling fluid for the heat exchanger is water while the RLV Ramp cooling fluid is LH2, fluid and therma models were constructed to correlate to the specific test set-up. The knowledge gained from this work will be helpful for analyzing the thermal response of the actual RLV Composite Ramp. The coolant thermal properties for the models are taken from test data. The heat exchanger’s cooling performance was analyzed using the Generalized Fluid System Simulation Program (GFSSP). Temperatures of the heat exchanger’s structure were predicted in finite element models using Patran and Sinda. Results from the analytical models and the tests show that RSC’s heat exchanger satisfied the combustion environments in a series of 16 tests. Author Heat Exchangers; Computerized Simulation; Performance Tests; Carbon-Carbon Composites; Thermal Analysis

20020022180 Boeing Reusable Space Systems, Seal Beach, CA USA Test Results for an Off-the-Shelf GPS/INS During Approach and Landing Testing of the X-40A Childers, Dave; Gelderloos, Hendrik; Turner, Susan, Technical Monitor; [2001]; In English; 20th Digital Avionics Conference, 16-18 Oct. 2001, Daytona Beach, FL, USA Contract(s)/Grant(s): NAS8-190; No Copyright; Avail: CASI; A03, Hardcopy The X-37 is an unpiloted, reusable space vehicle that will be launched into space, orbit the Earth, reenter the atmosphere, and land autonomously. At the heart of the Guidance, Navigation and Control (GN&C) will be the Space Integrated GPS/INS (SIGI) system, an off the shelf navigation grade GPS/INS that has been enhanced for space and reentry environments. SIGI will provide both navigation and flight control data to the X-37’s GN&C. The X-40A is an unpiloted experimental vehicle whose shape and performance are similar to the X-37’s and was flown earlier this year to develop and test the approach and landing phase of X-37. On board the X-40A is the X-37’s SIGI, which is riding along as an experiment. The X-40A SIGI experiment provided early characterization of SIGI operation and performance with differential GPS (Global Positioning System) during real world approach and landing. Characterization testing was geared toward assessment of reliability and performance of the system. The objectives were to demonstrate performance levels sufficient to meet the X-37 requirements for automatic, autonomous approach and landing, demonstrate reliability over repeated ground and flight tests, and reduce risk for integration of SIGI into the vehicle and support environment. This paper presents a summary of this testing and the results to date. Author Flight Tests; Autonomous Navigation; Spacecraft Reliability; X-37 Vehicle; Performance Tests; Systems Integration

63 20020021989 NASA Marshall Space Flight Center, Huntsville, AL USA X-33 Tank Failure During Autoclave Fabrication Nettles, Alan T.; Munafo, Paul, Technical Monitor; [2001]; In English; Joint NASA/AF Meeting on Honeycomb Panels, 2 Oct. 2001, El Segundo, CA, USA; No Copyright; Avail: CASI; A03, Hardcopy The composite liquid hydrogen tank (tank #1 of 2) for the X-33 flight vehicle is made up of four lobes that have a sandwich construction, bonded to a frame of longerons. Lobes 1 and 4 showed local disbonds to the longerons they were bonded to. The ‘bad’ areas were cut away and patched with new material. The new material was cured by placing the entire tank in a heated autoclave with no pressure. Upon removal from the autoclave, it was noted that lobe 1 had severe skin/core disbonds on the inner and outer skins. The skins on this lobe were cracked as well. The core was disbonded from the inner skin across the entire acreage, except for spots around the lobe perimeter. The outer skin was separated from the core in a region near the center of the lobe. Lobe 1 was removed from the tank on January 13, 1999. Bolts were placed through the lobe to hold it together and the cuts on the inner skin were not continuous, but ‘tabs’ were left for final cutting and removal. Upon closer inspection of the disbonded basesheet, it was noted that there was a lack of filleting into the honeycomb core. Good fillets are critical to bond strength. Derived from text Bolts; Debonding (Materials); Defects; Joints (Junctions); Sandwich Structures

20020021567 NASA Marshall Space Flight Center, Huntsville, AL USA NASA’s Space Launch Initiative Targets Toxic Propellants Hurlbert, Eric; McNeal, Curtis; Davis, Daniel J., Technical Monitor; [2001]; In English; 4th International Peroxide Propulsion Conference, 20-22 Jun. 2001, Noordwijk, Netherlands Report No.(s): CP-50; No Copyright; Avail: CASI; A02, Hardcopy When manned and unmanned space flight first began, the clear and overriding design consideration was performance. Consequently, propellant combinations of all kinds were considered, tested, and, when they lifted the payload a kilometer higher, or an extra kilogram to the same altitude, they became part of our operational inventory. Cost was not considered. And with virtually all of the early work being performed by the military, safety was hardly a consideration. After all, fighting wars has always been dangerous. Those days are past now. With space flight, and the products of space flight, a regular part of our lives today, safety and cost are being reexamined. NASA’s focus turns naturally to its Shuttle Space Transportation System. Designed, built, and flown for the first time in the 1970s, this system remains today America’s workhorse for manned space flight. Without its tremendous lift capability and mission flexibility, the International Space Station would not exist. And the Hubble telescope would be a monument to shortsighted management, rather than the clear penetrating eye on the stars it is today. But the Shuttle system fully represents the design philosophy of its period: it is too costly to operate, and not safe enough for regular long term access to space. And one of the key reasons is the utilization of toxic propellants. This paper will present an overview of the utilization of toxic propellants on the current Shuttle system. Derived from text Propellants; Space Transportation System; Reusable Launch Vehicles; Cryogenics

20020020963 NASA Langley Research Center, Hampton, VA USA Advanced Metallic Thermal Protection System Development Blosser, M. L.; Chen, R. R.; Schmidt, I. H.; Dorsey, J. T.; Poteet, C. C.; Bird, R. K.; [2002]; In English; 40th Aerospace Sciences Meeting and Exhibit, 14-17 Jan. 2002, Reno, NV, USA; Original contains color illustrations Report No.(s): AIAA Paper 2002-0504; Copyright; Avail: CASI; A03, Hardcopy A new Adaptable, Robust, Metallic, Operable, Reusable (ARMOR) thermal protection system (TPS) concept has been designed, analyzed, and fabricated. In addition to the inherent tailorable robustness of metallic TPS, ARMOR TPS offers improved features based on lessons learned from previous metallic TPS development efforts. A speciﬁc location on a single-stage-to-orbit reusable launch vehicle was selected to develop loads and requirements needed to design prototype ARMOR TPS panels. The design loads include ascent and entry heating rate histories, pressures, acoustics, and accelerations. Additional TPS design issues were identiﬁed and discussed. An iterative sizing procedure was used to size the ARMOR TPS panels for thermal and structural loads as part of an integrated TPS/cryogenic tank structural wall. The TPS panels were sized to maintain acceptable temperatures on the underlying structure and to operate under the design structural loading. Detailed creep analyses were also performed on critical components of the ARMOR TPS panels. A lightweight, thermally compliant TPS support system (TPSS) was designed to connect the TPS to the cryogenic tank structure. Four 18-inch-square ARMOR TPS panels were fabricated. Details of the fabrication process are presented. Details of the TPSS for connecting the ARMOR

64 TPS panels to the externally stiffened cryogenic tank structure are also described. Test plans for the fabricated hardware are presented. Author Fabrication; Panels; Thermal Protection; Cryogenic Fluid Storage; Creep Analysis

20020020427 NASA Marshall Space Flight Center, Huntsville, AL USA NASA Alternate Access to Station Service Concept Bailey, Michelle D.; Crumbly, Chris; [2002]; In English; 53rd International Astronautical Congress Conference, 10-19 Oct. 2002, Houston, TX, USA; No Copyright; Avail: Other Sources; Abstract Only The evolving nature of the NASA space enterprise compels the agency to develop new and innovative space systems concepts. NASA, working with increasingly strained budgets and a declining manpower base, is attempting to transform from operational activities to procurement of commercial services. NASA’s current generation reusable launch vehicle, the Shuttle, is in transition from a government owned and operated entity to a commercial venture to reduce the civil servant necessities for that program. NASA foresees its second generation launch vehicles being designed and operated by industry for commercial and government services. The ‘service’ concept is a pioneering effort by NASA. The purpose the ‘service’ is not only to reduce the civil servant overhead but will free up government resources for further research - and enable industry to develop a space business case so that industry can sustain itself beyond government programs. In addition, NASA desires a decreased responsibility thereby decreasing liability. The Second Generation Reusable Launch Vehicle (RLV) program is implementing NASA’s Space Launch Initiative (SLI) to enable industry to develop the launch vehicles of the future. The Alternate Access to Station (AAS) project office within this program is chartered with enabling industry to demonstrate an alternate access capability for the International Space Station (ISS). The project will not accomplish this by traditional government procurement methods, not by integrating the space system within the project office, or by providing the only source of business for the new capability. The project funds will ultimately be used to purchase a service to take re-supply cargo to the ISS, much the same as any business might purchase a service from FedEx to deliver a package to its customer. In the near term, the project will fund risk mitigation efforts for enabling technologies. AAS is in some ways a precursor to the 2nd Generation RLV. By accomplishing ISS resupply with existing technologies, not only will a new category of autonomous vehicles deliver cargo, but a commercial business base will be incubated that will improve the likelihood of commercial convergence with the next generation of RLVs. Traditional paradigms in government management and acquisition philosophy are being challenged in order to bring about the objective of the AAS project. The phased procurement approach is proving to be the most questionable aspect to date. This work addresses the fresh approach AAS is adopting in management and procurement through a study of the AAS history, current solutions, key technologies, procurement complications, and an incremental forward plan leading to the purchase of a service to deliver goods to ISS. Included in this work is a discussion of the Commercial Space Act of 1998 and how it affects government purchase of space launch and space vehicle services. Industry should ﬁnd these topics pertinent to their current state of business. Author Reusable Launch Vehicles; International Space Station; Space Commercialization

20020020362 NASA Marshall Space Flight Center, Huntsville, AL USA Second Generation RLV Space Vehicle Concept Bailey, Michelle; Daniel, Charles; Throckmorton, David A., Technical Monitor; [2002]; In English; 53rd International Aeronautical Congress Conference, 10-19 Oct. 2002, Houston, TX, USA; No Copyright; Avail: Other Sources; Abstract Only NASA has a long history of conducting development programs and projects in a consistent fashion. Systems Engineering within those programs and projects has also followed a given method outlined by such documents as the NASA Systems Engineering Handbook. The relatively new NASA Space Launch Initiative (SLI) is taking a new approach to developing a space vehicle, with innovative management methods as well as new Systems Engineering processes. With the program less than a year into its life cycle, the efficacy of these new processes has yet to be proven or disproven. At $776M for phase 1, SLI represents a major portion of the NASA focus; however, the new processes being incorporated are not reflected in the training provided by NASA to its engineers. The NASA Academy of Program and Project Leadership (APPL) offers core classes in program and project management and systems engineering to NASA employees with the purpose of creating a ‘knowledge community where ideas, skills, and experiences are exchanged to increase each other’s capacity for strong leadership’. The SLI program is, in one sense, a combination of a conceptual design program and a technology program. The program as a whole doesn’t map into the generic systems engineering project cycle as currently, and for some time, taught. For example, the NASA APPL Systems Engineering training course teaches that the ‘first step in developing an architecture is to define the external boundaries of the system’, which will require definition of the interfaces with other systems and the

65 next step will be to ‘define all the components that make up the next lower level of the system hierarchy’ where fundamental requirements are allocated to each component. Whereas, the SLI technology risk reduction approach develops architecture subsystem technologies prior to developing architectures. The higher level architecture requirements are not allowed to fully develop and undergo decomposition and allocation down to the subsystems before the subsystems must develop allocated requirements based on the highest level of requirements. In the vernacular of the project cycles prior to the mid 1990’s, the architecture definition portion of the program appears to be at a generic Phase A stage, while the subsystems are operating at Phase B. Even the management structure of the SLI program is innovative in its approach to Systems Engineering and is not reflected in the APPL training modules. The SLI program has established a Systems Engineering office as an office separate from the architecture development or the subsystem technology development, while that office does have representatives within these other offices. The distributed resources of the Systems Engineering Office are co-located with the respective Project Offices. This template is intended to provide systems engineering as an integrated function at the Program Level. the program management of SLI and the MAT agree that ‘program/project managers and the systems engineering team must work closely together towards the single objective of delivering quality products that meet the customer needs’. This paper will explore the differences between the methods being taught by NASA, which represent decades of ideas, and those currently in practice in SLI. Time will tell if the innovation employed by SLI will prove to be the model of the future. For now, it is suggested that the training of the present exercise the flexibility of recognizing the new processes employed by a major new NASA program. Author Reusable Launch Vehicles; Spacecraft Launching; Management Methods; Systems Engineering

20020019044 NASA Marshall Space Flight Center, Huntsville, AL USA Lightweight Chambers for Thrust Assemblies Elam, S.; Lee, J.; Holmes, R.; Zimmerman, F.; Turner, Jim, Technical Monitor; Jul. 12, 2001; In English; 52nd International Astronautical Congress, 1-5 Oct. 2001, Toulouse, France Contract(s)/Grant(s): NRA8-21 Report No.(s): IAF-01-S.3.05; Copyright; Avail: Other Sources New materials and fabrication techniques have been successfully used to create actively cooled thrust chambers that weigh less, and potentially operate hotter, than conventional designs. Results could improve the performance and payload capacity of many vehicles, including new Reusable Launch Vehicle (RLV) concepts. Polymer Matrix Composite (PMC) and Metal Matrix Composite (MMC) materials replaced traditional metal alloys for structural jackets and manifolds. Ceramic Matrix Composite (CMC) materials and an advanced copper alloy (GRCop-84) were used for chamber liners to significantly increase operating temperatures. Several small chambers were fabricated and tested with promising results to demonstrate that designs over 50% lighter are feasible. New fabrication techniques were demonstrated, including an advanced casting technique and a low cost vacuum plasma spray process. Based on the favorable results for the MMC, PMC and GRCop-84 materials, similar concepts were applied to larger chamber designs. These larger units are currently being fabricated for testing in 2001. Hot-fire testing on the larger units will directly compare performance results between a conventional chamber design and these ‘lightweight’ alternatives. The technology developed and demonstrated by this effort could benefit new RLV programs, as well as other existing and future engine designs. Author Engine Design; Thrust Chambers; Cooling Systems; Design Analysis; Metal Matrix Composites; Combustion Chambers

20020017580 NASA Marshall Space Flight Center, Huntsville, AL USA X-33 XRS-2200 Linear Aerospike Engine Sea Level Plume Radiation DAgostino, Mark G.; Lee, Young C.; Wang, Ten-See; Turner, Jim, Technical Monitor; [2001]; In English; 26th JANNAF Exhaust Plume Technology Subcommittee Meeting, 5-9 Nov. 2001, San Antonio, TX, USA; No Copyright; Avail: CASI; A02, Hardcopy Wide band plume radiation data were collected during ten sea level tests of a single XRS-2200 engine at the NASA Stennis Space Center in 1999 and 2000. The XRS-2200 is a liquid hydrogen/liquid oxygen fueled, gas generator cycle linear aerospike engine which develops 204,420 lbf thrust at sea level. Instrumentation consisted of six hemispherical radiometers and one narrow view radiometer. Test conditions varied from 100% to 57% power level (PL) and 6.0 to 4.5 oxidizer to fuel (O/F) ratio. Measured radiation rates generally increased with engine chamber pressure and mixture ratio. One hundred percent power level radiation data were compared to predictions made with the FDNS and GASRAD codes. Predicted levels ranged from 42% over to 7% under average test values. Author Aerospike Engines; Plumes; X-33 Reusable Launch Vehicle; Heat Flux

66 20020016181 NASA Marshall Space Flight Center, Huntsville, AL USA An Approach to Establishing System Benefits for Technology in NASA’s Hypersonics Investment Area Hueter, Uwe; Pannell, Bill; Cook, Stephen, Technical Monitor; Aug. 17, 2001; In English; 52nd International Astronautical Congress, 1-5 Oct. 2001, Toulouse, France Report No.(s): IAF-01-V.4.05; No Copyright; Avail: CASI; A02, Hardcopy NASA’s has established long term goals for access-to-space. The third generation launch systems are to be fully reusable and operational around 2025. The goals for the third generation launch system are to significantly reduce cost and improve safety over current systems. The Advanced Space Transportation Program (ASTP) Office at the NASA’s Marshall Space Flight Center in Huntsville, AL has the agency lead to develop space transportation technologies. Within ASTP, under the Hypersonics Investment Area, third generation technologies are being pursued. The Hypersonics Investment Area’s primary objective is to mature vehicle technologies to enable substantial increases in the design and operating margins of third generation RLVs (current Space Shuttle is considered the first generation RLV) by incorporating advanced propulsion systems, materials, structures, thermal protection systems, power, and avionics technologies. The paper describes the system process, tools and concepts used to determine the technology benefits. Preliminary results will be presented along with the current technology investments that are being made by ASTP’s Hypersonics Investment Area. Author Hypersonics; Space Transportation; Technology Assessment; Space Shuttles; Propulsion System Performance; NASA Space Programs; Avionics

20020012536 NASA Glenn Research Center, Cleveland, OH USA Testing of Densified Liquid Hydrogen Stratification in a Scale Model Propellant Tank Jurns, John M.; Tomsik, Thomas M.; Greene, William D.; September 2001; In English; 1999 Cryogenic Engineering and International Cryogenic Materials Conference, 12-16 Jul. 1999, Montreal, Quebec, Canada Contract(s)/Grant(s): RTOP 721-20-00 Report No.(s): NASA/TM-2001-209391; NAS 1.15:209391; E-11820; No Copyright; Avail: CASI; A03, Hardcopy This paper describes a test program that was conducted at NASA to demonstrate the ability to load densified LH2 into a subscale propellant tank. This work was done through a collaborative effort between NASA Glenn Research Center and the Lockheed Martin Michoud Space Systems (LMMSS). The Multilobe tank, which was made from composite materials similar to that to be used on X-33, was formed from two lobes with a center septum. Test results are shown for data that was collected on filling the subscale tank with densified liquid hydrogen (DLH2) propellant that was produced at the NASA Plum Brook Station. Data is compared to analytical predictions. Data collected for this test series agrees well with analytical predictions of the environmental heat leak into the tank and the thermal stratification characteristics of the hydrogen propellant in the tank as it was filled with DLH2. Author Liquid Hydrogen; Propellant Tanks; Scale Models; Stratification; Cryogenic Fluid Storage; Performance Tests

20020012437 NASA Ames Research Center, Moffett Field, CA USA The Evolution of Flexible Insulation as Thermal Protection Systems for Reusable Launch Vehicles: AFRSI (Advanced Flexible Reusable Surface Insulation) to CRI (Conformal Reusable Insulation) Rezin, Marc; Oka, Kris; Arnold, Jim, Technical Monitor; [2001]; In English; National Space and Missile Materials Symposium, 25 Jun. 2001; No Copyright; Avail: CASI; A03, Hardcopy This viewgraph presentation gives an overview of the evolution of flexible insulation as a thermal protection system for reusable launch vehicles, Advanced Flexible Reusable Surface Insulation (AFRSI) to Conformal Reusable Insulation (CRI). Details are given on the approved use of AFRIS on the Shuttle Orbiter in June 1980, the first flight of AFRIS on STS-6, windward blanket development, composite flexible blanket insulation, and flight demonstrations. CASI Thermal Insulation; Space Shuttle Orbiters

20020008402 NASA, Washington, DC USA NASA IVHM Technology Experiment for X-vehicles (NITEX) Sandra, Hayden; Bajwa, Anupa; [2001]; In English; No Copyright; Avail: CASI; A01, Hardcopy The purpose of the NASA IVHM Technology Experiment for X-vehicles (NITEX) is to advance the development of selected IVHM technologies in a ﬂight environment and to demonstrate the potential for reusable launch vehicle ground

67 processing savings. The technologies to be developed and demonstrated include system-level and detailed diagnostics for real-time fault detection and isolation, prognostics for fault prediction, automated maintenance planning based on diagnostic and prognostic results, and a microelectronics hardware platform. Complete ﬂight The Evolution of Flexible Insulation as IVHM consists of advanced sensors, distributed data acquisition, data processing that includes model-based diagnostics, prognostics and vehicle autonomy for control or suggested action, and advanced data storage. Complete ground IVHM consists of evolved control room architectures, advanced applications including automated maintenance planning and automated ground support equipment. This experiment will advance the development of a subset of complete IVHM. Derived from text Reusable Launch Vehicles; Diagnosis; Distributed Processing; Fault Detection; Ground Based Control; Real Time Operation; Technology Assessment

20010110022 Oklahoma State Univ., Stillwater, OK USA Using Neural Networks in Decision Making for a Reconfigurable Electro Mechanical Actuator (EMA) Latino, Carl D.; 2000 Research Reports: NASA/ASEE Summer Faculty Fellowship Program; October 2001; In English; No Copyright; Avail: CASI; A01, Hardcopy The objectives of this project were to demonstrate applicability and advantages of a neural network approach for evaluating the performance of an electro-mechanical actuator (EMA). The EMA in question was intended for the X-37 Advanced Technology Vehicle. It will have redundant components for safety and reliability. The neural networks for this application are to monitor the operation of the redundant electronics that control the actuator in real time and decide on the operating configuration. The system we proposed consists of the actuator, sensors, control circuitry and dedicated (embedded) processors. The main purpose of the study was to develop suitable hardware and neural network capable of allowing real time reconfiguration decisions to be made. This approach was to be compared to other methods such as fuzzy logic and knowledge based systems considered for the same application. Over the course of the project a more general objective was the identification of the other neural network applications and the education of interested NASA personnel on the topic of Neural Networks. Author Electromechanical Devices; Actuators; Neural Nets; Decision Making; Redundant Components; Real Time Operation

20010106933 NASA Ames Research Center, Moffett Field, CA USA Wireless Subsurface Sensors for Health Monitoring of Thermal Protection Systems on Hypersonic Vehicles Milos, Frank S.; Arnold, Jim, Technical Monitor; [2001]; In English; 3rd International Workshop on Structural Health Monitoring, 12-14 Sep. 2001, Stanford, CA, USA; No Copyright; Avail: Other Sources; Abstract Only Health diagnostics is an area where major improvements have been identified for potential implementation into the design of new reusable launch vehicles (RLVs) in order to reduce life cycle costs, to increase safety margins, and to improve mission reliability. NASA Ames is leading the effort to develop inspection and health management technologies for thermal protection systems. This paper summarizes a joint project between NASA Ames and industry partners to develop ‘wireless’ devices that can be embedded in the thermal protection system to monitor temperature or other quantities of interest. These devices are sensors integrated with radio-frequency identification (RFID) microchips to enable non-contact communication of sensor data to an external reader that may be a hand-held scanner or a large portal. Both passive and active prototype devices have been developed. The passive device uses a thermal fuse to indicate the occurrence of excessive temperature. This device has a diameter under 0.13 cm. (suitable for placement in gaps between ceramic TPS tiles on an RLV) and can withstand 370 C for 15 minutes. The active device contains a small battery to provide power to a thermocouple for recording a temperature history during flight. The bulk of the device must be placed beneath the TPS for protection from high temperature, but the thermocouple can be placed in a hot location such as near the external surface. Author Thermal Protection; Hypersonic Vehicles; Reusable Launch Vehicles; Diagnosis

20010106932 NASA Ames Research Center, Moffett Field, CA USA Thermal Protection Materials for Reentry and Planetary Applications Johnson, Sylvia M.; Arnold, Jim, Technical Monitor; [2001]; In English; Conference on Structural Ceramics and Ceramic Composites for High Temperature Applications, Oct. 2001, Spain; No Copyright; Avail: Other Sources; Abstract Only Thermal protection materials and systems (TPS) are used to protect spacecraft during reentry into Earth’s atmosphere or entry into planetary atmospheres. As such, these materials are subject to severe environments with high heat ﬂuxes and rapid

68 heating. Catalytic effects can increase the temperatures substantially. These materials are also subject to impact damage from micrometeorites or other debris during ascent, orbit, and descent, and thus must be able to withstand damage and function following damage. Thermal protection materials and coatings used in reusable launch vehicles will be reviewed, including the needs and directions for new materials to enable new missions that require faster turnaround and much greater reusability. The role of ablative materials for use in high heat ﬂux environments, especially for non-reusable applications and upcoming planetary missions, will be discussed. New thermal protection system materials may enable the use of sharp nose caps and leading edges on future reusable space transportation vehicles. Vehicles employing this new technology would have signiﬁcant increases in maneuverability and out-of-orbit cross range compared to current vehicles, leading to increased mission safety in the event of the need to abort during ascent or from orbit. Ultrahigh temperature ceramics, a family of materials based on HfB2and ZrB2 with SiC, will be discussed. The development, mechanical and thermal properties, and uses of these materials will be reviewed. Author Thermal Protection; Heat Flux; Reusable Launch Vehicles; Atmospheric Entry; Heating; Planetary Atmospheres

20010098302 NASA Langley Research Center, Hampton, VA USA Structures for the 3rd Generation Reusable Concept Vehicle Hrinda, Glenn A.; [2001]; In English; AIAA Space 2001 Conference and Exposition, 28-30 Aug. 2001, Albuquerque, NM, USA; Copyright; Avail: CASI; A02, Hardcopy A major goal of NASA is to create an advance space transportation system that provides a safe, affordable highway through the air and into space. The long-term plans are to reduce the risk of crew loss to 1 in 1,000,000 missions and reduce the cost of Low-Earth Orbit by a factor of 100 from today’s costs. A third generation reusable concept vehicle (RCV) was developed to assess technologies required to meet NASA’s space access goals. The vehicle will launch from Cape Kennedy carrying a 25,000 lb. payload to the International Space Station (ISS). The system is an air breathing launch vehicle (ABLV) hypersonic lifting body with rockets and uses triple point hydrogen and liquid oxygen propellant. The focus of this paper is on the structural concepts and analysis methods used in developing the third generation reusable launch vehicle (RLV). Member sizes, concepts and material selections will be discussed as well as analysis methods used in optimizing the structure. Analysis based on the HyperSizer structural sizing software will be discussed. Design trades required to optimize structural weight will be presented. Author Structural Design; Reusable Spacecraft; Spacecraft Design; Design Optimization; Structural Analysis

20010097314 NASA Ames Research Center, Moffett Field, CA USA Wireless Subsurface Microsensors for Health Monitoring of Thermal Protection Systems on Hypersonic Vehicles Milos, Frank S.; Watters, David G.; Pallix, Joan B.; Bahr, Alfred J.; Huestis, David L.; Arnold, Jim, Technical Monitor; [2001]; In English; SPIE 6th Annual International Symposium on Health Monitoring and Diagnostics, 4-8 Mar. 2001, Newport Beach, CA, USA Contract(s)/Grant(s): NAS2-99092; NAS2-10730; NAS2-10731; No Copyright; Avail: CASI; A02, Hardcopy Health diagnostics is an area where major improvements have been identiﬁed for potential implementation into the design of new reusable launch vehicles in order to reduce life cycle costs, to increase safety margins, and to improve mission reliability. NASA Ames is leading the effort to develop inspection and health management technologies for thermal protection systems. This paper summarizes a joint project between NASA Ames and SRI International to develop ‘SensorTags,’ radio frequency identiﬁcation devices coupled with event-recording sensors, that can be embedded in the thermal protection system to monitor temperature or other quantities of interest. Two prototype SensorTag designs containing thermal fuses to indicate a temperature overlimit are presented and discussed. Author Damage Assessment; Microinstrumentation; Hypersonic Vehicles; Smart Materials; Thermal Protection; Wireless Communication

20010095443 Science Applications International Corp., Huntsville, AL USA Pathﬁnder Technologies Specialist, X-37 French, James R.; Sep. 30, 2001; In English Contract(s)/Grant(s): NAS8-99060; No Copyright; Avail: CASI; A04, Hardcopy The X-37 is a technology demonstrator sponsored by NASA. It includes a number of experiments both imbedded (i.e.,

69 essential aspects of the vehicle) and separate. The technologies demonstrated will be useful in future operational versions as well as having broad applications to other programs. Mr. James R. French, of JRF Engineering Services and as a consultant to SAIC, has provided technical support to the X-37 NASA Program office since the beginning of the program. In providing this service, Mr. French has maintained close contact with the Boeing Seal Beach and Rocketdyne technical teams via telephone, e-mail, and periodic visits. His interfaces were primarily with the working engineers in order to provide NASA sponsors with a different view than that achieved through management channels. Mr. French’s periodic and highly detailed technical reports were submitted to NASA and SAIC (Science Applications International Corporation) on a weekly/monthly basis. These reports addressed a wide spectrum of programmatic and technical interests related to the X-37 Program including vehicle design, ﬂight sciences, propulsion, thermal protection, Guidance Navigation & Control (GN&C), structures, and operations. This deliverable is presented as a consolidation of the twelve monthly reports submitted during the Contract’s Option Year, Author X-37 Vehicle; Research and Development; NASA Programs; Contract Management; Technology Utilization

20010095419 Science Applications International Corp., Huntsville, AL USA Space Transportation Systems Technologies Laue, Jay H.; Sep. 30, 2001; In English Contract(s)/Grant(s): NAS8-99060; NRA8-21; No Copyright; Avail: CASI; A05, Hardcopy This document is the final report by the Science Applications International Corporation (SAIC) on contracted support provided to the National Aeronautics and Space Administration (NASA) under Contract NAS8-99060, ‘Space Transportation Systems Technologies’. This contract, initiated by NASA’s Marshall Space Flight Center (MSFC) on February 8, 1999, was focused on space systems technologies that directly support NASA’s space flight goals. It was awarded as a Cost-Plus- Incentive-Fee (CPIF) contract to SAIC, following a competitive procurement via NASA Research Announcement, NRA 8-21. This NRA was specifically focused on tasks related to Reusable Launch Vehicles (RLVs). Through Task Area 3 (TA-3), ‘Other Related Technology’ of this NRA contract, SAIC extensively supported the Space Transportation Directorate of MSFC in effectively directing, integrating, and setting its mission, operations, and safety priorities for future RLV-focused space flight. Following an initially contracted Base Year (February 8, 1999 through September 30, 1999), two option years were added to the contract. These were Option Year 1 (October 1, 1999 through September 30, 2000) and Option Year 2 (October 1, 2000 through September 30, 2001). This report overviews SAIC’s accomplishments for the Base Year, Option Year 1, and Option Year 2, and summarizes the support provided by SAIC to the Space Transportation Directorate, NASA/MSFC. Author Aerospace Systems; Reusable Launch Vehicles; Space Transportation System

20010092480 NASA Glenn Research Center, Cleveland, OH USA Performance Evaluation of the NASA GTX RBCC Flowpath Thomas, Scott R.; Palac, Donald T.; Trefny, Charles J.; Roche, Joseph M.; July 2001; In English; Fifteenth International Symposium on Airbreathing Engines, 2-7 Sep. 2001, Bangalore, India; Original contains color illustrations Contract(s)/Grant(s): RTOP 708-73-20 Report No.(s): NASA/TM-2001-210953; E-12807; NAS 1.15:210953; ISABE-2001-1070; No Copyright; Avail: CASI; A02, Hardcopy The NASA Glenn Research Center serves as NASAs lead center for aeropropulsion. Several programs are underway to explore revolutionary airbreathing propulsion systems in response to the challenge of reducing the cost of space transportation. Concepts being investigated include rocket-based combined cycle (RBCC), pulse detonation wave, and turbine-based combined cycle (TBCC) engines. The GTX concept is a vertical launched, horizontal landing, single stage to orbit (SSTO) vehicle utilizing RBCC engines. The propulsion pod has a nearly half-axisymmetric flowpath that incorporates a rocket and ram-scramjet. The engine system operates from lift-off up to above Mach 10, at which point the airbreathing engine flowpath is closed off, and the rocket alone powers the vehicle to orbit. The paper presents an overview of the research efforts supporting the development of this RBCC propulsion system. The experimental efforts of this program consist of a series of test rigs. Each rig is focused on development and optimization of the flowpath over a specific operating mode of the engine. These rigs collectively establish propulsion system performance over all modes of operation, therefore, covering the entire speed range. Computational Fluid Mechanics (CFD) analysis is an important element of the GTX propulsion system development and validation. These efforts guide experiments and flowpath design, provide insight into experimental data, and extend results to

70 conditions and scales not achievable in ground test facilities. Some examples of important CFD results are presented. Author Air Breathing Engines; Computational Fluid Dynamics; Propulsion System Conﬁgurations; Propulsion System Performance

20010091019 NASA Ames Research Center, Moffett Field, CA USA Health Monitoring Technology for Thermal Protection Systems on Reusable Hypersonic Vehicles Milos, Frank S.; Watters, D. G.; Heinemann, J. M.; Karunaratne, K. S.; Arnold, Jim, Technical Monitor; [2001]; In English; National Space and Missile Materials Symposium, 25-28 Jun. 2001, Monterey, CA, USA; No Copyright; Avail: CASI; A03, Hardcopy Integrated subsystem health diagnostics is an area where major improvements have been identified for potential implementation into the design of new reusable launch vehicles (RLVs) in order to reduce life cycle costs, to increase safety margins, and to improve mission reliability. This talk summarizes a joint effort between NASA Ames and industry partners to develop rapid non-contact diagnostic tools for health and performance monitoring of thermal protection systems (TPS) on future RLVs. The specific goals for TPS health monitoring are to increase the speed and reliability of TPS inspections for improved operability at lower cost. The technology being developed includes a 3-D laser scanner for examining the exterior surface of the TPS, and a subsurface microsensor suite for monitoring the health and performance of the TPS. The sensor suite consists of passive overlimit sensors and sensors for continuous parameter monitoring in flight. The sensors are integrated with radio-frequency identification (RFID) microchips to enable wireless communication of-the sensor data to an external reader that may be a hand-held scanner or a large portal. Prototypes of the laser system and both types of subsurface sensors have been developed. The laser scanner was tested on Shuttle Orbiter Columbia and was able to dimension surface chips and holes on a variety of TPS materials. The temperature-overlimit microsensor has a diameter under 0.05 inch (suitable for placement in gaps between ceramic TPS tiles) and can withstand 700 F for 15 minutes. Author Thermal Protection; Microinstrumentation; Diagnosis; Chips (Electronics); Ceramics; Systems Health Monitoring

20010086424 Science Applications International Corp., Huntsville, AL USA Second Generation RLV Program Laue, Jay; Feb. 26, 2001; In English; Original contains color illustrations Contract(s)/Grant(s): NAS8-99060; NRA8-21 Report No.(s): NAS8-99060-MPR26; No Copyright; Avail: CASI; A03, Hardcopy During the time period covered by this report, SAIC: 1) Continued to develop and assess processes and approaches that can be applied to Second Generation Reuseable Launch Vehicles (RLV) technologies prioritization. An approach based on the use of analytic Saaty scale functions has been defined and is being investigated. 2) Planned and facilitated technologies prioritization workshops, supported development of systems program algorithms based on the concept of influence diagramming, and assessment of analogies between aircraft and space systems developments. 3) Video interviews held with X-37 personnel at Dryden. The CD-ROM is being concluded and a near-final review disc is expected soon. 4) CD-ROMS were produced by Engineered Multimedia, Inc. (EMI) under subcontract to SAIC. These two CD-ROM products, ‘Microgravity’ and ‘New Horizons’, were delivered to MSFS technical representatives at a briefing at SAID on January 18, 2001. 5) Presentation brochure, ‘Aviation/Space Analog Team Interim Report’ was provided to Space Propulsion Synergy Team (SPST) members and to NASA personnel. Wrap-up of effort will be a mid-April briefing to both MSFC and the full SPST membership. Support to Phase 2 of CCPD ends 3/31/01. Phase 3 continuation effort planned with emphasis on technical info content. 6) Recommended that eliminating the B-52 flights of the X-37 in favor of alternate approached be evaluated. Discovered that a required change to the thruster valves of the X-37 had not been made. Repair work continues on the lower fuselage section of the X-37, in the areas that experienced core collapse. The currently catalogued potential weight increases and decreases are about equal, but this does not include the impact of the cable weight underestimate reported earlier. The CFD for RCS testing correlates well with the wind tunnel data. 7) The Level IV CCB approved SCN-4 to the MC-1 Engine SPecification. Approved changes have been incorporated into the specification, and Revision D has been released to the Repository. Billy Gonterman, SAIC, traveled to SSC to perform a five-day readiness assessment on the progress of facilities activation. Implemented modifications to the MC-1 Engine TPA Component Database. 8) Replaced 25 Combined-Cycle Propulsion Database (CCPD) documents; modified the low-on screen to include a direct email link for user contact; prepared a summary presentation of current and proposed tasks. Derived from text Reusable Launch Vehicles; Microgravity; X-37 Vehicle; Combined Cycle Power Generation; Spacecraft Propulsion; Systems Engineering; Rocket Engines

71 20010084776 NASA Dryden Flight Research Center, Edwards, CA USA Full Envelope Reconfigurable Control Design for the X-33 Vehicle Cotting, M. Christopher; Burken, John J.; Lee, Seung-Hee, Technical Monitor; [2001]; In English; AIAA Guidance, Navigation and Control Conference, 6-10 Aug. 2001, Montreal, Quebec, Canada Report No.(s): AIAA Paper 2001-1234; Copyright; Avail: CASI; A02, Hardcopy In the event of a control surface failure, the purpose of a reconfigurable control system is to redistribute the control effort among the remaining working surfaces such that satisfactory stability and performance are retained. An Off-line Nonlinear General Constrained Optimization (ONCO) approach was used for the reconfigurable X-33 control design method. Three example failures are shown using a high fidelity 6 DOF simulation (case I ascent with a left body flap jammed at 25 deg.; case 2 entry with a right inboard elevon jam at 25 deg.; and case 3, landing (TAEM) with a left rudder jam at -30 deg.) Failure comparisons between responses with the nominal controller and reconfigurable controllers show the benefits of reconfiguration. Single jam aerosurface failures were considered, and failure detection and identification is considered accomplished in the actuator controller. The X-33 flight control system will incorporate reconfigurable flight control in the baseline system. Author X-33 Reusable Launch Vehicle; Flight Control; Control Systems Design; Failure; Design Analysis; Aerodynamic Configurations; Aircraft Design; Controllers

20010081320 NASA Dryden Flight Research Center, Edwards, CA USA Reconfigurable Control Design for the Full X-33 Flight Envelope Cotting, M. Christopher; Burken, John J.; August 2001; In English; AIAA Guidance, Navigation and Control Conference, 6-10 Aug. 2001, Montreal, Quebec, Canada Contract(s)/Grant(s): RTOP 715-33-02-E8-23 Report No.(s): NASA/TM-2001-210396; H-2458; NAS 1.15:210396; AIAA Paper 2001-4379; No Copyright; Avail: CASI; A03, Hardcopy A reconfigurable control law for the full X-33 flight envelope has been designed to accommodate a failed control surface and redistribute the control effort among the remaining working surfaces to retain satisfactory stability and performance. An offline nonlinear constrained optimization approach has been used for the X-33 reconfigurable control design method. Using a nonlinear, six-degree-of-freedom simulation, three example failures are evaluated: ascent with a left body flap jammed at maximum deflection; entry with a right inboard elevon jammed at maximum deflection; and landing with a left rudder jammed at maximum deflection. Failure detection and identification are accomplished in the actuator controller. Failure response comparisons between the nominal control mixer and the reconfigurable control subsystem (mixer) show the benefits of reconfiguration. Single aerosurface jamming failures are considered. The cases evaluated are representative of the study conducted to prove the adequate and safe performance of the reconfigurable control mixer throughout the full flight envelope. The X-33 flight control system incorporates reconfigurable flight control in the existing baseline system. Author X-33 Reusable Launch Vehicle; Configuration Management; Control Systems Design; Flight Envelopes; Flight Control; Control Theory

20010080460 NASA Dryden Flight Research Center, Edwards, CA USA A Monte Carlo Dispersion Analysis of the X-33 Simulation Software Williams, Peggy S.; [2001]; In English; AIAA Atmospheric Flight Mechanics Conference, 6-9 Aug. 2001, Montreal, Canada Report No.(s): AIAA Paper 2001-4067; Copyright; Avail: CASI; A03, Hardcopy A Monte Carlo dispersion analysis has been completed on the X-33 software simulation. The simulation is based on a preliminary version of the software and is primarily used in an effort to define and refine how a Monte Carlo dispersion analysis would have been done on the final flight-ready version of the software. This report gives an overview of the processes used in the implementation of the dispersions and describes the methods used to accomplish the Monte Carlo analysis. Selected results from 1000 Monte Carlo runs are presented with suggestions for improvements in future work. Author Computerized Simulation; Monte Carlo Method; Computer Programs; Software Development Tools; Dispersion

20010080458 NASA Marshall Space Flight Center, Huntsville, AL USA Evaluation of Microcracking in Two Carbon-Fiber/Epoxy-Matrix Composite Cryogenic Tanks Hodge, A. J.; August 2001; In English; Original contains color illustrations Report No.(s): NASA/TM-2001-211194; M-1024; NAS 1.15:211194; No Copyright; Avail: CASI; A03, Hardcopy

72 Two graphite/epoxy cryogenic pressure vessels were evaluated for microcracking. The X-33 LH2 tank lobe skins were extensively examined for microcracks. Specimens were removed from the inner skin of the X-33 tank for tensile testing. The data obtained from these tests were used to model expected microcrack density as a function of stress. Additionally, the laminate used in the Marshall Space Flight Center (MSFC) Composite Conformal, Cryogenic, Common Bulkhead, Aerogel-Insulated Tank (CBAT) was evaluated. Testing was performed in an attempt to predict potential microcracking during testing of the CBAT. Author Graphite-Epoxy Composites; Pressure Vessels; X-33 Reusable Launch Vehicle; Storage Tanks; Cryogenic Fluid Storage; Microcracks; Liquid Hydrogen

20010079105 NASA Marshall Space Flight Center, Huntsville, AL USA National Rocket Propulsion Materials Plan: A NASA, Department of Defense, and Industry Partnership Clinton, Raymond G., Jr.; Munafo, Paul M., Technical Monitor; [2001]; In English; 2001 ASM/TMS Spring Symposium, 1-2 May 2001, Schenectady, NY, USA; No Copyright; Avail: Other Sources; Abstract Only NASA, Department of Defense, and rocket propulsion industry representatives are working together to create a national rocket propulsion materials development roadmap. This ‘living document’ will facilitate collaboration among the partners, leveraging of resources, and will be a highly effective tool for technology development planning. The structuring of the roadmap, and development plan, which will combine the signiﬁcant efforts of the Integrated High Payoff Rocket Propulsion Technology (IHPRPT) Program, and NASA’s Integrated Space Transportation Plan (ISTP), is being lead by the IHPRPT Materials Working Group (IMWG). The IHPRPT Program is a joint DoD, NASA, and industry effort to dramatically improve the nation’s rocket propulsion capabilities. This phased program is structured with increasingly challenging goals focused on performance, reliability, and cost to effectively double rocket propulsion capabilities by 2010. The IHPRPT program is focused on three propulsion application areas: Boost and Orbit Transfer (both liquid rocket engines and solid rocket motors), Tactical, and Spacecraft. Critical to the success of this initiative is the development and application of advanced materials, processes, and manufacturing technologies. NASA’s ISTP is a comprehensive strategy focusing on the aggressive safety, reliability, and affordability goals for future space transportation systems established by the agency. Key elements of this plan are the 2 nd and 3 d Generation Reusable Launch Vehicles (RLV). The affordability and safety goals of these generational systems are, respectively, 10X cheaper and 100X safer by 2010, and 100X cheaper and 10,000X safer by 2025. Accomplishment of these goals requires dramatic and sustained breakthroughs, particularly in the development and the application of advanced material systems. The presentation will provide an overview of the IHPRPT materials initiatives, NASA’s 2nd and 3 rd Generation RLV propulsion materials projects, and the approach for the development of the national rocket propulsion materials roadmap. Author Technological Forecasting; Solid Propellant Rocket Engines; Safety; Reliability; Propulsion; Planning; Liquid Propellant Rocket Engines

20010071334 NASA Langley Research Center, Hampton, VA USA High Temperature Polyimide Materials in Extreme Temperature Environments Johnson, Theodore F.; Gates, Thomas S.; [2001]; In English; 42nd AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference and Exhibit, 16-19 Apr. 2001, Seattle, WA, USA Report No.(s): AIAA Paper 2001-1214; Copyright; Avail: CASI; A03, Hardcopy At the end of the NASA High Speed Research (HSR) Program, NASA Langley Research Center (LaRC) began a program to screen the high-temperature Polymeric Composite Materials (PMCs) characterized by the HSR Durability Program for possible use in Reusable Launch Vehicles (RLVs) operating under extreme temperature conditions. The HSR Program focused on developing material-related technologies to enable a High Speed Civil Transport (HSCT) capable of operating temperatures ranging from 54 C (-65 F) to 177 C (350 F). A high-temperature polymeric resin, PETI-5 was used in the HSR Program to satisfy the requirements for performance and durability for a PMC. For RLVs, it was anticipated that this high temperature material would contribute to reducing the overall weight of a vehicle by eliminating or reducing the thermal protection required to protect the internal structural elements of the vehicle and increasing the structural strain limits. The tests were performed to determine temperature-dependent mechanical and physical proper-ties of IM7/PETI-5 composite over a temperature range from cryogenic temperature -253 C (-423F) to the material’s maximum use temperature of 230 C (450 F). This paper presents results from the test program for the temperature-dependent mechanical and physical properties of IM7/PETI-5 composite in the temperature range from -253 C (-423 F) to 27 C (80 F). Author High Temperature; Reusable Launch Vehicles; Polyimides; Refractory Materials; Supersonic Transports; Cryogenic Temperature

73 20010067392 NASA Marshall Space Flight Center, Huntsville, AL USA Predictor-Corrector Entry Guidance for Reusable Launch Vehicles Youssef, Hussein; Chowdhry, Rajiv; Lee, Howard; Zimmerman, Curtis; Brandon, Larry, Technical Monitor; [2001]; In English; AIAA GNC Conference, 6-9 Aug. 2001, Montreal, Canada Contract(s)/Grant(s): NCC8-115; No Copyright; Avail: CASI; A02, Hardcopy An online entry guidance algorithm has been developed using a predictor-corrector approach. The algorithm is designed for the Reusable Launch Vehicle (RLV) and is demonstrated by using, the X-33 model. The objective of the design is to handle widely dispersed entry conditions and deliver the vehicle at the Terminal Area Energy Management (TAEM) interface box within an acceptable tolerance and without violating any of the vehicle physical constraints. Combination of several control variables is used in testing the performance and computational requirement of the algorithm. The control variables are the bank angle, angle-of-attack and the time for roll reversal. The bank angle and angle-of-attack profiles are the nominal profiles plus the perturbations in each direction. The initial guess of the bank profile is a 45 degrees bank angle with reversal at 360 seconds from liftoff. A six-element state vector is propagated to the TAEM interface box through the integration of the equations of motion (EOM). Altitude, heading and range errors are computed between the desired and the achieved state at the TAEM interface. These errors are used to correct the initial guess of the control variables. This process is repeated until the errors meet an acceptable level at the TAEM interface. Several numerical optimization methods are used to evaluate the convergent property of the predictor-predictor methodology. Successful results are demonstrated using the X-33 model. Author X-33 Reusable Launch Vehicle; Predictor-Corrector Methods; Numerical Analysis; Algorithms

20010067274 NASA Marshall Space Flight Center, Huntsville, AL USA X-33 LH2 Tank Failure Investigation Findings Niedermeyer, M.; Proceedings of The 4th Conference on Aerospace Materials, Processes, and Environmental Technology; February 2001; In English; No Copyright; Avail: CASI; A03, Hardcopy The X-33 liquid hydrogen tank failure investigation found the following: (1) The inner skin microcracked and hydrogen inﬁltrated into it; (2) The cracks grew larger under pressure; (3) When pressure was removed, the cracks closed slightly; (4) When the tank was drained and warmed, the cracks closed and blocked the leak path; (5) Foreign object debris (FOD) and debond areas provided an opportunity for a leak path; and (6) There is still hydrogen in the other three lobes today. Derived from text Microcracks; X-33 Reusable Launch Vehicle; Accident Investigation; Propellant Tanks

20010067271 NASA Marshall Space Flight Center, Huntsville, AL USA Development of Metal Matrix Composites for NASA’s Advanced Propulsion Systems Lee, J.; Elam, S.; Proceedings of The 4th Conference on Aerospace Materials, Processes, and Environmental Technology; February 2001; In English; No Copyright; Avail: CASI; A02, Hardcopy The state-of-the-art development of several Metal Matrix Composites (MMC) for NASA’s advanced propulsion systems will be presented. The goal is to provide an overview of NASA-Marshall Space Flight Center’s on-going activities in MMC components for advanced liquid rocket engines such as the X-33 vehicle’s Aerospike engine and X-34’s Fastrac engine. The focus will be on lightweight, low cost, and environmental compatibility with oxygen and hydrogen of key MMC materials, within each of NASA’s new propulsion application, that will provide a high payoff for NASA’s Reusable Launch Vehicles and space access vehicles. In order to fabricate structures from MMC, effective joining methods must be developed to join MMC to the same or to different monolithic alloys. Therefore, a qualitative assessment of MMC’s welding and joining techniques will be outlined. Author Metal Matrix Composites; Welding; Joints (Junctions)

20010067262 NASA Marshall Space Flight Center, Huntsville, AL USA Composite,Cryogenic, Conformal, Common Bulkhead, Aerogel-Insulated Tank (CBAT) Roberts, J. K.; Kovach, M. P.; McMahon, W. M.; Finckenor, J. L.; Proceedings of The 4th Conference on Aerospace Materials, Processes, and Environmental Technology; February 2001; In English; Copyright; Avail: Other Sources The objective of the Composite, Cryogenic, Conformal, Common Bulkhead, Aerogel-insulated Tank (CBAT) Program is to evaluate the potential for using various new technologies in next generation Reusable Launch Vehicles (RLVs) through design, fabrication, and testing of a subscale system. The new technologies include polymer matrix composites (PMCs),

74 conformal propellant storage, common bulkhead packaging, and aerogel insulation. The National Aeronautics and Space Administration (NASA) and Thiokol Propulsion from Cordant Technologies are working together to develop a design and the processing methodologies which will allow integration of these technologies into a single structural component assembly. Such integration will signiﬁcantly decrease subsystem weight and reduce shape, volume, and placement restrictions, thereby enhancing overall launch system performance. This paper/presentation focuses on the challenges related to materials and processes that were encountered and overcome during this program to date. Author Technology Assessment; Reusable Launch Vehicles; Technology Utilization; Fabrication; Propellant Tanks

20010067258 NASA Marshall Space Flight Center, Huntsville, AL USA Fabrication and Testing of Ceramic Matrix Composite Rocket Propulsion Components Effinger, M. R.; Clinton, R. C., Jr.; Dennis, J.; Elam, S.; Genge, G.; Eckel, A.; Jaskowiak, M. H.; Kiser, J. D.; Lang, J.; Proceedings of The 4th Conference on Aerospace Materials, Processes, and Environmental Technology; February 2001; In English; No Copyright; Avail: CASI; A03, Hardcopy NASA has established goals for Second and Third Generation Reusable Launch Vehicles. Emphasis has been placed on signiﬁcantly improving safety and decreasing the cost of transporting payloads to orbit. Ceramic matrix composites (CMC) components are being developed by NASA to enable signiﬁcant increases in safety and engineer performance, while reducing costs. The development of the following CMC components are being pursued by NASA: (1) Simplex CMC Blisk; (2) Cooled CMC Nozzle Ramps; (3) Cooled CMC Thrust Chambers; and (4) CMC Gas Generator. These development efforts are application oriented, but have a strong underpinning of fundamental understanding of processing-microstructure-property relationships relative to structural analyses, nondestructive characterization, and material behavior analysis at the coupon and component and system operation levels. As each effort matures, emphasis will be placed on optimizing and demonstrating material/component durability, ideally using a combined Building Block Approach and Build and Bust Approach. Derived from text Ceramic Matrix Composites; Reusable Launch Vehicles; Fabrication; Performance Tests; Engine Parts

20010067250 NASA Marshall Space Flight Center, Huntsville, AL USA Large Composite Structures Processing Technologies for Reusable Launch Vehicles Clinton, R. C., Jr.; Vickers, J. H.; McMahon, W. M.; Hulcher, A. B.; Johnston, N. J.; Cano, R. J.; Belvin, H. L.; McIver, K.; Franklin, W.; Sidwell, D., et al.; Proceedings of The 4th Conference on Aerospace Materials, Processes, and Environmental Technology; February 2001; In English; No Copyright; Avail: CASI; A03, Hardcopy Signiﬁcant efforts have been devoted to establishing the technology foundation to enable the progression to large scale composite structures fabrication. We are not capable today of fabricating many of the composite structures envisioned for the second generation reusable launch vehicle (RLV). Conventional ‘aerospace’ manufacturing and processing methodologies (ﬁber placement, autoclave, tooling) will require substantial investment and lead time to scale-up. Out-of-autoclave process techniques will require aggressive efforts to mature the selected technologies and to scale up. Focused composite processing technology development and demonstration programs utilizing the building block approach are required to enable envisioned second generation RLV large composite structures applications. Government/industry partnerships have demonstrated success in this area and represent best combination of skills and capabilities to achieve this goal. Derived from text Composite Structures; Fabrication; Reusable Launch Vehicles

20010067229 NASA Marshall Space Flight Center, Huntsville, AL USA Thermographic Analysis of Composite Cobonds on the X-33 Russell, S. S.; Walker, J. L.; Lansing, M. D.; Proceedings of The 4th Conference on Aerospace Materials, Processes, and Environmental Technology; February 2001; In English; No Copyright; Avail: CASI; A02, Hardcopy During the manufacture of the X-33 liquid hydrogen (LH2) Tank 2, a total of 36 reinforcing caps were inspected thermographically. The cured reinforcing sheets of graphite/epoxy were bonded to the tank using a wet cobond process with vacuum bagging and low temperature curing. A foam ﬁller material wedge separated the reinforcing caps from the outer skin of the tank. Manufacturing difficulties caused by a combination of the size of the reinforcing caps and their complex geometry lead to a potential for trapping air in the bond line. An inspection process was desired to ensure that the bond line was free of voids before it had cured so that measures could be taken to rub out the entrapped air or remove the cap and perform additional surface matching. Infrared thermography was used to perform the procure ‘wet bond’ inspection as well a to

75 document the final ‘cured’ condition of the caps. The thermal map of the bond line was acquired by heating the cap with either a flash lamp or a set of high intensity quartz lamps and then viewing it during cool down. The inspections were performed through the vacuum bag and voids were characterized by localized hot spots. In order to ensure that the cap had bonded to the tank properly, a post cure ‘flash heating’ thermographic investigation was performed with the vacuum bag removed. Any regions that had opened up after the preliminary inspection or that were hidden during the bagging operation were marked and filled by drilling small holes in the cap and injecting resin. This process was repeated until all critical sized voids were filled. Author Thermography; Voids; Joints (Junctions); Infrared Inspection; Nondestructive Tests

20010063955 NASA Dryden Flight Research Center, Edwards, CA USA Two Reconfigurable Flight-Control Design Methods: Robust Servomechanism and Control Allocation Burken, John J.; Lu, Ping; Wu, Zheng-Lu; Bahm, Cathy; Journal of Guidance, Control and Dynamics; May-June 2001; ISSN 0731-5090; Volume 24, No. 3; In English; AIAA Guidance Navigation and Control Conference, 9-11 Aug. 1999, Portland, OR, USA Contract(s)/Grant(s): NCC4-117; RTOP 242-33-02-00-23 Report No.(s): H-2433; AIAA Paper 99-4134; Copyright; Avail: CASI; A03, Hardcopy Two methods for control system reconfiguration have been investigated. The first method is a robust servomechanism control approach (optimal tracking problem) that is a generalization of the classical proportional-plus-integral control to multiple input-multiple output systems. The second method is a control-allocation approach based on a quadratic programming formulation. A globally convergent fixed-point iteration algorithm has been developed to make onboard implementation of this method feasible. These methods have been applied to reconfigurable entry flight control design for the X-33 vehicle. Examples presented demonstrate simultaneous tracking of angle-of-attack and roll angle commands during failures of the fight body flap actuator. Although simulations demonstrate success of the first method in most cases, the control-allocation method appears to provide uniformly better performance in all cases. Author Flight Control; Servomechanisms; Proportional Control; Mimo (Control Systems); Quadratic Programming; X-33 Reusable Launch Vehicle

20010060161 Old Dominion Univ., Norfolk, VA USA N-Factor Computations for the X-33 Vehicle Balakumar, P.; Hamilton, H. Harris, Technical Monitor; May 24, 2001; In English Contract(s)/Grant(s): NAG1-2297 Report No.(s): ODURF-103331; No Copyright; Avail: CASI; A05, Hardcopy N-Factor computations for the X-33 vehicle on the lower side of the body were per-formed. The transition prediction code e-Malik was used to compute the N-Factors and the meanflow field was computed using the CFD code LAURA. The computations are done at two angles of attack 30 and 40 degrees and the freestream Mach number and Reynolds number are M= 6 and Re = 7.7* 10(exp 6). The N-Factors obtained are very low in the range of 2-3 near the experimentally observed transition onset regions. The stability computations show that at 30 degrees angle of attack, the transition occurs due to crossflow instability and at higher angles of attack the transition occurs due to first mode inviscid instability. Author Angle of Attack; Computational Fluid Dynamics; Inviscid Flow; Stability; X-33 Reusable Launch Vehicle; Boundary Layers

20010059347 NASA Marshall Space Flight Center, Huntsville, AL USA Progress and Status of the 2nd Generation Reusable Launch Vehicle Program Dumbacher, Daniel L.; Rogacki, John R., Technical Monitor; [2001]; In English; 52nd International Astronautical Congress, 1-5 Oct. 2001, Toulouse, France; No Copyright; Avail: Other Sources; Abstract Only Safe, low-cost space transportation is the key enabler of the commercial development and civil exploration of space. Human space ﬂight remains a hazardous endeavor in spite of advances in aerospace technology. NASA intends to reduce the resources devoted to routine space operations, thus enabling additional science research, technology development, and exploration activities. NASA developed the Integrated Space Transportation Plan (ISTP) to address the needs for space access. ISTP is a balanced and comprehensive plan that embodies strategies for achieving goals for earth-to-orbit and in-space transportation operations. ISTP calls for near-term safety related upgrades to the Space Shuttle, development of a Crew Return Vehicle, a new Space Launch Initiative (SLI), and development of an Aero-Space Base to provide foundation technology for

76 future space systems. NASA formed the SLI to coordinate the development of the 2nd Generation Reusable Launch Vehicle (RLV) architecture, The SLI is implemented by the 2nd Generation RLV Program Office. Author Reusable Launch Vehicles; Spacecraft Launching; Aerospace Systems; Space Transportation; Low Cost; Space Shuttles

20010054791 Georgia Inst. of Tech., Atlanta, GA USA A Method of Integrating Aeroheating into Conceptual Reusable Launch Vehicle Design: Evaluation of Advanced Thermal Protection Techniques for Future Reusable Launch Vehicles Olds, John R.; Cowart, Kris; June 2001; In English Contract(s)/Grant(s): NCC2-5332; No Copyright; Avail: CASI; A09, Hardcopy A method for integrating Aeroheating analysis into conceptual reusable launch vehicle (RLV) design is presented in this thesis. This process allows for faster turn-around time to converge a RLV design through the advent of designing an optimized thermal protection system (TPS). It consists of the coupling and automation of four computer software packages: MINIVER, TPSX, TCAT, and ADS. MINIVER is an Aeroheating code that produces centerline radiation equilibrium temperatures, convective heating rates, and heat loads over simplified vehicle geometries. These include flat plates and swept cylinders that model wings and leading edges, respectively. TPSX is a NASA Ames material properties database that is available on the World Wide Web. The newly developed Thermal Calculation Analysis Tool (TCAT) uses finite difference methods to carry out a transient in-depth 1-D conduction analysis over the center mold line of the vehicle. This is used along with the Automated Design Synthesis (ADS) code to correctly size the vehicle’s thermal protection system (TPS). The numerical optimizer ADS uses algorithms that solve constrained and unconstrained design problems. The resulting outputs for this process are TPS material types, unit thicknesses, and acreage percentages. TCAT was developed for several purposes. First, it provides a means to calculate the transient in-depth conduction seen by the surface of the TPS material that protects a vehicle during ascent and reentry. Along with the in-depth conduction, radiation from the surface of the material is calculated along with the temperatures at the backface and interior parts of the TPS material. Secondly, TCAT contributes added speed and automation to the overall design process. Another motivation in the development of TCAT is optimization. In some vehicles, the TPS accounts for a high percentage of the overall vehicle dry weight. Optimizing the weight of the TPS will thereby lower the percentage of the dry weight accounted for by the TPS. Also, this will lower the cost of the TPS and the overall cost of the vehicle. Derived from text Aerodynamic Heating; Computer Programs; Thermal Protection

20010050134 Old Dominion Univ., Norfolk, VA USA Panel Flutter and Sonic Fatigue Analysis for RLV Mei, Chuh; Cheng, Guangfeng; May 2001; In English Contract(s)/Grant(s): NAG1-2150; ODURF Proj. 191881; No Copyright; Avail: CASI; A03, Hardcopy A methodology is presented for the flutter analysis of the seal of thermal protection system (TPS) panel of X-33 Advanced Technology Demonstrator test vehicle. The seal is simulated as a two-dimensional cantilevered panel with an elastic stopper, which is modeled as an equivalent spring. This cantilever beam-spring model under the aerodynamic pressure at supersonic speeds turns out to be an impact nonlinear dynamic system. The flutter analysis of the seal is thus carried out using, time domain numerical simulation with a displacement stability criterion. The flutter boundary of the seal is further verified with a family of three traditional and one nontraditional panel flutter models. The frequency domain method that applies eigenanalysis on the traditional panel flutter problem was used. The results showed that the critical dynamic pressure could be more than doubled with properly chosen material for the base stopper. The proposed methodology can be easily extended to three-dimensional panel seals with flow angularity. Author Acoustic Fatigue; Boundaries; Displacement; Flutter Analysis; Panel Flutter; Supersonic Speed; Thermal Protection

20010048747 NASA Marshall Space Flight Center, Huntsville, AL USA An Approach to Establishing System Beneﬁts for Technologies In NASA’s Spaceliner Investment Area Hueter, Uwe; Pannell, Bill; Lyles, Garry M., Technical Monitor; [2001]; In English; 52nd International Aeronautical Congress, 1-5 Oct. 2001, Toulouse, France; No Copyright; Avail: Other Sources; Abstract Only NASA’s has established long term goals for access-to-space. The third generation launch systems are to be fully reusable and operational around 2025. The goals for the third generation launch system are to signiﬁcantly reduce cost and improve

77 safety over current systems. The Advanced Space Transportation Program Office (ASTP) at the NASA’s Marshall Space Flight Center in Huntsville, AL has the agency lead to develop space transportation technologies. Within ASTP, under the Spaceliner Investment Area, third generation technologies are being pursued. The Spaceliner Investment Area’s primary objective is to mature vehicle technologies to enable substantial increases in the design and operating margins of third generation RLVs (current Space Shuttle is considered the first generation RLV) by incorporating advanced propulsion systems, materials, structures, thermal protection systems, power, and avionics technologies. Advancements in design tools and better characterization of the operational environment will result in reduced design and operational variabilities leading to improvements in margins. Improvements in operational efficiencies will be obtained through the introduction of integrated vehicle health management, operations and range technologies. Investments in these technologies will enable the reduction in the high operational costs associated with today’s vehicles by allowing components to operate well below their design points resulting in improved component operating life, reliability, and safety which in turn reduces both maintenance and refurbishment costs. The introduction of advanced technologies may enable horizontal takeoff by significantly reducing the takeoff weight and allowing use of existing infrastructure. This would be a major step toward the goal of airline-like operation. These factors in conjunction with increased flight rates, resulting from reductions in transportation costs, will result in significant improvements of future vehicles. The real-world problem is that resources are limited and technologies need to be prioritized to assure the resources are spent on technologies that provide the highest system level benefits. Toward that end, a systems approach is being taken to determine the benefits of technologies for the Spaceliner Investment Area. Technologies identified to be enabling will be funded. However, the other technologies will be funded based on their system’s benefits. Since the final launch system concept will not be decided for many years, several vehicle concepts are being evaluated to determine technology benefits. Not only performance, but also cost and operability are being assessed. This will become an annual process to assess these technologies against their goals and the benefits to various launch systems concepts. The paper describes the system process, tools and concepts used to determine the technology benefits. Preliminary results will be presented along with the current technology investments that are being made by ASTP’s Spaceliner Investment Area. Author Cost Reduction; Operating Costs; Space Transportation System; Reusable Launch Vehicles; Reusable Spacecraft; Cost Effectiveness; Cost Analysis

20010046963 NASA Marshall Space Flight Center, Huntsville, AL USA An Overview of In-Stu Treatability Studies at Marshall Space Flight Center, Huntsville, Alabama McElroy, Bill; Keith, Amy; Glasgow, J. K.; Dasappa, Srini; McCaleb, Rebecca, Technical Monitor; [2001]; In English; Oxidation and Reduction Technologies for In-Situ Treatment of Soil and Groundwater, 26-29 Jun. 2001, Niagra Falls, Canada; No Copyright; Avail: Other Sources; Abstract Only Marshall Space Flight Center (MSFC) is located in Huntsville, Alabama (north-central Alabama), on approximately 1,840 acres near the center of the U.S. Army’s Redstone Arsenal (RSA). MSFC is the National Aeronautics and Space Administration’s (NASA’s) principal propulsion development center. Its scientists, engineers, and support personnel play a major role in the National Space Transportation System by managing space shuttle mission activities, including the microgravity laboratory. In addition, MSFC will be a significant contributor to several of NASA’s future programs, including the Reusable Launch Vehicle (X-33), International Space Station, and Advanced X-ray Astrophysics Facility, as well as research on a variety of space science applications. MSFC has been used to develop, test and manufacture space vehicles and components since 1960, when civilian rocketry and missile activities were transferred from RSA to MSFC. In 1994, MSFC was placed on the National Priority List for the management of hazardous waste sites, under the requirements of the Comprehensive Environmental Response, Compensation, and Liability Act (CERCLA). One requirement of the CERCLA program is to evaluate the nature and extent of environmental contamination resulting from identified CERCLA sites, assess the public health and environmental risks associated with the identified contamination, and identify potential remedial actions. A CERCLA remedial investigation (RI) for the groundwater system has identified at least five major plumes of chlorinated volatile organic compounds (CVOCs) in the groundwater beneath the facility. These plumes are believed to be the result of former management practices at 14 main facility locations (termed ‘source areas’) where CVOCs were released to the subsurface. Trichloroethene (TCE) is the predominant CVOC and is common to all the plumes. Perchloroethene (PCE) also exists in two of the plumes. In addition to TCE and PCE, carbon tetrachloride and 1,1,2,2-tetrachloroethane are contained in one of the plumes. The CVOCs are believed to exist as dense non-aqueous phase liquids (DNAPLs) beneath many of the source areas. Author NASA Programs; Environment Pollution

78 20010046959 NASA Marshall Space Flight Center, Huntsville, AL USA X-33 XRS-2200 Linear Aerospike Engine Sea Level Plume Radiation DAgostino, Mark; Lee, Young C.; Wang, Ten-See; Turner, Jim E., Technical Monitor; [2001]; In English; Fluids Workshop, 5 Apr. 2001, Huntsville, AL, USA; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy This paper presents radiation data for the linear aerospike engine. CASI Aerospike Engines; Plumes

20010046955 NASA Marshall Space Flight Center, Huntsville, AL USA 2nd Generation Reusable Launch Vehicle Potential Commercial Development Scenarios Creech, Stephen D.; Rogacki, John R., Technical Monitor; [2001]; In English; IAA 31st Symposium of Economics and Commercialization of Space Activities; No Copyright; Avail: Other Sources; Abstract Only The presentation will discuss potential commercial development scenarios for a Second Generation Reusable Launch Vehicle. The analysis of potential scenarios will include commercial rates of return, government return on investment, and market considerations. The presentation will include policy considerations in addition to analysis of Second Generation Reusable Launch Vehicle economics. The data discussed is being developed as a part of NASA’s Second Generation Reusable Launch Vehicle Program, for consideration as potential scenarios for enabling a next generation system. Material will include potential scenarios not previously considered by NASA or presented at other conferences. Candidate paper has not been presented at a previous meeting, and conference attendance of the author has been approved by NASA. Author Reusable Launch Vehicles; Economics; Commercialization; Conferences

20010045816 NASA Marshall Space Flight Center, Huntsville, AL USA 2nd Generation RLV: Program Goals and Acquisition Strategy Graham, J. Bart; Dumbacher, D. L., Technical Monitor; [2001]; In English; Space Technology and Applications International Forum, 11-14 Feb. 2001, Albuquerque, NM, USA; No Copyright; Avail: CASI; A02, Hardcopy The risk to loss of life for Space Shuttle crewmembers is approximately one in 245 missions. U.S. launch service providers captured nearly 100%, of the commercial launch market revenues in the mid 1980s. Today, the U.S. captures less than 50% of that market. A launch system architecture is needed that will dramatically increase the safety of space flight while significantly reducing the cost. NASA’s Space Launch Initiative, which is implemented by the 2nd Generation RLV Program Office at Marshall Space Flight Center, seeks to develop technology and reusable launch vehicle concepts which satisfy the commercial launch market needs and the unique needs of NASA. Presented in this paper are the five primary elements of NASA’s Integrated Space Transportation Plan along with the highest level goals and the acquisition strategy of the 2nd Generation RLV Program. Approval of the Space Launch Initiative FY01 budget of $290M is seen as a major commitment by the Agency and the Nation to realize the commercial potential that space offers and to move forward in the exploration of space. Author Space Transportation; Reusable Launch Vehicles; Market Research

20010043670 NASA Langley Research Center, Hampton, VA USA Detection of Micro-Leaks Through Complex Geometries Under Mechanical Load and at Cryogenic Temperature Rivers, H. Kevin; Sikora, J. G.; Sankaran, S. N.; [2001]; In English; 42nd Structures, Structural Dynamics, and Materials Conference and Exhibit, 16-19 Apr. 2001, Seattle, WA, USA Report No.(s): AIAA Paper 2001-1218; Copyright; Avail: CASI; A01, Hardcopy Polymer Matrix Composite (PMC) hydrogen tanks have been proposed as an enabling technology for reducing the weight of Single-Stage-to-Orbit reusable launch vehicles where structural mass has a large impact on vehicle performance. A key development issue of these lightweight structures is the leakage of hydrogen through the composite material. The rate of hydrogen leakage can be a function of the material used, method of 6 fabrication used to manufacture the tank, mechanical load the tank must react, internal damage-state of the material, and the temperatures at which the tank must operate. A method for measuring leakage through a geometrically complex structure at cryogenic temperature and under mechanical load was developed, calibrated and used to measure hydrogen leakage through complex X-33 liquid-hydrogen tank structure sections. Author Polymer Matrix Composites; Leakage; Propellant Tanks; Cryogenic Temperature; Liquid Hydrogen; Fault Detection; Composite Structures

79 20010041325 NASA Marshall Space Flight Center, Huntsville, AL USA Development of Aluminum Composites for a Rocket Engine’s Lightweight Thrust Cell Lee, Jonathan A.; Elam, Sandy; Munafo, Paul M., Technical Monitor; [2001]; In English; 25th Annual Conference on Composites, Materials and Structures, 21-26 Jan. 2001, Cocoa Beach, FL, USA Contract(s)/Grant(s): RTOP 713-23-53; No Copyright; Avail: Other Sources; Abstract Only The Aerospike liquid fueled rocket engine for the X-33 aerospace vehicle consists of several thrust cells, which can comprise as much as 25% of the engine weight. The interior wall of the thrust cell chamber is exposed to high temperature combustion products and must be cooled by using liquid hydrogen. Ultimately, reducing engine weight would increase vehicle performance and allow heavier payload capabilities. Currently, the thrust cell’s structural jacket and manifolds are made of stainless steel 347, which can potentially be replaced by a lighter material such as an Aluminum (Al) Metal Matrix Composites (MMC). Up to 50% weight reduction can be expected for each of the thrust cell chambers using particulate SiC reinforced Al MMC. Currently, several Al MMC thrust cell structural jackets have been produced, using cost-effective processes such as gravity casting and plasma spray deposition, to demonstrate MMC technology readiness for NASA’s advanced propulsion systems. Author Aerospike Engines; Aluminum; Metal Matrix Composites; Thrust Chambers; Weight Reduction

20010041080 NASA Ames Research Center, Moffett Field, CA USA A Two-Stage-to-Orbit Spaceplane Concept With Growth Potential Mehta, Unmeel B.; Bowles, Jeffrey V.; February 2001; In English; 10 International Space Planes and Hypersonic Systems and Technologies Conference, 24-27 Apr. 2001, Kyoto, Japan; Original contains color illustrations Contract(s)/Grant(s): RTOP 713-74-00 Report No.(s): NASA/TM-2001-209620; NAS 1.15:209620; A-00V0029; No Copyright; Avail: CASI; A03, Hardcopy A two-stage-to-orbit (TSTO) spaceplane concept developed in 1993 is revisited, and new information is provided to assist in the development of the next-generation space transportation vehicles. The design philosophy, TSTO spaceplane concept, and the design method are briefly described. A trade study between cold and hot structures leads to the choice of cold structures with external thermal protection systems. The optimal Mach number for staging the second stage of the TSTO spaceplane (with air-breathing propulsion on the first stage) is 10, based on life-cycle cost analysis. The performance and specification of a prototype/experimental (P/X) TSTO spaceplane with a turbo/ram/scramjet propulsion system and built-in growth potential are presented and discussed. The internal rate of return on investment is the highest for the proposed TSTO spaceplane, vis-A-vis a single-stage-to-orbit (SSTO) rocket vehicle and a TSTO spaceplane without built-in growth. Additional growth potentials for the proposed spaceplane are suggested. This spaceplane can substantially decrease access-to-space cost and risk, and increase safety and reliability in the near term It can be a serious candidate for the next-generation space transportation system. Author Single Stage to Orbit Vehicles; Space Transportation System; Air Breathing Engines; Aerospace Planes; Design Analysis

20010039539 NASA Marshall Space Flight Center, Huntsville, AL USA Permeability Testing of Impacted Composite Laminates for Use on Reusable Launch Vehicles Nettles, Alan T.; Munafo, Paul, Technical Monitor; February 2001; In English Report No.(s): NASA/TM-2001-210799; M-1001; NAS 1.15:210799; No Copyright; Avail: Other Sources; Abstract Only Since composite laminates are beginning to be identified for use in reusable launch vehicle propulsion systems, an understanding of their permeance is needed. A foreign object impact event can cause a localized area of permeability (leakage) in a polymer matrix composite, and it is the aim of this study to assess a method of quantifying permeability-after-impact results. A simple test apparatus is presented, and variables that could affect the measured values of permeability-after-impact were assessed. Once it was determined that valid numbers were being measured, a fiber/resin system was impacted at various impact levels and the resulting permeability measured, first with a leak check solution (qualitative) then using the new apparatus (quantitative). The results showed that as the impact level increased, so did the measured leakage. As the pressure to the specimen was increased, the leak rate was seen to increase in a nonlinear fashion for almost all the specimens tested. Author Laminates; Permeability; Polymer Matrix Composites; Impact Tests

80 20010039031 Boeing Co., Huntington Beach, CA USA, NASA Marshall Space Flight Center, Huntsville, AL USA A Cryogenic Propellant Production Depot for Low Earth Orbit Potter, Seth D.; Henley, Mark; Guitierrez, Sonia; Fikes, John; Carrington, Connie; Smitherman, David; Gerry, Mark; Sutherlin, Steve; Beason, Phil; Howell, Joe, Technical Monitor; [2001]; In English; International Space Development Conference, 24-28 May 2001, Albuquerque, NM, USA; No Copyright; Avail: CASI; A03, Hardcopy The cost of access to space beyond low Earth orbit can be lowered if vehicles can refuel in orbit. The power requirements for a propellant depot that electrolyzes water and stores cryogenic oxygen and hydrogen can be met using technology developed for space solar power. A propellant depot is described that will be deployed in a 400 km circular equatorial orbit, receive tanks of water launched into a lower orbit from Earth by gun launch or reusable launch vehicle, convert the water to liquid hydrogen and oxygen, and store Lip to 500 metric tonnes of cryogenic propellants. The propellant stored in the depot can support transportation from low Earth orbit to geostationary Earth orbit, the Moon, LaGrange points, Mars, etc. The tanks are conﬁgured in an inline gravity-gradient conﬁguration to minimize drag and settle the propellant. Temperatures can be maintained by body-mounted radiators; these will also provide some shielding against orbital debris. Power is supplied by a pair of solar arrays mounted perpendicular to the orbital plane, which rotate once per orbit to track the Sun. In the longer term, cryogenic propellant production technology can be applied to a larger LEO depot, as well as to the use of lunar water resources at a similar depot elsewhere. Author Cryogenic Rocket Propellants; Low Earth Orbits; Refueling

20010038424 NASA Marshall Space Flight Center, Huntsville, AL USA Unshrouded Impeller Technology Task Status Williams, Robert W.; Turner, James E., Technical Monitor; [2001]; In English; TD Fluids Workshop, 4-5 Apr. 2001, Huntsville, AL, USA; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy The objective of this presentation is to develop an unshrouded impeller design that will meet the performance requirements of a 3-stage fuel pump with a 2-stage pump design. CASI Shrouded Propellers; Impellers; Fuel Pumps; Turbine Pumps

20010038264 NASA Marshall Space Flight Center, Huntsville, AL USA NASA’s Spaceliner Investment Area Technology Activities Hueter, Uwe; Lyles, Garry M., Technical Monitor; [2001]; In English; 10th International Space Planes and Hypersonics Systems and Technologies Conference, 24-27 Apr. 2001, Kyoto, Japan; Copyright; Avail: CASI; A02, Hardcopy NASA’s has established long term goals for access-to-space. The third generation launch systems are to be fully reusable and operational around 2025. The goals for the third generation launch system are to significantly reduce cost and improve safety over current conditions. The Advanced Space Transportation Program Office (ASTP) at the NASA’s Marshall Space Flight Center in Huntsville, AL has the agency lead to develop space transportation technologies. Within ASTP, under the Spaceliner Investment Area, third generation technologies are being pursued in the areas of propulsion, airframes, integrated vehicle health management (IVHM), avionics, power, operations, and range. The ASTP program will mature these technologies through both ground and flight system testing. The Spaceliner Investment Area plans to mature vehicle technologies to reduce the implementation risks for future commercially developed reusable launch vehicles (RLV). The plan is to substantially increase the design and operating margins of the third generation RLV (the Space Shuttle is the first generation) by incorporating advanced technologies in propulsion, materials, structures, thermal protection systems, avionics, and power. Advancements in design tools and better characterization of the operational environment will allow improvements in design margins. Improvements in operational efficiencies will be provided through use of advanced integrated health management, operations, and range technologies. The increase in margins will allow components to operate well below their design points resulting in improved component operating life, reliability, and safety which in turn reduces both maintenance and refurbishment costs. These technologies have the potential of enabling horizontal takeoff by reducing the takeoff weight and achieving the goal of airline-like operation. These factors in conjunction with increased flight rates from an expanding market will result in significant improvements in safety and reductions in operational costs of future vehicles. The paper describes current status, future plans and technologies that are being matured by the Spaceliner Investment Area under the Advanced Space Transportation Program Office. Author Space Transportation; Technology Utilization; NASA Programs; Avionics; Systems Analysis; Spacecraft Propulsion

81 20010037825 Engineering Sciences, Inc., Huntsville, AL USA Unic Unstructured CFD Methodology Development Chen, Y. S.; Liu, J.; Zhang, S.; Mallapragada, P.; Turner, James E., Technical Monitor; Wang, T. S., Technical Monitor; D’Agostino, M., Technical Monitor; [2001]; In English; Fluids, 4-5 Apr. 2001, Huntsville, AL, USA; Original contains color illustrations Contract(s)/Grant(s): NAS8-00002; No Copyright; Avail: CASI; A03, Hardcopy Base heating characteristics is crucial to the success and the overall performance of the X-33 engine. Base heating is important throughout the entire flight trajectory due to the aerospike engine design of X-33. The base region is surrounded by the hot-gas plume, which expands, circulates and impinges on the base. An advanced computation fluid dynamics method is employed in an effort to develop a robust, accurate and efficient tool for the X-33 base heating performance predictions. This computational tool is developed based on a Navier-Stokes flow solver, which is suitable for general complex geometry and includes turbulence, finite-rate chemistry, and radiation models. To fulfill the fast turnaround requirement as a design analysis tool, adaptive mesh refinement method and parallel-computing algorithm are also incorporated in the present model. Case study for the X-33 base-region fluid dynamics and heat transfer characteristics are presented. Author Computational Fluid Dynamics; X-33 Reusable Launch Vehicle; Base Heating; Engine Design; Heat Transfer; Design Analysis

20010028815 NASA Langley Research Center, Hampton, VA USA Evaluation of Computational Aeroelasticity Codes for Loads and Flutter Huttsell, Larry; Schuster, Dave; Volk, John; Giesing, Joe; Love, Mike; [2001]; In English Report No.(s): AIAA Paper 01-0569; Copyright; Avail: CASI; A03, Hardcopy This paper presents an overview of a joint effort to evaluate computational aeroelasticity codes for loads and flutter. Computations are performed on realistic problems for static aeroelasticity, classic flutter, Limit Cycle Oscillations (LCO), and control surface buzz. The codes being evaluated involve a transonic small disturbance code with an interactive boundary layer method and two Euler/Navier-Stokes codes. To evaluate the accuracy of these codes, comparisons are made with available wind tunnel or flight test data. This paper presents the results/status of the following applications: static aeroelasticity - aeroelastic tailored wing, flutter - AV-8B-like wing and F-I5-like tail, LCO - B-1-like and B-2, and control surface buzz - NASP wind tunnel model and Global Hawk wing. Author Aeroelasticity; Flutter; Loads (Forces); Control Surfaces; Wing Oscillations

20010025824 Pratt and Whitney Aircraft, West Palm Beach, FL USA 2nd Generation Reusable Launch Vehicle (2G RLV) Matlock, Steve; Sides, Steve; Kmiec, Tom; Arbogast, Tim; Mayers, Tom; Doehnert, Bill; Feb. 23, 2001; In English; Original contains color illustrations Contract(s)/Grant(s): NAS8-00168; NRA8-27 Report No.(s): LS-50-Rev; No Copyright; Avail: CASI; A06, Hardcopy This is a revised final report and addresses all of the work performed on this program. Specifically, it covers vehicle architecture background, definition of six baseline engine cycles, reliability baseline (space shuttle main engine QRAS), and component level reliability/performance/cost for the six baseline cycles, and selection of 3 cycles for further study. This report further addresses technology improvement selection and component level reliability/performance/cost for the three cycles selected for further study, as well as risk reduction plans, and recommendation for future studies. Author Reusable Launch Vehicles; Rocket Engine Design; Liquid Rocket Propellants; Space Shuttle Main Engine; Performance Prediction

20010021342 NASA Marshall Space Flight Center, Huntsville, AL USA Lightweight Chambers for Thrust Assemblies Elam, Sandra K.; Lee, Jonathan; Holmes, Richard; Zimmerman, Frank; Effinger, Mike; Turner, James E., Technical Monitor; [2001]; In English; 52nd International Astronautical Congress, 1-5 Oct. 2001, Toulouse, France Contract(s)/Grant(s): NRA8-21; No Copyright; Avail: Other Sources; Abstract Only The Marshall Space Flight Center (MSFC) of the National Aeronautics and Space Administration (NASA) has

82 successfully applied new materials and fabrication techniques to create actively cooled thrust chambers that operate 200-400 degrees hotter and weigh 50% lighter than conventional designs. In some vehicles, thrust assemblies account for as much as 20% of the engine weight. So, reducing the weight of these components and increasing their operating range will benefit many engines and vehicle designs, including Reusable Launch Vehicle (RLV) concepts. Obviously, copper and steel alloys have been used successfully for many years in the chamber components of thrust assemblies. Yet, by replacing the steel alloys with Polymer Matrix Composite (PMC) and/or Metal Matrix Composite (MMC) materials, design weights can be drastically reduced. In addition, replacing the traditional copper alloys with a Ceramic Matrix Composite (CMC) or an advanced copper alloy (Cu-8Cr-4Nb, also known as GRCop-84) significantly increases allowable operating temperatures. Several small MMC and PMC demonstration chambers have recently been fabricated with promising results. Each of these designs included GRCop-84 for the cooled chamber liner. These units successfully verified that designs over 50% lighter are feasible. New fabrication processes, including advanced casting technology and a low cost vacuum plasma spray (VPS) process, were also demonstrated with these units. Hot-fire testing at MSFC is currently being conducted on the chambers to verify increased operating temperatures available with the GRCop-84 liner. Unique CMC chamber liners were also successfully fabricated and prepared for hot-fire testing. Yet, early results indicate these CMC liners need significantly more development in order to use them in required chamber designs. Based on the successful efforts with the MMC and PMC concepts, two full size ‘lightweight’ chambers are currently being designed and fabricated for hot-fire testing at MSFC in 2001. These ‘full size’ chambers will be similar in size to those used on the X33 engine (RS2200). One will be fabricated with a MMC structural jacket, while the other uses a PMC jacket. Each will be designed for thrust levels of 15,000 pounds in an oxygen/hydrogen environment with liquid hydrogen coolant. Both chambers will use GRCop-84 for its channel wall liner. Each unit is expected to be at least 60% lighter than a conventional design with traditional materials. Hot-fire testing on the full size units in late 2001 will directly compare performance results between a conventional chamber design and these ‘lightweight’ alternatives. The technology developed and demonstrated by this effort will not only benefit next generation RLV programs, but it can be applied to other existing and future engine programs, as well. Efforts were sponsored by the Advanced Space Transportation Program for RLV Focused Technologies. The task team was led by MSFC with additional members from NASA-Glenn Research Center and the Rocketdyne Division of The Boeing Company. Specific materials development and fabrication processes were provided by Aerojet, Lockheed Martin Astronautics, Composite Optics, Inc., Hyper-Therm, Ceramic Composites, Inc., MSE Technology Applications, and Plasma Processes, Inc. Author Technology Assessment; Fabrication; Ceramic Matrix Composites; Copper Alloys; Metal Matrix Composites; Polymer Matrix Composites; Technology Utilization; Thrust Chambers

20010020948 NASA Marshall Space Flight Center, Huntsville, AL USA Permeability Testing of Impacted Composite Laminates for Use on Reusable Launch Vehicles Nettles, A. T.; February 2001; In English Report No.(s): NASA/TM-2001-210799; M-1001; NAS 1.15:210799; No Copyright; Avail: CASI; A03, Hardcopy Since composite laminates are beginning to be identified for use in reusable launch vehicle propulsion systems, an understanding of their permeance is needed. A foreign object impact event can cause a localized area of permeability (leakage) in a polymer matrix composite, and it is the aim of this study to assess a method of quantifying permeability-after-impact results. A simple test apparatus is presented, and variables that could affect the measured values of permeability-after-impact were assessed. Once it was determined that valid numbers were being measured, a fiber/resin system was impacted at various impact levels and the resulting permeability measured, first with a leak check solution (qualitative) then using the new apparatus (quantitative). The results showed that as the impact level increased, so did the measured leakage. As the pressure to the specimen was increased, the leak rate was seen to increase in a nonlinear fashion for almost all the specimens tested. Author Laminates; Leakage; Permeability; Microcracks; Impact Tests

20010020386 NASA Marshall Space Flight Center, Huntsville, AL USA RLV Turbine Performance Optimization Griffin, Lisa W.; Dorney, Daniel J.; [2000]; In English; 12th PERC Propulsion Symposium, 26-27 Oct. 2000, Cleveland, OH, USA; Original contains color illustrations Contract(s)/Grant(s): NRA8-21; No Copyright; Avail: CASI; A02, Hardcopy A task was developed at NASA/Marshall Space Flight Center (MSFC) to improve turbine aerodynamic performance through the application of advanced design and analysis tools. There are four major objectives of this task: 1) to develop, enhance, and integrate advanced turbine aerodynamic design and analysis tools; 2) to develop the methodology for application

83 of the analytical techniques; 3) to demonstrate the beneﬁts of the advanced turbine design procedure through its application to a relevant turbine design point; and 4) to verify the optimized design and analysis with testing. Final results of the preliminary design and the results of the two-dimensional (2D) detailed design of the ﬁrst-stage vane of a supersonic turbine suitable for a reusable launch vehicle (R-LV) are presented. Analytical techniques for obtaining the results are also discussed. Author Reusable Launch Vehicles; Computational Fluid Dynamics; Gas Turbine Engines; Aerodynamic Characteristics; Design Optimization

20010019998 NASA Marshall Space Flight Center, Huntsville, AL USA Testing the Preliminary X-33 Navigation System Lomas, James J.; Mitchell, Daniel W.; Freestone, Todd M.; Lee, Charles; Lessman, Craig; Foster, Lee D., Technical Monitor; [2001]; In English; 24th Guidance and Control, 31 Jan. - 4 Feb. 2001, Breckenridge, CO, USA; No Copyright; Avail: Other Sources; Abstract Only The X-33 Reusable Launch Vehicle (RLV) must meet the demanding requirements of landing autonomously on a narrow landing strip following a flight that reaches an altitude of up to 200,000 feet and a speed in excess of Mach 9 with significant in-flight energy bleed-off maneuvers. To execute this flight regimen a highly reliable avionics system has been designed that includes three LN-100G Inertial Navigation System/Global Positioning System (INS/GPS) units as the primary navigation system for the X-33. NASA’s Marshall Space Flight Center (MSFC) tested an INS/GPS system in real-time simulations to determine the ability of this navigation suite to meet the in flight and autonomous landing requirements of the X-33 RLV. A total of sixty-one open loop tests were performed to characterize the navigation accuracy of the LN-100G. Twenty-seven closed-loop tests were also performed to evaluate the performance of the X-33 Guidance, Navigation and Control (GN&C) algorithms with the real navigation hardware. These closed-loop tests were also designed to expose any integration or operational issues with the real-time X-33 vehicle simulation. Dynamic road tests of the INS/GPS were conducted by Litton to assess the performance of differential and nondifferential INS/GPS hybrid navigation solutions. The results of the simulations and road testing demonstrate that this novel solution is capable of meeting the demanding requirements of take-off, in-flight navigation, and autonomous landing of the X-33 RLV. This paper describes the test environment developed to stimulate the LN-100G and discusses the results of this test effort. This paper also presents recommendations for a navigation system suitable to an operational RLV system. Author X-33 Reusable Launch Vehicle; Autonomy; Spacecraft Landing; Navigation Aids; Performance Tests

20010018568 NASA Langley Research Center, Hampton, VA USA Infrared Sensing Aeroheating Flight Experiement: STS-96 Flight Results Blanchard, Robert C.; Wilmoth, Richard G.; Glass, Christopher E.; Merski, N. Ronald, Jr.; Berry, Scott A.; Bozung, Timothy J.; Tietjen, Alan; Wendt, Jodean; Dawson, Don; [2001]; In English; 39th Aerospace Sciences, 8-11 Jan. 2001, Reno, NV, USA; Original contains color illustrations Report No.(s): AIAA Paper 2001-0352; Copyright; Avail: CASI; A03, Hardcopy Major elements of an experiment called the Infrared Sensing Aeroheating Flight Experiment are discussed. The primary experiment goal is to provide reentry global temperature images from infrared measurements to define the characteristics of hypersonic boundary-layer transition during flight. Specifically, the experiment is to identify, monitor, and quantify hypersonic boundary layer windward surface transition of the X-33 vehicle during flight. In addition, the flight data will serve as a calibration and validation of current boundary layer transition prediction techniques, provide benchmark laminar, transitional, and fully turbulent global aeroheating data in order to validate existing wind tunnel and computational results, and to advance aeroheating technology. Shuttle Orbiter data from STS-96 used to validate the data acquisition and data reduction to global temperatures, in order to mitigate the experiment risks prior to the maiden flight of the X-33, is discussed. STS-96 reentry mid-wave (3-5 Pm) infrared data were collected at the Ballistic Missile Defense Organization/Innovative Sciences and Technology Experimentation Facility site at NASA-Kennedy Space Center and subsequently mapped into global temperature contours using ground calibrations only. A series of image mapping techniques have been developed in order to compare each frame of infrared data with thermocouple data collected during the flight. Comparisons of the ground calibrated global temperature images with the corresponding thermocouple data are discussed. The differences are shown to be generally less than about 5%, which is comparable to the expected accuracy of both types of aeroheating measurements. Author Infrared Radiation; Performance Tests; Heating; Reentry Physics; Temperature Measurement; Surface Temperature; Data Acquisition; Flight Characteristics

84 20010017162 NASA Glenn Research Center, Cleveland, OH USA Preliminary Sizing of Vertical Take-Off Rocket-Based Combined-Cycle Powered Launch Vehicles Roche, Joseph M.; McCurdy, David R.; January 2001; In English; 37th Combustion, 13-17 Nov. 2000, Monterey, CA, USA Contract(s)/Grant(s): RTOP 708-73-20 Report No.(s): NASA/TM-2001-210668; E-12503-1; NAS 1.15:210668; No Copyright; Avail: CASI; A03, Hardcopy The task of single-stage-to-orbit has been an elusive goal due to propulsion performance, materials limitations, and complex system integration. Glenn Research Center has begun to assemble a suite of relationships that tie Rocket-Based Combined-Cycle (RBCC) performance and advanced material data into a database for the purpose of preliminary sizing of RBCC-powered launch vehicles. To accomplish this, a near optimum aerodynamic and structural shape was established as a baseline. The program synthesizes a vehicle to meet the mission requirements, tabulates the results, and plots the derived shape. A discussion of the program architecture and an example application is discussed herein. Author Single Stage to Orbit Vehicles; Mission Planning; Aerodynamics

20010016865 NASA Glenn Research Center, Cleveland, OH USA Structural Sizing of a 25,000-lb Payload, Air-Breathing Launch Vehicle for Single-Stage-to-Orbit Roche, Joseph M.; Kosareo, Daniel N.; January 2001; In English; 37th Combustion, 13-17 Nov. 2000, Monterey, CA, USA Contract(s)/Grant(s): RTOP 708-73-20 Report No.(s): NASA/TM-2001-210667; E-12501-1; NAS 1.15:210667; No Copyright; Avail: CASI; A03, Hardcopy In support of NASA’s Air-Breathing Launch Vehicle (ABLV) study, a 25,000-lb payload version of the GTX (formerly Trailblazer) reference vehicle concept was developed. The GTX is a vertical lift-off, reusable, single-stage-to-orbit launch vehicle concept that uses hypersonic air-breathing propulsion in a rocket-based combined-cycle (RBCC) propulsion system to reduce the required propellant fraction. To achieve this goal the vehicle and propulsion system must be well integrated both aerodynamically and structurally to reduce weight. This study demonstrates the volumetric and structural efficiency of a vertical takeoff, horizontal landing, hypersonic vehicle with a circular cross section. A departure from the lifting body concepts, this design philosophy is even extended to the engines, which have semicircular nacelles symmetrically mounted on the vehicle. Material candidates with a potential for lightweight and simplicity have been selected from a set of near term technologies (five to ten years). To achieve the mission trajectory, preliminary weight estimates show the vehicle’s gross lift-off weight is 1.26 x 10(exp 6) lb. The structural configuration of the GTX vehicle and its propulsion system are described. The vehicle design benefits are presented, and key technical issues are highlighted. Author Air Breathing Boosters; Reusable Launch Vehicles; Propulsion System Configurations; Space Transportation

20010014167 NASA Dryden Flight Research Center, Edwards, CA USA Wind-Tunnel Investigations of Blunt-Body Drag Reduction Using Forebody Surface Roughness Whitmore, Stephen A.; Sprague, Stephanie; Naughton, Jonathan W.; Curry, Robert E., Technical Monitor; January 2001; In English; 39th Aerospace Sciences, 8-11 Jan. 2001, Reno, NV, USA; Original contains color illustrations Contract(s)/Grant(s): RTOP 718-20-00-E8-53; RTOP 529-35-24-M1-00-RR Report No.(s): NASA/TM-2001-210390; H-2439; NAS 1.15:210390; AIAA Paper 2001-0252; Copyright; Avail: CASI; A03, Hardcopy This paper presents results of wind-tunnel tests that demonstrate a novel drag reduction technique for blunt-based vehicles. For these tests, the forebody roughness of a blunt-based model was modified using micomachined surface overlays. As forebody roughness increases, boundary layer at the model aft thickens and reduces the shearing effect of external flow on the separated flow behind the base region, resulting in reduced base drag. For vehicle configurations with large base drag, existing data predict that a small increment in forebody friction drag will result in a relatively large decrease in base drag. If the added increment in forebody skin drag is optimized with respect to base drag, reducing the total drag of the configuration is possible. The wind-tunnel tests results conclusively demonstrate the existence of a forebody dragbase drag optimal point. The data demonstrate that the base drag coefficient corresponding to the drag minimum lies between 0.225 and 0.275, referenced to the base area. Most importantly, the data show a drag reduction of approximately 15% when the drag optimum is reached. When this drag reduction is scaled to the X-33 base area, drag savings approaching 45,000 N (10,000 lbf) can be realized. Author Wind Tunnel Tests; Aerodynamic Drag; Base Flow; Drag Reduction

85 20010003921 NASA Langley Research Center, Hampton, VA USA Infrared Sensing Aeroheating Flight Experiment: STS-96 Flight Results Blanchard, Robert C.; Wilmoth, Richard G.; Glass, Christopher E.; Merski, N. Ronald, Jr.; Berry, Scott A.; Bozung, Timothy J.; Tietjen, Alan; Wendt, Jodean; Dawson, Don; [2001]; In English; 39th Aerospace Sciences, 8-11 Jan. 2001, Reno, NV, USA; Original contains color illustrations Report No.(s): AIAA Paper 2001-0352; Copyright; Avail: CASI; A03, Hardcopy Major elements of an experiment called the Infrared Sensing Aeroheating Flight Experiment are discussed. The primary experiment goal is to provide reentry global temperature images from infrared measurements to define the characteristics of hypersonic boundary-layer transition during flight. Specifically, the experiment is to identify, monitor, and quantity hypersonic boundary layer windward surface transition of the X-33 vehicle during flight. In addition, the flight data will serve as a calibration and validation of current boundary layer transition prediction techniques, provide benchmark laminar, transitional, and fully turbulent global aeroheating data in order to validate existing wind tunnel and computational results, and to advance aeroheating technology. Shuttle Orbiter data from STS-96 used to validate the data acquisition and data reduction to global temperatures, in order to mitigate the experiment risks prior to the maiden flight of the X-33, is discussed. STS-96 reentry midwave (3-5 micron) infrared data were collected at the Ballistic Missile Defense Organization/Innovative Sciences and Technology Experimentation Facility site at NASA-Kennedy Space Center and subsequently mapped into global temperature contours using ground calibrations only. A series of image mapping techniques have been developed in order to compare each frame of infrared data with thermocouple data collected during the flight. Comparisons of the ground calibrated global temperature images with the corresponding thermocouple data are discussed. The differences are shown to be generally less than about 5%, which is comparable to the expected accuracy of both types of aeroheating measurements. Author Infrared Detectors; Aerodynamic Heating; Boundary Layer Transition; Turbulent Boundary Layer; Hypersonic Boundary Layer; X-33 Reusable Launch Vehicle; Space Shuttle Orbiters

20010000337 Lee and Associates, LLC, USA Support to X-33/Resusable Launch Vehicle Technology Program Dec. 05, 2000; In English Contract(s)/Grant(s): NASA Order H-29854-D; No Copyright; Avail: CASI; A03, Hardcopy The X-33 Guidance, Navigation, and Control (GN&C) Peer Review Team (PRT) was formed to assess the integrated X-33 vehicle GN&C system in order to identify any areas of disproportionate risk for initial flight. The eventual scope of the PRT assessment encompasses the GN&C algorithms, software, avionics, control effectors, applicable models, and testing. The initial (phase 1) focus of the PRT was on the GN&C algorithms and the Flight Control Actuation Subsystem (FCAS). The PRT held meetings during its phase 1 assessment at X-33 assembly facilities in Palmdale, California on May 17-18, 2000 and at Honeywell facilities in Tempe, Arizona on June 7, 2000. The purpose of these meetings was for the PRT members to get background briefings on the X-33 vehicle and for the PRT team to be briefed on the design basis and current status of the X-33 GN&C algorithms as well as the FCAS. The following material is covered in this PRT phase 1 final report. Some significant GN&C-related accomplishments by the X-33 development team are noted. Some topics are identified that were found during phase 1 to require fuller consideration when the PRT reconvenes in the future. Some new recommendations by the PRT to the X-33 program will likely result from a thorough assessment of these subjects. An initial list of recommendations from the PRT to the X-33 program is provided. These recommendations stem from topics that received adequate review by the PRT in phase 1. Significant technical observations by the PRT members as a result of the phase 1 meetings are detailed. (These are covered in an appendix.) There were many X-33 development team members who contributed to the technical information used by the PRT during the phase 1 assessment, who supported presentations to the PRT, and who helped to address the many questions posed by the PRT members at and after the phase 1 meetings. In all instances the interaction between the PRT and the X-33 development team members was cordial and very professional. The members of the PRT are grateful for the time and effort applied by all of these individuals and hope that the contents of this report will help to make the X-33 program a success. Derived from text X-33 Reusable Launch Vehicle; Guidance (Motion); Flight Control; Navigation

20000120758 Lee and Associates, LLC, USA Support to X-33/Reusable Launch Vehicle Technology Program Nov. 29, 2000; In English Contract(s)/Grant(s): NASA Order H-31417-D; No Copyright; Avail: CASI; A03, Hardcopy The Primary activities of Lee & Associates for the referenced Purchase Order has been in direct support of the

86 X-33/Reusable Launch Vehicle Technology Program. An independent review to evaluate the X-33 liquid hydrogen fuel tank failure, which recently occurred after-test of the starboard tank has been provided. The purpose of the Investigation team was to assess the tank design modifications, provide an assessment of the testing approach used by MSFC (Marshall Space Flight Center) in determining the flight worthiness of the tank, assessing the structural integrity, and determining the cause of the failure of the tank. The approach taken to satisfy the objectives has been for Lee & Associates to provide the expertise of Mr. Frank Key and Mr. Wayne Burton who have relevant experience from past programs and a strong background of experience in the fields critical to the success of the program. Mr. Key and Mr. Burton participated in the NASA established Failure Investigation Review Team to review the development and process data and to identify any design, testing or manufacturing weaknesses and potential problem areas. This approach worked well in satisfying the objectives and providing the Review Team with valuable information including the development of a Fault Tree. The detailed inputs were made orally in real time in the Review Team daily meetings. The results of the investigation were presented to the MSFC Center Director by the team on February 15, 2000. Attached are four charts taken from that presentation which includes 1) An executive summary, 2) The most probable cause, 3) Technology assessment, and 4) Technology Recommendations for Cryogenic tanks. Author X-33 Reusable Launch Vehicle; Fuel Tanks; Technology Assessment; Cryogenic Fluid Storage

20000120402 Boeing Co., Seattle, WA USA Advanced High Temperature Structural Seals Newquist, Charles W.; Verzemnieks, Juris; Keller, Peter C.; Shorey, Mark W.; Steinetz, Bruce, Technical Monitor; November 2000; In English Contract(s)/Grant(s): F33615-95-D-3203; RTOP 242-90-0D; AF Proj. 0012 Report No.(s): NASA/CR-2000-210522; E-12489; NAS 1.26:210522; No Copyright; Avail: CASI; A03, Hardcopy This program addresses the development of high temperature structural seals for control surfaces for a new generation of small reusable launch vehicles. Successful development will contribute significantly to the mission goal of reducing launch cost for small, 200 to 300 lb payloads. Development of high temperature seals is mission enabling. For instance, ineffective control surface seals can result in high temperature (3100 F) flows in the elevon area exceeding structural material limits. Longer sealing life will allow use for many missions before replacement, contributing to the reduction of hardware, operation and launch costs. During the first phase of this program the existing launch vehicle control surface sealing concepts were reviewed, the aerothermal environment for a high temperature seal design was analyzed and a mock up of an arc-jet test fixture for evaluating seal concepts was fabricated. Author Control Surfaces; Reusable Launch Vehicles; Seals (Stoppers); Structural Design; Research and Development

20000120039 FDC/NYMA, Inc., Hampton, VA USA International Space Station Evolution Data Book, Volume 2, Evolution Concepts Jorgensen, Catherine A., Editor; Antol, Jeffrey, Technical Monitor; October 2000; In English; Original contains color illustrations Contract(s)/Grant(s): NAS1-96013; RTOP 953-20-00 Report No.(s): NASA/SP-2000-6109/VOL2/REV1; L-18039B/VOL2/REV1; NAS 1.21:6109/VOL2/REV1; No Copyright; Avail: CASI; A07, Hardcopy This report provides a focused and in-depth look at the opportunities and drivers for the enhancement and evolution of the International Space Station (ISS) during assembly and beyond the assembly complete stage. These enhancements would expand and improve the current baseline capabilities of the ISS and help to facilitate the commercialization of the ISS by the private sector. Volume 1 provides the consolidated overview of the ISS baseline systems; information on the current facilities available for pressurized and unpressurized payloads; and information on current plans for crew availability and utilization, resource timelines and margin summaries including power, thermal, and storage volumes; and an overview of the vehicle traffic model. Volume 2 includes discussions of advanced technologies being investigated for use on the ISS and potential commercial utilization activities being examined including proposed design reference missions (DRM’s) and the technologies being assessed by the Pre-planned Program Improvement (P(sup 3) I) Working Group. This information is very high level and does not provide the relevant information necessary for detailed design efforts. This document is meant to educate readers on the ISS and to stimulate the generation of ideas for enhancement and utilization of the ISS, either by or for the government, academia, and commercial industry. Author International Space Station; Evolution (Development); Space Commercialization; Space Station Modules; Space Station Structures; Structural Design; Space Industrialization

87 20000108738 NASA Marshall Space Flight Center, Huntsville, AL USA A Damage Resistance Comparison Between Candidate Polymer Matrix Composite Feedline Materials Nettles, A. T; September 2000; In English Report No.(s): NASA/TM-2000-210482; NAS 1.15:210482; M-992; No Copyright; Avail: CASI; A03, Hardcopy As part of NASAs focused technology programs for future reusable launch vehicles, a task is underway to study the feasibility of using the polymer matrix composite feedlines instead of metal ones on propulsion systems. This is desirable to reduce weight and manufacturing costs. The task consists of comparing several prototype composite feedlines made by various methods. These methods are electron-beam curing, standard hand lay-up and autoclave cure, solvent assisted resin transfer molding, and thermoplastic tape laying. One of the critical technology drivers for composite components is resistance to foreign objects damage. This paper presents results of an experimental study of the damage resistance of the candidate materials that the prototype feedlines are manufactured from. The materials examined all have a 5-harness weave of IM7 as the ﬁber constituent (except for the thermoplastic, which is unidirectional tape laid up in a bidirectional conﬁguration). The resin tested were 977-6, PR 520, SE-SA-1, RS-E3 (e-beam curable), Cycom 823 and PEEK. The results showed that the 977-6 and PEEK were the most damage resistant in all tested cases. Author Composite Materials; Curing; Damage; Technology Assessment; Harnesses; Manufacturing; Polymer Matrix Composites

20000064109 NASA Glenn Research Center, Cleveland, OH USA Ceramic Matrix Composites (CMC) Life Prediction Method Development Levine, Stanley R.; Calomino, Anthony M.; Ellis, John R.; Halbig, Michael C.; Mital, Subodh K.; Murthy, Pappu L.; Opila, Elizabeth J.; Thomas, David J.; Thomas-Ogbuji, Linus U.; Verrilli, Michael J.; May 2000; In English; 24th Composites, Materials and Structures, 24-28 Jan. 2000, Cape Canaveral, FL, USA Contract(s)/Grant(s): RTOP 242-82-77 Report No.(s): NASA/TM-2000-210052; E-12253; NAS 1.15:210052; No Copyright; Avail: CASI; A03, Hardcopy Advanced launch systems (e.g., Reusable Launch Vehicle and other Shuttle Class concepts, Rocket-Based Combine Cycle, etc.), and interplanetary vehicles will very likely incorporate ﬁber reinforced ceramic matrix composites (CMC) in critical propulsion components. The use of CMC is highly desirable to save weight, to improve reuse capability, and to increase performance. CMC candidate applications are mission and cycle dependent and may include turbopump rotors, housings, combustors, nozzle injectors, exit cones or ramps, and throats. For reusable and single mission uses, accurate prediction of life is critical to mission success. The tools to accomplish life prediction are very immature and not oriented toward the behavior of carbon ﬁber reinforced silicon carbide (C/SiC), the primary system of interest for a variety of space propulsion applications. This paper describes an approach to satisfy the need to develop an integrated life prediction system for CMC that addresses mechanical durability due to cyclic and steady thermomechanical loads, and takes into account the impact of environmental degradation. Author Ceramic Matrix Composites; Technology Utilization; Silicon Carbides; Prediction Analysis Techniques; Life (Durability); Fiber Composites; Durability

20000062016 NASA Langley Research Center, Hampton, VA USA Parametric Weight Comparison of Advanced Metallic, Ceramic Tile, and Ceramic Blanket Thermal Protection Systems Myers, David E.; Martin, Carl J.; Blosser, Max L.; June 2000; In English Contract(s)/Grant(s): RTOP 242-33-01-50 Report No.(s): NASA/TM-2000-210289; L-17954; NAS 1.15:210289; No Copyright; Avail: CASI; A03, Hardcopy A parametric weight assessment of advanced metallic panel, ceramic blanket, and ceramic tile thermal protection systems (TPS) was conducted using an implicit, one-dimensional (I-D) ﬁnite element sizing code. This sizing code contained models to account for coatings fasteners, adhesives, and strain isolation pads. Atmospheric entry heating proﬁles for two vehicles, the Access to Space (ATS) vehicle and a proposed Reusable Launch Vehicle (RLV), were used to ensure that the trends were not unique to a certain trajectory. Ten TPS concepts were compared for a range of applied heat loads and substructural heat capacities to identify general trends. This study found the blanket TPS concepts have the lightest weights over the majority of their applicable ranges, and current technology ceramic tiles and metallic TPS concepts have similar weights. A proposed, state-of-the-art metallic system which uses a higher temperature alloy and efficient multilayer insulation was predicted to be

88 signiﬁcantly lighter than the ceramic tile stems and approaches blanket TPS weights for higher integrated heat loads. Author Ceramics; Fasteners; Isolation; Loads (Forces); Multilayer Insulation; Thermal Protection; Tiles; Trajectories

20000052923 NASA Langley Research Center, Hampton, VA USA Fatigue Crack Growth Rate Test Results for Al-Li 2195 Parent Metal, Variable Polarity Plasma Arc Welds and Friction Stir Welds Haﬂey, Robert A.; Wagner, John A.; Domack, Marcia S.; May 2000; In English Contract(s)/Grant(s): RTOP 242-20-05 Report No.(s): NASA/TM-2000-210098; NAS 1.15:210098; L-17984; No Copyright; Avail: CASI; A03, Hardcopy The fatigue crack growth rate of aluminum-lithium (Al-Li) alloy 2195 plate and weldments was determined at 200-F, ambient temperature and -320-F. The effects of stress ratio (R), welding process, orientation and thickness were studied. Results are compared with plate data from the Space Shuttle Super Lightweight Tank (SLWT) allowables program. Data from the current series of tests, both plate and weldment, falls within the range of data generated during the SLWT allowables program. Author Aluminum-Lithium Alloys; Crack Propagation; Welded Joints; Weld Tests; Cracking (Fracturing); Fatigue (Materials); Cracks

20000037777 NASA Marshall Space Flight Center, Huntsville, AL USA Test Report for NASA MSFC Support of the Linear Aerospike SR-71 Experiment (LASRE) Elam, S. K.; February 2000; In English Report No.(s): NASA/TM-2000-210076; NAS 1.15:210076; M-973; No Copyright; Avail: CASI; A07, Hardcopy The Linear Aerospike SR-71 Experiment (LASRE) was performed in support of the Reusable Launch Vehicle (RLV) program to help develop a linear aerospike engine. The objective of this program was to operate a small aerospike engine at various speeds and altitudes to determine how slipstreams affect the engine’s performance. The joint program between government and industry included NASA!s Dryden Flight Research Center, The Air Force’s Phillips Laboratory, NASA’s Marshall Space Flight Center, Lockheed Martin Skunkworks, Lockheed-Martin Astronautics, and Rocketdyne Division of Boeing North American. Ground testing of the LASRE engine produced two successful hot-fire tests, along with numerous cold flows to verify sequencing and operation before mounting the assembly on the SR-71. Once installed on the aircraft, flight testing performed several cold flows on the engine system at altitudes ranging from 30,000 to 50,000 feet and Mach numbers ranging from 0.9 to 1.5. The program was terminated before conducting hot-fires in flight because excessive leaks in the propellant supply systems could not be fixed to meet required safety levels without significant program cost and schedule impacts. Author Aerospike Engines; Reusable Launch Vehicles; SR-71 Aircraft; Liquid Propellant Rocket Engines; Propulsion System Performance; Engine Tests

20000033847 NASA Glenn Research Center, Cleveland, OH USA Recent Advances and Applications in Cryogenic Propellant Densification Technology Tomsik, Thomas M.; March 2000; In English; 12th Intersociety Cryogenic, 5-9 Mar. 2000, Atlanta, GA, USA Contract(s)/Grant(s): RTOP 242-33-0C Report No.(s): NASA/TM-2000-209941; E-12189; NAS 1.15:209941; No Copyright; Avail: CASI; A03, Hardcopy This purpose of this paper is to review several historical cryogenic test programs that were conducted at the NASA Glenn Research Center (GRC), Cleveland, Ohio over the past fifty years. More recently these technology programs were intended to study new and improved denser forms of liquid hydrogen (LH2) and liquid oxygen (LO2) cryogenic rocket fuels. Of particular interest are subcooled cryogenic propellants. This is due to the fact that they have a significantly higher density (eg. triple-point hydrogen, slush etc.), a lower vapor pressure and improved cooling capacity over the normal boiling point cryogen. This paper, which is intended to be a historical technology overview, will trace the past and recent development and testing of small and large-scale propellant densification production systems. Densifier units in the current GRC fuels program, were designed and are capable of processing subcooled LH2 and L02 propellant at the X33 Reusable Launch Vehicle (RLV) scale. One final objective of this technical briefing is to discuss some of the potential benefits and application which propellant densification technology may offer the industrial cryogenics production and end-user community. Density enhancements to

89 cryogenic propellants (LH2, LO2, CH4) in rocket propulsion and aerospace application have provided the opportunity to either increase performance of existing launch vehicles or to reduce the overall size, mass and cost of a new vehicle system. Author Cryogenic Rocket Propellants; Densiﬁcation; Technology Assessment; Liquid Hydrogen; Liquid Oxygen

20000024930 NASA Kennedy Space Center, Cocoa Beach, FL USA Design, Development,and Testing of Umbillical System Mechanisms for the X-33 Advanced Technology Demonstrator Littleﬁeld, Alan C.; Melton, Gregory S.; Dec. 1999; In English; 34th Aerospace Mechanisms, 10-12 May 2000, Greenbelt, MD, USA Report No.(s): NASA/TM-1999-208562; NAS1.15:208562; No Copyright; Avail: CASI; A03, Hardcopy The X-33 Advanced Technology Demonstrator is an un-piloted, vertical take-off, horizontal landing spacecraft. The purpose of the X-33 program is to demonstrate technologies that will dramatically lower the cost of access to space. The rocket-powered X-33 will reach an altitude of up to 100 km and speeds between Mach 13 and 15. Fifteen ﬂight tests are planned, beginning in 2000. Some of the key technologies demonstrated will be the linear aerospike engine, improved thermal protection systems, composite fuel tanks and reduced operational timelines. The X-33 vehicle umbilical connections provide monitoring, power, cooling, purge, and fueling capability during horizontal processing and vertical launch operations. Two ‘rise-ofF’ umbilicals for the X-33 have been developed, tested, and installed. The X-33 umbilical systems mechanisms incorporate several unique design features to simplify horizontal operations and provide reliable disconnect during launch. Author Cost Reduction; Technology Assessment; Flight Tests; Proving; Spacecraft Launching; Vertical Landing; X-33 Reusable Launch Vehicle

20000017931 NASA Glenn Research Center, Cleveland, OH USA Advanced Electric Propulsion for RLV Launched Geosynchronous Spacecraft Oleson, Steven; December 1999; In English; 26th 26th International Electric Propulsion Conference, 17-21 Oct. 1999, Kitakyushu, Japan Contract(s)/Grant(s): RTOP 632-1B-1C Report No.(s): NASA/TM-1999-209646; NAS 1.15:209646; IEPC-99-185; E-11994; No Copyright; Avail: CASI; A03, Hardcopy Solar Electric Propulsion (SEP) when used for station keeping and final orbit insertion has been shown to increase a geostationary satellite’s payload when launched by existing expendable launch vehicles. In the case of reusable launch vehicles or expendable launch vehicles where an upper stage is an expensive option, this methodology can be modified by using the existing on-board apogee chemical system to perform a perigee burn and then letting the electric propulsion system complete the transfer to geostationary orbit. The elimination of upper stages using on-board chemical and electric propulsion systems was thus examined for GEO spacecraft. Launch vehicle step-down from an Atlas IIAR to a Delta 7920 (no upper stage) was achieved using expanded on-board chemical tanks, 40 kW payload power for electric propulsion, and a 60 day elliptical to GEO SEP orbit insertion. Optimal combined chemical and electric trajectories were found using SEPSPOT. While Hall and ion thrusters provided launch vehicle step-down and even more payload for longer insertion times, NH3 arcjets had insufficient performance to allow launch vehicle step-down. Degradation levels were only 5% to 7% for launch step-down cases using advanced solar arrays. Results were parameterized to allow comparisons for future reusable launch vehicles. Results showed that for an 8 W/kg initial power/launch mass power density spacecraft, 50% to 100% more payload can be launched using this method. Author Electric Propulsion; Propulsion System Configurations; Propulsion System Performance; Spacecraft Propulsion; Reusable Launch Vehicles; Geosynchronous Orbits; Synchronous Platforms; Synchronous Satellites

20000004220 NASA Langley Research Center, Hampton,VA USA, NASA Langley Research Center, Hampton,VA USA Compute as Fast as the Engineers Can Think! ULTRAFAST COMPUTING TEAM FINAL REPORT Biedron, R. T.; Mehrotra, P.; Nelson, M. L.; Preston, M. L.; Rehder, J. J.; Rogersm J. L.; Rudy, D. H.; Sobieski, J.; Storaasli, O. O.; Sep. 1999; In English Contract(s)/Grant(s): RTOP 509-10-31-03 Report No.(s): NASA/TM-1999-209715; L-17891; NAS 1.15:209715; No Copyright; Avail: CASI; A04, Hardcopy This report documents ﬁndings and recommendations by the Ultrafast Computing Team (UCT). In the period 10-12/98,

90 UCT reviewed design case scenarios for a supersonic transport and a reusable launch vehicle to derive computing requirements necessary for support of a design process with efficiency so radically improved that human thought rather than the computer paces the process. Assessment of the present computing capability against the above requirements indicated a need for further improvement in computing speed by several orders of magnitude to reduce time to solution from tens of hours to seconds in major applications. Evaluation of the trends in computer technology revealed a potential to attain the postulated improvement by further increases of single processor performance combined with massively parallel processing in a heterogeneous environment. However, utilization of massively parallel processing to its full capability will require redevelopment of the engineering analysis and optimization methods, including invention of new paradigms. To that end UCT recommends initiation of a new activity at LaRC called Computational Engineering for development of new methods and tools geared to the new computer architectures in disciplines, their coordination, and validation and beneﬁt demonstration through applications. Author Computer Systems Design; Architecture (Computers); Axioms; Heterogeneity; Massively Parallel Processors

19990117247 NASA Langley Research Center, Hampton, VA USA Multidisciplinary Optimization Branch Experience Using iSIGHT Software Padula, S. L.; Korte, J. J.; Dunn, H. J.; Salas, A. O.; November 1999; In English Contract(s)/Grant(s): RTOP 242-33-03-05; RTOP 581-20-31-01; RTOP 509-10-11-01 Report No.(s): NASA/TM-1999-209714; NAS 1.15:209714; L-17914; No Copyright; Avail: CASI; A03, Hardcopy The Multidisciplinary Optimization (MDO) Branch at NASA Langley is investigating frameworks for supporting multidisciplinary analysis and optimization research. A framework provides software and system services to integrate computational tasks and allows the researcher to concentrate more on the application and less on the programming details. A framework also provides a common working environment and a full range of optimization tools, and so increases the productivity of multidisciplinary research teams. Finally, a framework enables staff members to develop applications for use by disciplinary experts in other organizations. This year, the MDO Branch has gained experience with the iSIGHT framework. This paper describes experiences with four aerospace applications, including: (1) reusable launch vehicle sizing, (2) aerospike nozzle design, (3) low-noise rotorcraft trajectories, and (4) acoustic liner design. Brief overviews of each problem are provided, including the number and type of disciplinary codes and computation time estimates. In addition, the optimization methods, objective functions, design variables, and constraints are described for each problem. For each case, discussions on the advantages and disadvantages of using the iSIGHT framework are provided as well as notes on the ease of use of various advanced features and suggestions for areas of improvement. Author Design Analysis; Multidisciplinary Design Optimization; Computer Aided Design; Systems Engineering; Computer Programs; Evaluation; Computer Systems Performance

19990113121 NASA Dryden Flight Research Center, Edwards, CA USA Evaluation of the Linear Aerospike SR-71 Experiment (LASRE) Oxygen Sensor Ennix, Kimberly A.; Corpening, Griffin P.; Jarvis, Michele; Chiles, Harry R.; November 1999; In English; 9th 9th International Space Planes and Hypersonic Systems, 1-5 Nov. 1999, Norfolk, VA, USA Contract(s)/Grant(s): RTOP 529-31-44-00-32 Report No.(s): NASA/TM-1999-206589; NAS 1.15:206589; H-2377; No Copyright; Avail: CASI; A03, Hardcopy The Linear Aerospike SR-71 Experiment (LASRE) was a propulsion flight experiment for advanced space vehicles such as the X-33 and reusable launch vehicle. A linear aerospike rocket engine was integrated into a semi-span of an X-33-like lifting body shape (model), and carried on top of an SR-71 aircraft at NASA Dryden Flight Research Center. Because no flight data existed for aerospike nozzles, the primary objective of the LASRE flight experiment was to evaluate flight effects on the engine performance over a range of altitudes and Mach numbers. Because it contained a large quantity of energy in the form of fuel, oxidizer, hypergolics, and gases at very high pressures, the LASRE propulsion system posed a major hazard for fire or explosion. Therefore, a propulsion-hazard mitigation system was created for LASRE that included a nitrogen purge system. Oxygen sensors were a critical part of the nitrogen purge system because they measured purge operation and effectiveness. Because the available oxygen sensors were not designed for flight testing, a laboratory study investigated oxygen-sensor characteristics and accuracy over a range of altitudes and oxygen concentrations. Laboratory test data made it possible to properly calibrate the sensors for flight. Such data also provided a more accurate error prediction than the manufacturer’s

91 specification. This predictive accuracy increased confidence in the sensor output during critical phases of the flight. This paper presents the findings of this laboratory test. Author Aerospike Engines; Liquid Nitrogen; Gas Detectors; Oxygen Analyzers; Purging

19990113117 Starcraft Boosters, Inc., Houston, TX USA, Starcraft Boosters, Inc., Canyon Lake, TX USA The StarBooster System: A Cargo Aircraft for Space Davis, Hubert P.; Dula, Arthur M.; McLaughlin, Don; Frassanito, John; Andrews, Jason, Editor; Oct. 27, 1999; In English Contract(s)/Grant(s): NASA Order H-30550-D Report No.(s): SBI-1999-001A; No Copyright; Avail: CASI; A04, Hardcopy Starcraft Boosters has developed a different approach for lowering the cost of access to space. We propose developing a new aircraft that will house an existing expendable rocket stage. This vehicle, termed StarBooster, will be the ﬁrst stage of a family of launch vehicles. By combining these elements, we believe we can reduce the cost and risk of ﬁelding a new partially reusable launch system. This report summarizes the work performed on the StarBooster concept since the company’s inception in 1996. Detailed analyses are on-going and future reports will focus on the maturation of the vehicle and system design. Derived from text Booster Rocket Engines; Reusable Launch Vehicles; Propulsion System Performance; Space Transportation

19990110310 NASA Dryden Flight Research Center, Edwards, CA USA Flush Airdata Sensing (FADS) System Calibration Procedures and Results for Blunt Forebodies Cobleigh, Brent R.; Whitmore, Stephen A.; Haering, Edward A., Jr.; Borrer, Jerry; Roback, V. Eric; November 1999; In English; 9th International Space Planes and Hypersonic Systems and Technologies, 1-5 Nov. 1999, Norfolk, VA, USA Contract(s)/Grant(s): RTOP 242-33-02-00-23 Report No.(s): NASA/TP-1999-209012; NAS 1.60:209012; H-2379; AIAA Paper 99-4816; No Copyright; Avail: CASI; A03, Hardcopy Blunt-forebody pressure data are used to study the behavior of the NASA Dryden Flight Research Center flush airdata sensing (FADS) pressure model and solution algorithm. The model relates surface pressure measurements to the airdata state. Spliced from the potential flow solution for uniform flow over a sphere and the modified Newtonian impact theory, the model was shown to apply to a wide range of blunt-forebody shapes and Mach numbers. Calibrations of a sphere, spherical cones, a Rankine half body, and the F-14, F/A-18, X-33, X-34, and X-38 configurations are shown. The three calibration parameters are well-behaved from Mach 0.25 to Mach 5.0, an angle-of-attack range extending to greater than 30 deg, and an angle-of-sideslip range extending to greater than 15 deg. Contrary to the sharp calibration changes found on traditional pitot-static systems at transonic speeds, the FADS calibrations are smooth, monotonic functions of Mach number and effective angles of attack and sideslip. Because the FADS calibration is sensitive to pressure port location, detailed measurements of the actual pressure port locations on the flight vehicle are required and the wind-tunnel calibration model should have pressure ports in similar locations. The procedure for calibrating a FADS system is outlined. Author Calibrating; Detection; Forebodies; Mathematical Models; Pressure Measurement

19990108597 NASA Dryden Flight Research Center, Edwards,CA USA Evaluation of the Linear Aerospike SR-71 Experiment (LASRE) Oxygen Sensor Ennix, Kimberly A.; Corpening, Griffin P.; Jarvis, Michele; Chiles, Harry R.; November 1999; In English; 9th Space Planes and Hypersonic Systems, 1-5 Nov. 1999, Norfolk, VA, USA; Original contains color illustration Contract(s)/Grant(s): RTOP 529-31-44 Report No.(s): NASA/TM-1999-206589; H-2377; NAS 1.15:206589; Copyright; Avail: CASI; A03, Hardcopy The Linear Aerospike SR-71 Experiment (LASRE) was a propulsion flight experiment for advanced space vehicles such as the X-33 and reusable launch vehicle. A linear aerospike rocket engine was integrated into a semi-span of an X-33-like lifting body shape (model), and carried on top of an SR-71 aircraft at NASA Dryden Flight Research Center. Because no flight data existed for aerospike nozzles, the primary objective of the LASRE flight experiment was to evaluate flight effects on the engine performance over a range of altitudes and Mach numbers. Because it contained a large quantity of energy in the form of fuel, oxidizer, hypergolics, and gases at very high pressures, the LASRE propulsion system posed a major hazard for fire or explosion. Therefore, a propulsion-hazard mitigation system was created for LASRE that included a nitrogen purge system.

92 Oxygen sensors were a critical part of the nitrogen purge system because they measured purge operation and effectiveness. Because the available oxygen sensors were not designed for flight testing, a laboratory study investigated oxygen-sensor characteristics and accuracy over a range of altitudes and oxygen concentrations. Laboratory test data made it possible to properly calibrate the sensors for flight. Such data also provided a more accurate error prediction than the manufacturer’s specification. This predictive accuracy increased confidence in the sensor output during critical phases of the flight. This paper presents the findings of this laboratory test. Author Aerospike Engines; Oxygen; SR-71 Aircraft; Gas Detectors; Propulsion System Configurations

19990106559 NASA Dryden Flight Research Center, Edwards, CA USA Propellant Feed System Leak Detection: Lessons Learned From the Linear Aerospike SR-71 Experiment (LASRE) Hass, Neal; Mizukami, Masashi; Neal, Bradford A.; St. John, Clinton; Beil, Robert J.; Griffin, Timothy P.; November 1999; In English; 9th International Space Planes and Hypersonic Systems, 1-5 Nov. 1999, Norfolk, VA, USA Contract(s)/Grant(s): RTOP 243-33-02 Report No.(s): NASA/TM-1999-206590; NAS 1.15:206590; H-2378; Copyright; Avail: CASI; A03, Hardcopy This paper presents pertinent results and assessment of propellant feed system leak detection as applied to the Linear Aerospike SR-71 Experiment (LASRE) program flown at the NASA Dryden Flight Research Center, Edwards, California. The LASRE was a flight test of an aerospike rocket engine using liquid oxygen and high-pressure gaseous hydrogen as propellants. The flight safety of the crew and the experiment demanded proven technologies and techniques that could detect leaks and assess the integrity of hazardous propellant feed systems. Point source detection and systematic detection were used. Point source detection was adequate for catching gross leakage from components of the propellant feed systems, but insufficient for clearing LASRE to levels of acceptability. Systematic detection, which used high-resolution instrumentation to evaluate the health of the system within a closed volume, provided a better means for assessing leak hazards. Oxygen sensors detected a leak rate of approximately 0.04 cubic inches per second of liquid oxygen. Pressure sensor data revealed speculated cryogenic boiloff through the fittings of the oxygen system, but location of the source(s) was indeterminable. Ultimately, LASRE was cancelled because leak detection techniques were unable to verify that oxygen levels could be maintained below flammability limits. Author Leakage; Aerospike Engines; Gas Detectors; Pressure Sensors; Feed Systems; Flight Tests

19990105883 NASA Dryden Flight Research Center, Edwards, CA USA Flush Airdata Sensing (FADS) System Calibration Procedures and Results for Blunt Forebodies Cobleigh, Brent R.; Whitmore, Stephen A.; Haering, Edward A., Jr.; Borrer, Jerry; Roback, V. Eric; November 1999; In English; 9th Space Planes and Hypersonic Systems and Technologies, 1-5 Nov. 1999, Norfolk, VA, USA Contract(s)/Grant(s): RTOP 242-33-02-00-23 Report No.(s): NASA/TP-1999-209012; NAS 1.60:209012; H-2379; No Copyright; Avail: CASI; A03, Hardcopy Blunt-forebody pressure data are used to study the behavior of the NASA Dryden Flight Research Center flush airdata sensing (FADS) pressure model and solution algorithm. The model relates surface pressure measurements to the airdata state. Spliced from the potential flow solution for uniform flow over a sphere and the modified Newtonian impact theory, the model was shown to apply to a wide range of blunt-forebody shapes and Mach numbers. Calibrations of a sphere, spherical cones, a Rankine half body, and the F-14, F/A-18, X-33, X-34, and X-38 configurations are shown. The three calibration parameters are well-behaved from Mach 0.25 to Mach 5.0, an angle-of-attack range extending to greater than 30 deg, and an angle-of-sideslip range extending to greater than 15 deg. Contrary to the sharp calibration changes found on traditional pitot-static systems at transonic speeds, the FADS calibrations are smooth, monotonic functions of Mach number and effective angles of attack and sideslip. Because the FADS calibration is sensitive to pressure port location, detailed measurements of the actual pressure port locations on the flight vehicle are required and the wind-tunnel calibration model should have pressure ports in similar locations. The procedure for calibrating a FADS system is outlined. Author Procedures; Calibrating; Forebodies; Mathematical Models; Algorithms; Potential Flow; Pressure Measurement; Newton Theory

93 19990100643 Lockheed Martin Engineering and Sciences Co., Hampton, VA USA An Inviscid Computational Study of an X-33 Configuration at Hypersonic Speeds Prabhu, Ramadas K.; July 1999; In English Contract(s)/Grant(s): NAS1-96014; RTOP 242-80-01-01 Report No.(s): NASA/CR-1999-209366; NAS 1.26:209366; No Copyright; Avail: CASI; A03, Hardcopy This report documents the results of a study conducted to compute the inviscid longitudinal aerodynamic characteristics of a simplified X-33 configuration. The major components of the X-33 vehicle, namely the body, the canted fin, the vertical fin, and the body-flap, were simulated in the CFD (Computational Fluid Dynamic) model. The rear-ward facing surfaces at the base including the aerospike engine surfaces were not simulated. The FELISA software package consisting of an unstructured surface and volume grid generator and two inviscid flow solvers was used for this study. Computations were made for Mach 4.96, 6.0, and 10.0 with perfect gas air option, and for Mach 10 with equilibrium air option with flow condition of a typical point on the X-33 flight trajectory. Computations were also made with CF4 gas option at Mach 6.0 to simulate the CF4 tunnel flow condition. An angle of attack range of 12 to 48 deg was covered. The CFD results were compared with available wind tunnel data. Comparison was good at low angles of attack; at higher angles of attack (beyond 25 deg) some differences were found in the pitching moment. These differences progressively increased with increase in angle of attack, and are attributed to the viscous effects. However, the computed results showed the trends exhibited by the wind tunnel data. Author Computational Fluid Dynamics; Inviscid Flow; Viscous Flow; Unstructured Grids (Mathematics); X-33 Reusable Launch Vehicle; Venturestar Launch Vehicle; Aerospike Engines

19990094229 NASA Langley Research Center, Hampton,VA USA Cyclic Cryogenic Thermal-Mechanical Testing of an X-33/RLV Liquid Oxygen Tank Concept Rivers, H. Kevin; September 1999; In English; Original contains color illustrations Contract(s)/Grant(s): RTOP 242-23-02-13 Report No.(s): NASA/TM-1999-209560; L-17818; NAS 1.15:209560; No Copyright; Avail: CASI; A03, Hardcopy An important step in developing a cost-effective, reusable, launch vehicle is the development of durable, lightweight, insulated, cryogenic propellant tanks. Current cryogenic tanks are expendable so most of the existing technology is not directly applicable to future launch vehicles. As part of the X-33/Reusable Launch Vehicle (RLV) Program, an experimental apparatus developed at the NASA Langley Research Center for evaluating the effects of combined, cyclic, thermal and mechanical loading on cryogenic tank concepts was used to evaluate cryogenic propellant tank concepts for Lockheed-Martin Michoud Space Systems. An aluminum-lithium (Al 2195) liquid oxygen tank concept, insulated with SS-1171 and PDL-1034 cryogenic insulation, is tested under simulated mission conditions, and the results of those tests are reported. The tests consists of twenty-five simulated Launch/Abort missions and twenty-five simulated flight missions with temperatures ranging from -320 F to 350 F and a maximum mechanical load of 71,300 lb. in tension. Author Cyclic Loads; Storage Tanks; Propellant Tanks; Load Tests; Load Testing Machines; Flight Simulation

19990081112 NASA Glenn Research Center, Cleveland, OH USA Analysis of a New Rocket-Based Combined-Cycle Engine Concept at Low Speed Yungster, S.; Trefny, C. J.; September 1999; In English; 35th Joint Propulsion, 20-24 Jun. 1999, Los Angeles, CA, USA Contract(s)/Grant(s): NCC3-542; RTOP 505-90-5K Report No.(s): NASA/TM-1999-209393; NAS 1.15:209393; E-11824; AIAA Paper 99-2393; ICOMP-99-05; No Copyright; Avail: CASI; A03, Hardcopy An analysis of the Independent Ramjet Stream (IRS) cycle is presented. The IRS cycle is a variation of the conventional ejector-Ramjet, and is used at low speed in a rocket-based combined-cycle (RBCC) propulsion system. In this new cycle, complete mixing between the rocket and ramjet streams is not required, and a single rocket chamber can be used without a long mixing duct. Furthermore, this concept allows flexibility in controlling the thermal choke process. The resulting propulsion system is intended to be simpler, more robust, and lighter than an ejector-ramjet. The performance characteristics of the IRS cycle are analyzed for a new single-stage-to-orbit (SSTO) launch vehicle concept, known as ’Trailblazer.‘ The study is based on a quasi-one-dimensional model of the rocket and air streams at speeds ranging from lift-off to Mach 3. The numerical formulation is described in detail. A performance comparison between the IRS and ejector-ramjet cycles is also presented. Author Ramjet Engines; Rocket Engines; Single Stage to Orbit Vehicles; Propulsion System Performance; Propulsion System Configurations; Spacecraft Configurations

94 19990062664 NASA Glenn Research Center, Cleveland, OH USA Inlet Development for a Rocket Based Combined Cycle, Single Stage to Orbit Vehicle Using Computational Fluid Dynamics DeBonis, J. R.; Trefny, C. J.; Steffen, C. J., Jr.; June 1999; In English; 35th Joint Propulsion, 20-24 Jun. 1999, Los Angeles, CA, USA Contract(s)/Grant(s): RTOP 523-61-13 Report No.(s): NASA/TM-1999-209279; NAS 1.15:209279; AIAA Paper 99-2239; E-11744; No Copyright; Avail: CASI; A03, Hardcopy Design and analysis of the inlet for a rocket based combined cycle engine is discussed. Computational ﬂuid dynamics was used in both the design and subsequent analysis. Reynolds averaged Navier-Stokes simulations were performed using both perfect gas and real gas assumptions. An inlet design that operates over the required Mach number range from 0 to 12 was produced. Performance data for cycle analysis was post processed using a stream thrust averaging technique. A detailed performance database for cycle analysis is presented. The effect ot vehicle forebody compression on air capture is also examined. Author Inlet Flow; Data Bases; Reliability Analysis; Reynolds Averaging; Simulation; Orbital Maneuvers; Design Analysis

19990062176 NASA Dryden Flight Research Center, Edwards,CA USA Reconfigurable Flight Control Designs With Application to the X-33 Vehicle Burken, John J.; Lu, Ping; Wu, Zhenglu; August 1999; In English; Guidance Navigation and Control, 9-11 Aug. 1999, Portland, OR, USA Contract(s)/Grant(s): RTOP 242-33-02-00-23 Report No.(s): NASA/TM-1999-206582; H-2345; NAS 1.15:206582; AIAA Paper 99-4134; No Copyright; Avail: CASI; A03, Hardcopy Two methods for control system reconfiguration have been investigated. The first method is a robust servomechanism control approach (optimal tracking problem) that is a generalization of the classical proportional-plus-integral control to multiple input-multiple output systems. The second method is a control-allocation approach based on a quadratic programming formulation. A globally convergent fixed-point iteration algorithm has been developed to make onboard implementation of this method feasible. These methods have been applied to reconfigurable entry flight control design for the X-33 vehicle. Examples presented demonstrate simultaneous tracking of angle-of-attack and roll angle commands during failures of the right body flap actuator. Although simulations demonstrate success of the first method in most cases, the control-allocation method appears to provide uniformly better performance in all cases. Author Control Systems Design; Mimo (Control Systems); X-33 Reusable Launch Vehicle; Servomechanisms; Proportional Control; Flight Control; Launch Vehicle Configurations

19990053005 NASA Glenn Research Center, Cleveland, OH USA NASA Glenn Research Center’s Hypersonic Propulsion Program Palac, Donald T.; May 1999; In English; Airbreathing Engines, 5-10 Sep. 1999, Florence, Italy Contract(s)/Grant(s): RTOP 523-61-13 Report No.(s): NASA/TM-1999-209185; E-11696; NAS 1.15:209185; No Copyright; Avail: CASI; A02, Hardcopy NASA Glenn Research Center (GRC), as NASA’s lead center for aeropropulsion, is responding to the challenge of reducing the cost of space transportation through the integration of air-breathing propulsion into launch vehicles. Air- breathing launch vehicle (ABLV) propulsion requires a marked departure from traditional propulsion applications. and stretches the technology of both rocket and air-breathing propulsion. In addition, the demands of the space launch mission require an unprecedented level of integration of propulsion and vehicle systems. GRC is responding with a program with rocket-based combined cycle (RBCC) propulsion technology as its main focus. RBCC offers the potential for simplicity, robustness, and performance that may enable low-cost single-stage-to-orbit (SSTO) transportation. Other technologies, notably turbine-based combined cycle (TBCC) propulsion, offer beneﬁts such as increased robustness and greater mission ﬂexibility, and are being advanced, at a slower pace, as part of GRC’s program in hypersonics. Author Air Breathing Boosters; Air Breathing Engines; Launch Vehicles; Spacecraft Launching

95 19990036776 NASA Langley Research Center, Hampton,VA USA X-33 (Rev-F) Aeroheating Results of Test 6770 in NASA Langley 20-Inch Mach 6 Air Tunnel Berry, Scott A.; Horvath, Thomas J.; Kowalkowski, Matthew K.; Liechty, Derek S.; March 1999; In English; Original contains color illustrations Contract(s)/Grant(s): RTOP 242-80-01-01 Report No.(s): NASA/TM-1999-209122; L-17824; NAS 1.15:209122; No Copyright; Avail: CASI; A05, Hardcopy Aeroheating characteristics of the X-33 Rev-F configuration have been experimentally examined in the Langley 20-Inch Mach 6 Air Tunnel (Test 6770). Global surface heat transfer distributions, surface streamline patterns, and shock shapes were measured on a 0.013-scale model at Mach 6 in air. Parametric variations include angles-of-attack of 20-deg, 30-deg, and 40-deg; Reynolds numbers based on model length of 0.9 to 4.9 million; and body-flap deflections of 10-deg and 20-deg. The effects of discrete roughness elements on boundary layer transition, which included trip height, size, and location, both on and off the windward centerline, were investigated. This document is intended to serve as a quick release of preliminary data to the X-33 program; analysis is limited to observations of the experimental trends in order to expedite dissemination. Author X-33 Reusable Launch Vehicle; Aerodynamic Heating; Aerodynamic Configurations; Wind Tunnel Tests; Hypersonic Speed; Heat Transfer; Boundary Layer Transition; Aerothermodynamics

19990032557 NASA Marshall Space Flight Center, Huntsville, AL USA General Public Space Travel and Tourism, Volume 1, Executive Summary ONeil, Daniel, Compiler; Bekey, Ivan; Mankins, John; Rogers, Thomas F.; Stallmer, Eric W.; March 1998; In English; Summary of a Space Act Agreement Study, 19-21 Feb. 1997, Washington, DC, USA Report No.(s): NASA/NP-1998-03-11-MSFC/VOL1; NAS 1.83:03-11-MSFC/VOL1; No Copyright; Avail: CASI; A03, Hardcopy Travel and tourism is one of the world’s largest businesses. Its gross revenues exceed $400 billion per year in the U.S. alone, and it is our second largest employer. U.S. private sector business revenues in the space information area now approximate $10 billion per year, and are increasing rapidly. Not so in the human spaceflight area. After spending $100s of billions (1998 dollars) in public funds thereon, and continuing to spend over $5 billion per year, the government is still the only customer for human spaceflight goods and services. Serious and detailed consideration was first given to the possibility of space being opened up to trips by the general public three decades ago, and some initial attempts to do so were made a dozen years ago. But the difficulties were great and the Challenger disaster put an end to them. In recent years professional space tourism studies have been conducted in the UK, Germany and, especially, Japan. In the U.S., technological progress has been pronounced; we have had nearly a decade’s experience in seeing our astronauts travel to-from low Earth orbit (LEO) safely, and we expect to commence assembly of a LEO space station housing a half-dozen people this year. Too, NASA and our space industry now have new and promising space transportation development programs underway, especially the X-33 and X-34 programs, and some related, further generation, basic technology development programs. And five private companies are also working on the design of new surface - LEO vehicles. The first professional space tourism market studies have been conducted in several countries in the past few years, especially in Japan and here. The U.S. study makes it clear that, conceptually, tens of millions of us would like to take a trip to space if we could do so with reasonable safety, comfort and reliability, and at an acceptable price. Initial businesses will address the desires of those willing to pay a greater price and accept a greater risk. A two-year cooperative Space Act agreement study has been conducted by our National Aeronautics and Space Administration and the Space Transportation Association. It was conducted by NASA and STA study leaders drawing upon the competence, experience and hard-nosed imagination of a national Steering Group and scores of attendees at a multi-day Workshop. The study has involved scores of professionals and business people from various areas: astronauts; space booster technology and operations professionals; a hotel architect and a hotel operator; an airline planner; insurance underwriters; space sickness experts; space theme park designers; space and travel and tourism association and business executives; a space-related financier; university tourism and space policy experts; present and former space-responsible government officials; space entrepreneurs; space writers; This study concludes that serious national attention should now be given to activities that would enable the expansion of today’s terrestrial space tourism businesses, and the creation of in-space travel and tourism businesses. Indeed, it concludes that, in time, it should become a very important part of our Country’s overall commercial and civil space business-program structure. Author Aerospace Engineering; Low Earth Orbits; NASA Programs; Space Flight; Space Stations; Space Transportation; Travel; Space Habitats

96 19990028385 NASA Marshall Space Flight Center, Huntsville, AL USA A Strategy for Integrating a Large Finite Element Model Using MSC NASTRAN/PATRAN: X-33 Lessons Learned McGhee, D. S.; March 1999; In English Report No.(s): NASA/TM-1999-209201; NAS 1.15:209201; M-917; No Copyright; Avail: CASI; A03, Hardcopy The X-33 vehicle is an advanced technology demonstrator sponsored by NASA. For the past 3 years the Structural Dynamics and Loads Branch of NASA’s Marshall Space Flight Center has had the task of integrating the X-33 vehicle structural finite element model. In that time, five versions of the integrated vehicle model have been produced and a strategy has evolved that would benefit anyone given the task of integrating structural finite element models that have been generated by various modelers and companies. The strategy that has been presented here consists of six decisions that need to be made: purpose of models, units, common materials list, model numbering, interface control, and archive format. This strategy has been proven and expanded from experience on the X-33 vehicle. Author X-33 Reusable Launch Vehicle; Finite Element Method; Models; Nastran

19990026605 NASA Dryden Flight Research Center, Edwards,CA USA A Base Drag Reduction Experiment on the X-33 Linear Aerospike SR-71 Experiment (LASRE) Flight Program Whitmore, Stephen A.; Moes, Timothy R.; Mar. 1999; In English; 37th Aerospace Sciences, 11-14 Jan. 1999, Reno, NV, USA Contract(s)/Grant(s): RTOP 242-33-02 Report No.(s): NASA/TM-1999-206575; H-2333; NAS 1.15:206575; AIAA Paper 99-0277; No Copyright; Avail: CASI; A03, Hardcopy Drag reduction tests were conducted on the LASRE/X-33 flight experiment. The LASRE experiment is a flight test of a roughly 20% scale model of an X-33 forebody with a single aerospike engine at the rear. The experiment apparatus is mounted on top of an SR-71 aircraft. This paper suggests a method for reducing base drag by adding surface roughness along the forebody. Calculations show a potential for base drag reductions of 8-14%. Flight results corroborate the base drag reduction, with actual reductions of 15% in the high-subsonic flight regime. An unexpected result of this experiment is that drag benefits were shown to persist well into the supersonic flight regime. Flight results show no overall net drag reduction. Applied surface roughness causes forebody pressures to rise and offset base drag reductions. Apparently the grit displaced streamlines outward, causing forebody compression. Results of the LASRE drag experiments are inconclusive and more work is needed. Clearly, however, the forebody grit application works as a viable drag reduction tool. Author Aerodynamic Drag; Drag Reduction; Flight Tests; Drag Measurement

19990025759 NASA Langley Research Center, Hampton, VA USA Effective Thermal Conductivity of High Temperature Insulations for Reusable Launch Vehicles Daryabeigi, Kamran; February 1999; In English Contract(s)/Grant(s): RTOP 242-33-03-29 Report No.(s): NASA/TM-1999-208972; NAS 1.15:208972; L-17808; No Copyright; Avail: CASI; A03, Hardcopy An experimental apparatus was designed to measure the effective thermal conductivity of various high temperature insulations subject to large temperature gradients representative of typical launch vehicle re-entry aerodynamic heating conditions. The insulation sample cold side was maintained around room temperature, while the hot side was heated to temperatures as high as 1800 degrees Fahrenheit. The environmental pressure was varied from 0.0001 to 760 torr. All the measurements were performed in a dry gaseous nitrogen environment. The effective thermal conductivity of Saffil, Q-Fiber felt, Cerachrome, and three multi-layer insulation conﬁgurations were measured. Author Thermal Conductivity; Multilayer Insulation; Reusable Launch Vehicles; Reentry Effects; Thermal Insulation

19990018443 NASA Langley Research Center, Hampton,VA USA Aeronautical Engineering: A Continuing Bibliography with Indexes Feb. 05, 1999; In English Report No.(s): NASA/SP-1999-7037/SUPPL393; NAS 1.21:7037/SUPPL393; No Copyright; Avail: CASI; A04, Hardcopy This supplemental issue of Aeronautical Engineering: A Continuing Bibliography with Indexes lists reports, articles, and other documents recently announced in the NASA STI Database. The coverage includes documents on the engineering and theoretical aspects of design, construction, evaluation, testing, operation, and performance of aircraft (including aircraft

97 engines) and associated components, equipment, and systems. It also includes research and development in aerodynamics, aeronautics, and ground support equipment for aeronautical vehicles. Each entry in the publication consists of a standard bibliographic citation accompanied, in most cases, by an abstract. The NASA CASI price code table, addresses of organizations, and document availability information are included before the abstract section. Two indexes-subject and author are included after the abstract section. Derived from text Bibliographies; Aeronautical Engineering

19990009863 NASA Lewis Research Center, Cleveland,OH USA A Hazardous Gas Detection System for Aerospace and Commercial Applications Hunter, G. W.; Neudeck, P. G.; Chen, L. - Y.; Makel, D. B.; Liu, C. C.; Wu, Q. H.; Knight, D.; Nov. 1998; In English; 34th Joint Propulsion, 12-15 Jul. 1998, Cleveland, OH, USA; Original contains color illustrations Contract(s)/Grant(s): RTOP 254-02-0A Report No.(s): NASA/TM-1998-208817; E-11419; NAS 1.15:208817; AIAA Paper 98-3614; No Copyright; Avail: CASI; A03, Hardcopy The detection of explosive conditions in aerospace propulsion applications is important for safety and economic reasons. Microfabricated hydrogen, oxygen, and hydrocarbon sensors as well as the accompanying hardware and software are being developed for a range of aerospace safety applications. The development of these sensors is being done using MEMS (Micro ElectroMechanical Systems) based technology and SiC-based semiconductor technology. The hardware and software allows control and interrogation of each sensor head and reduces accompanying cabling through multiplexing. These systems are being applied on the X-33 and on an upcoming STS-95 Shuttle mission. A number of commercial applications are also being pursued. It is concluded that this MEMS-based technology has signiﬁcant potential to reduce costs and increase safety in a variety of aerospace applications. Author Aerospace Engineering; Gas Detectors; Aerospace Safety; Space Transportation System; Space Transportation System Flights

19990008856 NASA Lewis Research Center, Cleveland, OH USA A Hazardous Gas Detection System for Aerospace and Commercial Applications Hunter, G. W.; Neudeck, P. G.; Chen, L.-Y.; Makel, D. B.; Liu, C. C.; Wu, Q. H.; Knight, D.; Nov. 1998; In English; 34th Propulsion, 12-15 Jul. 1998, Cleveland, OH, USA Contract(s)/Grant(s): RTOP 254-02-0A Report No.(s): NASA/TM-1998-208817; E-11419; NAS 1.15:208817; AIAA Paper 98-3614; No Copyright; Avail: CASI; A03, Hardcopy The detection of explosive conditions in aerospace propulsion applications is important for safety and economic reasons. Microfabricated hydrogen, oxygen, and hydrocarbon sensors as well as the accompanying hardware and software are being, developed for a range of aerospace safety applications. The development of these sensors is being done using MEMS (Micro ElectroMechanical Systems) based technology and SiC-based semiconductor technology. The hardware and software allows control and interrocation of each sensor head and reduces accompanying cabling through multiplexing. These systems are being, applied on the X-33 and on an upcoming STS-95 Shuttle mission. A number of commercial applications are also being pursued. It is concluded that this MEMS-based technology has signiﬁcant potential to reduce costs and increase safety in a variety of aerospace applications. Author Gas Detectors; Hazards; Computer Programs; Hardware; Aerospace Safety

19990008538 NASA Lewis Research Center, Cleveland, OH USA Real-Time Sensor Validation, Signal Reconstruction, and Feature Detection for an RLV Propulsion Testbed Jankovsky, Amy L.; Fulton, Christopher E.; Binder, Michael P.; Maul, William A., III; Meyer, Claudia M.; Oct. 1998; In English Contract(s)/Grant(s): RTOP 242-33-01 Report No.(s): NASA/TM-1998-208807; NAS 1.15:208807; E-11392; No Copyright; Avail: CASI; A03, Hardcopy A real-time system for validating sensor health has been developed in support of the reusable launch vehicle program. This system was designed for use in a propulsion testbed as part of an overall effort to improve the safety, diagnostic capability, and cost of operation of the testbed. The sensor validation system was designed and developed at the NASA Lewis Research

98 Center and integrated into a propulsion checkout and control system as part of an industry-NASA partnership, led by Rockwell International for the Marshall Space Flight Center. The system includes modules for sensor validation, signal reconstruction, and feature detection and was designed to maximize portability to other applications. Review of test data from initial integration testing veriﬁed real-time operation and showed the system to perform correctly on both hard and soft sensor failure test cases. This paper discusses the design of the sensor validation and supporting modules developed at LeRC and reviews results obtained from initial test cases. Author Systems Health Monitoring; Real Time Operation; Signal Processing; Checkout; Expert Systems; Sensors

19990008476 Computer Sciences Corp., Huntsville, AL USA A Compendium of Wind Statistics and Models for the NASA Space Shuttle and Other Aerospace Vehicle Programs Smith, O. E.; Adelfang, S. I.; Oct. 1998; In English Contract(s)/Grant(s): NAS8-60000 Report No.(s): NASA/CR-1998-208859; NAS 1.26:208859; M-899; No Copyright; Avail: CASI; A06, Hardcopy The wind profile with all of its variations with respect to altitude has been, is now, and will continue to be important for aerospace vehicle design and operations. Wind profile databases and models are used for the vehicle ascent flight design for structural wind loading, flight control systems, performance analysis, and launch operations. This report presents the evolution of wind statistics and wind models from the empirical scalar wind profile model established for the Saturn Program through the development of the vector wind profile model used for the Space Shuttle design to the variations of this wind modeling concept for the X-33 program. Because wind is a vector quantity, the vector wind models use the rigorous mathematical probability properties of the multivariate normal probability distribution. When the vehicle ascent steering commands (ascent guidance) are wind biased to the wind profile measured on the day-of-launch, ascent structural wind loads are reduced and launch probability is increased. This wind load alleviation technique is recommended in the initial phase of vehicle development. The vehicle must fly through the largest load allowable versus altitude to achieve its mission. The Gumbel extreme value probability distribution is used to obtain the probability of exceeding (or not exceeding) the load allowable. The time conditional probability function is derived from the Gumbel bivariate extreme value distribution. This time conditional function is used for calculation of wind loads persistence increments using 3.5-hour Jimsphere wind pairs. These increments are used to protect the commit-to-launch decision. Other topics presented include the Shuttle Shuttle load-response to smoothed wind profiles, a new gust model, and advancements in wind profile measuring systems. From the lessons learned and knowledge gained from past vehicle programs, the development of future launch vehicles can be accelerated. However, new vehicle programs by their very nature will require specialized support for new databases and analyses for wind, atmospheric parameters (pressure, temperature, and density versus altitude), and weather. It is for this reason that project managers are encouraged to collaborate with natural environment specialists early in the conceptual design phase. Such action will give the lead time necessary to meet the natural environment design and operational requirements, and thus, reduce development costs. Author Wind Profiles; Wind Effects; Mathematical Models; Wind Velocity Measurement; Gust Loads; Climbing Flight

19980236873 NASA Dryden Flight Research Center, Edwards, CA USA The X-33 Extended Flight Test Range Mackall, Dale A.; Sakahara, Robert; Kremer, Steven E.; Oct. 1998; In English; 39th International Telemetry Conference, 26-29 Oct. 1980, San Diego, CA, USA Contract(s)/Grant(s): RTOP 242-33-02-00-23 Report No.(s): NASA/TM-1998-206557; NAS 1.15:206557; H-2261; No Copyright; Avail: CASI; A03, Hardcopy Development of an extended test range, with range instrumentation providing continuous vehicle communications, is required to ﬂight-test the X-33, a scaled version of a reusable launch vehicle. The extended test range provides vehicle communications coverage from California to landing at Montana or Utah. This paper provides an overview of the approaches used to meet X-33 program requirements, including using multiple ground stations, and methods to reduce problems caused by reentry plasma radio frequency blackout. The advances used to develop the extended test range show other hypersonic and access-to-space programs can beneﬁt from the development of the extended test range. Author X-33 Reusable Launch Vehicle; Reusable Launch Vehicles; Hypersonics; Flight Tests; Recoverable Launch Vehicles; Plasma Frequencies; Communication

99 19980236871 NASA Ames Research Center, Moffett Field, CA USA, Boeing Co., Long Beach, CA USA Results of Two-Stage Light-Gas Gun Development Efforts and Hypervelocity Impact Tests of Advanced Thermal Protection Materials Cornelison, C. J.; Watts, Eric T.; Oct. 1998; In English; Original contains color illustrations Contract(s)/Grant(s): RTOP 242-80-01 Report No.(s): NASA/TM-1998-112234; NAs 1.15:112234; A-9811760; No Copyright; Avail: CASI; A03, Hardcopy Gun development efforts to increase the launching capabilities of the NASA Ames 0.5-inch two-stage light-gas gun have been investigated. A gun performance simulation code was used to guide initial parametric variations and hardware modifications, in order to increase the projectile impact velocity capability to 8 km/s, while maintaining acceptable levels of gun barrel erosion and gun component stresses. Concurrent with this facility development effort, a hypervelocity impact testing series in support of the X-33/RLV program was performed in collaboration with Rockwell International. Specifically, advanced thermal protection system materials were impacted with aluminum spheres to simulate impacts with on-orbit space debris. Materials tested included AETB-8, AETB-12, AETB-20, and SIRCA-25 tiles, tailorable advanced blanket insulation (TABI), and high temperature AFRSI (HTA). The ballistic limit for several Thermal Protection System (TPS) configurations was investigated to determine particle sizes which cause threshold TPS/structure penetration. Crater depth in tiles was measured as a function of impact particle size. The relationship between coating type and crater morphology was also explored. Data obtained during this test series was used to perform a preliminary analysis of the risks to a typical orbital vehicle from the meteoroid and space debris environment. Author Hypervelocity Impact; Thermal Protection; Space Debris; Meteoroid Concentration; Impact Velocity; Impact Tests; Computerized Simulation

19980236696 NASA Dryden Flight Research Center, Edwards, CA USA X-33 Integrated Test Facility Extended Range Simulation Sharma, Ashley; Oct. 1998; In English; 34th International Telemetry Conference, 26-29 Oct. 1998, San Diego, CA, USA Contract(s)/Grant(s): RTOP 242-33-02-00-23 Report No.(s): NASA/TM-1998-206564; NAS 1.15:206564; H-2273; No Copyright; Avail: CASI; A03, Hardcopy In support of the X-33 single-stage-to-orbit program, NASA Dryden Flight Research Center was selected to provide continuous range communications of the X-33 vehicle from launch at Edwards Air Force Base, California, through landing at Malmstrom Air Force Base Montana, or at Michael Army Air Field, Utah. An extensive real-time range simulation capability is being developed to ensure successful communications with the autonomous X-33 vehicle. This paper provides an overview of various levels of simulation, integration, and test being developed to support the X-33 extended range subsystems. These subsystems include the ﬂight termination system, L-band command uplink subsystem, and S-band telemetry downlink subsystem. Author X-33 Reusable Launch Vehicle; Ground Support Systems; Military Air Facilities; Test Facilities; Uplinking; Radio Telemetry; Downlinking; Telecommunication

19980223961 NASA Dryden Flight Research Center, Edwards, CA USA Flight Testing the Linear Aerospike SR-71 Experiment (LASRE) Corda, Stephen; Neal, Bradford A.; Moes, Timothy R.; Cox, Timothy H.; Monaghan, Richard C.; Voelker, Leonard S.; Corpening, Griffin P.; Larson, Richard R.; Powers, Bruce G.; Sep. 1998; In English; 30th, 15-17 Sep. 1998, Reno, NV, USA Contract(s)/Grant(s): RTOP 242-33-02-00-23 Report No.(s): NASA/TM-1998-206567; H-2280; NAS 1.15:206567; No Copyright; Avail: CASI; A03, Hardcopy The design of the next generation of space access vehicles has led to a unique flight test that blends the space and flight research worlds. The new space vehicle designs, such as the X-33 vehicle and Reusable Launch Vehicle (RLV), are powered by linear aerospike rocket engines. Conceived of in the 1960’s, these aerospike engines have yet to be flown, and many questions remain regarding aerospike engine performance and efficiency in flight. To provide some of these data before flying on the X-33 vehicle and the RLV, a spacecraft rocket engine has been flight-tested atop the NASA SR-71 aircraft as the Linear Aerospike SR-71 Experiment (LASRE). A 20 percent-scale, semispan model of the X-33 vehicle, the aerospike engine, and all the required fuel and oxidizer tanks and propellant feed systems have been mounted atop the SR-71 airplane for this experiment. A major technical objective of the LASRE flight test is to obtain installed-engine performance flight data for comparison to wind-tunnel results and for the development of computational fluid dynamics-based design methodologies. The ultimate goal of firing the aerospike rocket engine in flight is still forthcoming. An extensive design and development phase

100 of the experiment hardware has been completed, including approximately 40 ground tests. Five ﬂights of the LASRE and ﬁring the rocket engine using inert liquid nitrogen and helium in place of liquid oxygen and hydrogen have been successfully completed. Author SR-71 Aircraft; X-33 Reusable Launch Vehicle; Reusable Launch Vehicles; Flight Tests; Computational Fluid Dynamics; Aerospike Engines; Aerodynamic Characteristics

19980218686 NASA Marshall Space Flight Center, Huntsville, AL USA Reusable Rocket Engine Operability Modeling and Analysis Christenson, R. L.; Komar, D. R.; Jul. 1998; In English Report No.(s): NASA/TP-1998-208530; M-879; NAS 1.60:208530; No Copyright; Avail: CASI; A05, Hardcopy This paper describes the methodology, model, input data, and analysis results of a reusable launch vehicle engine operability study conducted with the goal of supporting design from an operations perspective. Paralleling performance analyses in schedule and method, this requires the use of metrics in a validated operations model useful for design, sensitivity, and trade studies. Operations analysis in this view is one of several design functions. An operations concept was developed given an engine concept and the predicted operations and maintenance processes incorporated into simulation models. Historical operations data at a level of detail suitable to model objectives were collected, analyzed, and formatted for use with the models, the simulations were run, and results collected and presented. The input data used included scheduled and unscheduled timeline and resource information collected into a Space Transportation System (STS) Space Shuttle Main Engine (SSME) historical launch operations database. Results reﬂect upon the importance not only of reliable hardware but upon operations and corrective maintenance process improvements. Author Reusable Rocket Engines; Models; Procedures; Data Bases; Design Analysis

19980218658 California Univ., Santa Barbara, CA USA An Investigation of the Effect of Surface Impurities on the Adsorption Kinetics of Hydrogen Chemisorbed onto Iron Shanabarger, M. R.; 1997; In English Contract(s)/Grant(s): NCC2-63 Report No.(s): NASA/CR-1998-208255; NAS 1.26:208255; No Copyright; Avail: CASI; A03, Hardcopy The original goal of this program was to investigate the effect surface impurities have on the heterogeneous kinetic processes of those molecular species which produce gaseous hydrogen degradation of the mechanical properties of metallic structural materials. However, shortly after the initiation of the original program, the program’s NASA Technical Monitor, Dr. Howard Nelson, requested that the effort supported by this Co-operative Agreement be redirected to study more pressing materials issues associated to the development of the National Aero-Space Plane (NASP). The results of these efforts are outlined in this report. Detailed discussions of speciﬁc work, including experimental techniques and procedures, will be found in the publications listed with the subsection discussing that speciﬁc work as well and in Section 5. No inventions were generated or disclosed within this Agreement. Author Research; Surface Defects; Impurities; Adsorption; Hydrogen; Iron; Reaction Kinetics

19980217098 NASA Dryden Flight Research Center, Edwards, CA USA Flight Stability and Control and Performance Results from the Linear Aerospike SR-71 Experiment (LASRE) Moes, Timothy R.; Cobleigh, Brent R.; Cox, Timothy H.; Conners, Timothy R.; Iliff, Kenneth W.; Powers, Bruce G.; Aug. 1998; In English; Atmosphere Flight Mechanics, 10-12 Aug. 1998, Boston, MA, USA Contract(s)/Grant(s): RTOP 244-33-02 Report No.(s): NASA/TM-1998-206565; H-2276; NAS 1.15:206565; No Copyright; Avail: CASI; A03, Hardcopy The Linear Aerospike SR-71 Experiment (LASRE) is presently being conducted to test a 20-percent-scale version of the Linear Aerospike rocket engine. This rocket engine has been chosen to power the X-33 Single Stage to Orbit Technology Demonstrator Vehicle. The rocket engine was integrated into a lifting body configuration and mounted to the upper surface of an SR-71 aircraft. This paper presents stability and control results and performance results from the envelope expansion flight tests of the LASRE configuration up to Mach 1.8 and compares the results with wind tunnel predictions. Longitudinal stability and elevator control effectiveness were well-predicted from wind tunnel tests. Zero-lift pitching moment was mispredicted transonically. Directional stability, dihedral stability, and rudder effectiveness were overpredicted. The SR-71

101 handling qualities were never significantly impacted as a result of the missed predictions. Performance results confirmed the large amount of wind-tunnel-predicted transonic drag for the LASRE configuration. This drag increase made the performance of the vehicle so poor that acceleration through transonic Mach numbers could not be achieved on a hot day without depleting the available fuel. Author Flight Tests; Controllability; Aerospike Engines; Aerodynamic Stability; SR-71 Aircraft; Wind Tunnel Tests; Aircraft Performance

19980215576 NASA Langley Research Center, Hampton, VA USA Thrust Enhancement in Hypervelocity Nozzles by Chemical Catalysis Singh, D. J.; Carpenter, Mark H.; Drummond, J. P.; Journal of Propulsion and Power; Aug. 1997; Volume 13, No. 4; In English Report No.(s): NASA/TM-1997-208084; NAS 1.15:208084; No Copyright; Avail: CASI; A01, Hardcopy In the hypersonic flight regime, the air-breathing supersonic combustion ramjet (scramjet) has been shown to be a viable propulsion system. The current designs of scramjet engines provide performance benefits only up to a Mach number of 14. Performance losses increase rapidly as the Mach number increases. To extend the applicability of scram’jets beyond Mach 14, research is being conducted in the area of inlet and wave drag reduction, skin-friction and heat-transfer reduction, nozzle loss minimization, low-loss mixing, and combustion enhancement. For high Mach number applications, hydrogen is the obvious fuel choice because of its high energy content per unit mass in comparison with conventional fuels. These flight conditions require engines to operate at supersonic internal velocities, high combustor temperatures, and low static pressures. The high static temperature condition enhances the production of radicals such as H and OH, and the low-pressure condition slows the reaction rates, particularly the recombination reactions. High-temperature and low-pressure constraints, in combination with a small residence time, result in a radical-rich exhaust gas mixture exiting the combustor. At high Mach number conditions (due to low residence time), H and OH do not have enough time to recombine ; thus, a significant amount of energy is lost as these high-energy free radical are exhausted. The objective of the present study is to conduct a flowfield analysis for a typical nozzle geometry for NASP-type vehicle to assess for thrust enhancement in hypervelocity nozzles by substituting small amount of phosphine for hydrogen. Derived from text Hypervelocity Flow; Nozzle Geometry; Catalysis; Flow Distribution; Hypersonic Flight; Intake Systems; Thrust

19980210501 NASA Dryden Flight Research Center, Edwards, CA USA Linear Aerospike SR-71 Experiment (LASRE): Aerospace Propulsion Hazard Mitigation Systems Mizukami, Masashi; Corpening, Griffin P.; Ray, Ronald J.; Hass, Neal; Ennix, Kimberly A.; Lazaroff, Scott M.; Jul. 1998; In English; 34th Joint Propulsion, 13-15 Jul. 1998, Cleveland, OH, USA Contract(s)/Grant(s): RTOP 242-33-02-00-33 Report No.(s): NASA/TM-1998-206561; H-2268; NAS 1.15:206561; No Copyright; Avail: CASI; A01, Hardcopy A major hazard posed by the propulsion system of hypersonic and space vehicles is the possibility of fire or explosion in the vehicle environment. The hazard is mitigated by minimizing or detecting, in the vehicle environment, the three ingredients essential to producing fire: fuel, oxidizer, and an ignition source. The Linear Aerospike SR-71 Experiment (LASRE) consisted of a linear aerospike rocket engine integrated into one-half of an X-33-like lifting body shape, carried on top of an SR-71 aircraft. Gaseous hydrogen and liquid oxygen were used as propellants. Although LASRE is a one-of-a-kind experimental system, it must be rated for piloted flight, so this test presented a unique challenge. To help meet safety requirements, the following propulsion hazard mitigation systems were incorporated into the experiment: pod inert purge, oxygen sensors, a hydrogen leak detection algorithm, hydrogen sensors, fire detection and pod temperature thermocouples, water misting, and control room displays. These systems are described, and their development discussed. Analyses, ground test, and flight test results are presented, as are findings and lessons learned. Author Aerospike Engines; SR-71 Aircraft; Spacecraft Propulsion; Propulsion System Performance; Propulsion System Configurations; Liquid Oxygen; Lifting Bodies; Hypersonic Vehicles; Flight Tests

19980206153 Alabama Univ., Tuscaloosa, AL USA, Alabama Univ., Huntsville, AL USA Research Reports: 1996 NASA/ASEE Summer Faculty Fellowship Program Freeman, M., Editor; Chappell, C. R., Editor; Six, F., Editor; Karr, G. R., Editor; Oct. 1996; In English; 32nd, 28 May - 2 Aug. 1996, Huntsville, AL, USA Contract(s)/Grant(s): NGT8-52819 Report No.(s): NASA-CR-205205; NAS 1.26:205205; No Copyright; Avail: CASI; A18, Hardcopy

102 For the 32nd consecutive year, a NASA/ASEE Summer Faculty Fellowship Program was conducted at the Marshall Space Flight Center (MSFC). The program was conducted by the University of Alabama and MSFC during the period May 28, 1996 through August 2, 1996. Operated under the auspices of the American Society for Engineering Education, the MSFC program, as well as those at other NASA centers, was sponsored by the Higher Education Branch, Education Division, NASA Headquarters, Washington, D.C. The basic objectives of the programs, which are in the 33rd year of operation nationally, are (1) to further the professional knowledge of qualiﬁed engineering and science faculty members; (2) to stimulate an exchange of ideas between participants and NASA; (3) to enrich and refresh the research and teaching activities of the participants’ institutions; and (4) to contribute to the research objectives of the NASA centers. The Faculty Fellows spent 10 weeks at MSFC engaged in a research project compatible with their interests and background and worked in collaboration with a NASA/MSFC colleague. This document is a compilation of Fellows’ reports on their research during the summer of 1996. The University of Alabama presents the Co-Directors’ report on the administrative operations of the program. Further information can be obtained by contacting any of the editors. Author NASA Programs; Education; University Program; Astrionics; Materials Science; Structural Engineering; Systems Integration; Systems Analysis; Propulsion; Technology Transfer; Aerospace Sciences; Neural Nets; Systems Health Monitoring; Welding; Space Processing

19980202962 Georgia Inst. of Tech., Atlanta, GA USA Highly Reusable Space Transportation System Concept Evaluation (The Argus Launch Vehicle) Olds, John R.; Bellini, Peter X.; Apr. 30, 1998; In English Contract(s)/Grant(s): NCC8-108 Report No.(s): NASA/CR-1998-208093; NAS 1.26:208093; No Copyright; Avail: CASI; A03, Hardcopy This paper summarizes the results of a conceptual design study that was performed in support of NASA’s recent Highly Reusable Space Transportation study. The Argus concept uses a Maglifter magnetic-levitation sled launch assist system to accelerate it to a takeoff ground speed of 800 fps on its way to delivering a payload of 20,000 lb. to low earth orbit. Main propulsion is provided by two supercharged ejector rocket engines. The vehicle is autonomous and is fully reusable. A conceptual design exercise determined the vehicle gross weight to be approximately 597,250 lb. and the dry weight to be 75,500 lb. Aggressive weight and operations cost assumptions were used throughout the design process consistent with a second-generation reusable system that might be deployed in 10-15 years. Drawings, geometry, and weight of the concept are included. Preliminary development, production, and operations costs along with a business scenario assuming a price-elastic payload market are also included. A fleet of three Argus launch vehicles flying a total of 149 flights per year is shown to have a financial internal rate of return of 28%. At $169/lb., the recurring cost of Argus is shown to meet the study goal of $100/lb.-$200/lb., but optimum market price results in only a factor of two to five reduction compared to today’s launch systems. Author Space Transportation; Argus Project; Design Analysis; Reusable Launch Vehicles; Spacecraft Design

19980201356 Pennsylvania State Univ., University Park, PA USA An Experimental Study of Characteristic Combustion-Driven Flow for CFD Validation Santoro, Robert J.; Oct. 07, 1997; In English Contract(s)/Grant(s): NAS8-38862 Report No.(s): NASA/CR-97-208258; NAS 1.26:208258; NAS8-38862-2; No Copyright; Avail: CASI; A07, Hardcopy A series of uni-element rocket injector studies were completed to provide benchmark quality data needed to validate computational fluid dynamic models. A shear coaxial injector geometry was selected as the primary injector for study using gaseous hydrogen/oxygen and gaseous hydrogen/liquid oxygen propellants. Emphasis was placed on the use of non-intrusive diagnostic techniques to characterize the flowfields inside an optically-accessible rocket chamber. Measurements of the velocity and species fields were obtained using laser velocimetry and Raman spectroscopy, respectively Qualitative flame shape information was also obtained using laser-induced fluorescence excited from OH radicals and laser light scattering studies of aluminum oxide particle seeded combusting flows. The gaseous hydrogen/liquid oxygen propellant studies for the shear coaxial injector focused on breakup mechanisms associated with the liquid oxygen jet under sub-critical pressure conditions. Laser sheet illumination techniques were used to visualize the core region of the jet and a Phase Doppler Particle Analyzer was utilized for drop velocity, size and size distribution characterization. The results of these studies indicated that the shear coaxial geometry configuration was a relatively poor injector in terms of mixing. The oxygen core was observed to extend well downstream of the injector and a significant fraction of the mixing occurred in the near nozzle region where

103 measurements were not possible to obtain Detailed velocity and species measurements were obtained to allow CFD model validation and this set of benchmark data represents the most comprehensive data set available to date As an extension of the investigation, a series of gas/gas injector studies were conducted in support of the X-33 Reusable Launch Vehicle program. A Gas/Gas Injector Technology team was formed consisting of the Marshall Space Flight Center, the NASA Lewis Research Center, Rocketdyne and Penn State. Injector geometries studied under this task included shear and swirl coaxial conﬁgurations as well as an impinging jet injector. Author Computational Fluid Dynamics; Gas-Gas Interactions; Dynamic Models; Flow Distribution; Injectors; Laser Doppler Velocimeters; Laser Induced Fluorescence; Nonintrusive Measurement; Liquid Oxygen; Raman Spectroscopy; X-33 Reusable Launch Vehicle; Thrust Chambers

19980200836 Rohr Corp., Chula Vista, CA USA Thermal Management Design for the X-33 Lifting Body Bouslog, S.; Mammano, J.; Strauss, B.; 1998; In English; 3rd European Workshop on Thermal Protection Systems, 25-27 Mar. 1998, Netherlands Contract(s)/Grant(s): NCC8-115 Report No.(s): NASA/CR-1998-208247; NAS 1.26:208247; No Copyright; Avail: CASI; A03, Hardcopy The X-33 Advantage Technology Demonstrator offers a rare and exciting opportunity in Thermal Protection System development. The experimental program incorporates the latest design innovation in re-useable, low life cycle cost, and highly dependable Thermal Protection materials and constructions into both ground based and flight test vehicle validations. The unique attributes of the X-33 demonstrator for design application validation for the full scale Reusable Launch Vehicle, (RLV), are represented by both the configuration of the stand-off aeroshell, and the extreme exposures of sub-orbital hypersonic re-entry simulation. There are several challenges of producing a sub-orbital prototype demonstrator of Single Stage to Orbit/Reusable Launch Vehicle (SSTO/RLV) operations. An aggressive schedule with budgetary constraints precludes the opportunity for an extensive verification and qualification program of vehicle flight hardware. However, taking advantage of off the shelf components with proven technologies reduces some of the requirements for additional testing. The effects of scale on thermal heating rates must also be taken into account during trajectory design and analysis. Described in this document are the unique Thermal Protection System (TPS) design opportunities that are available with the lifting body configuration of the X-33. The two principal objectives for the TPS are to shield the primary airframe structure from excessive thermal loads and to provide an aerodynamic mold line surface. With the relatively benign aeroheating capability of the lifting body, an integrated stand-off aeroshell design with minimal weight and reduced procurement and operational costs is allowed. This paper summarizes the design objectives of the X-33 TPS, the flight test requirements driven configuration, and design benefits. Comparisons are made of the X-33 flight profiles and Space Shuttle Orbiter, and lifting body Reusable Launch Vehicle aerothermal environments. The X-33 TPS is based on a design to cost configuration concept. Only RLV critical technologies are verified to conform to cost and schedule restrictions. The one-off prototype vehicle configuration has evolved to minimize the tooling costs by reducing the number of unique components. Low cost approaches such as a composite/blanket leeward aeroshell and the use of Shuttle technology are implemented where applicable. The success of the X-33 will overcome the ballistic re-entry TPS mindset. The X-33 TPS is tailored to an aircraft type mission while maintaining sufficient operational margins. The flight test program for the X-33 will demonstrate that TPS for the RLV is not simply a surface insulation but rather an integrated aeroshell system. Author Thermal Protection; Temperature Control; X-33 Reusable Launch Vehicle; Design to Cost; Lifting Reentry Vehicles; Structural Design; Structural Design Criteria; Airframes; Aircraft Construction Materials

19980174935 NASA Lewis Research Center, Cleveland, OH USA Preliminary Assessment of Using Gelled and Hybrid Propellant Propulsion for VTOL/SSTO Launch Systems Palaszewski, Bryan; OLeary, Robert; Pelaccio, Dennis G.; Feb. 1998; In English; 33rd Joint Propulsion Conference, 7-9 Jul. 1997, Seattle, WA, USA Contract(s)/Grant(s): RTOP 260-98-09 Report No.(s): NASA/TM-1998-206306; NAS 1.15:206306; AIAA Paper 97-3216; E-11004; No Copyright; Avail: CASI; A03, Hardcopy A novel, reusable, Vertical-Takeoff-and-Vertical-Takeoff-and-Landing, Single-Stage-to-Orbit (VTOL/SSTO) launch system concept, named AUGMENT-SSTO, is presented in this paper to help quantify the advantages of employing gelled and hybrid propellant propulsion system options for such applications. The launch vehicle system concept considered uses a highly

104 coupled, main high performance liquid oxygen/liquid hydrogen (LO2/LH2) propulsion system, that is used only for launch, while a gelled or hybrid propellant propulsion system auxiliary propulsion system is used during final orbit insertion, major orbit maneuvering, and landing propulsive burn phases of flight. Using a gelled or hybrid propellant propulsion system for major orbit maneuver burns and landing has many advantages over conventional VTOL/SSTO concepts that use LO2/LH2 propulsion system(s) burns for all phases of flight. The applicability of three gelled propellant systems, O2/H2/Al, O2/RP-1/Al, and NTO/MMH/Al, and a state-of-the-art (SOA) hybrid propulsion system are examined in this study. Additionally, this paper addresses the applicability of a high performance gelled O2/H2 propulsion system to perform the primary, as well as the auxiliary propulsion system functions of the vehicle. Author Gelled Rocket Propellants; Cryogenic Rocket Propellants; Hypergolic Rocket Propellants; Vertical Takeoff Aircraft; Single Stage to Orbit Vehicles; Propulsion; Liquid Oxygen; Liquid Hydrogen; Hybrid Propulsion

19980174934 NASA Marshall Space Flight Center, Huntsville, AL USA The Control System for the X-33 Linear Aerospike Engine Jackson, Jerry E.; Espenschied, Erich; Klop, Jeffrey; 1998; In English Contract(s)/Grant(s): NCC8-115 Report No.(s): NASA/CR-1998-207923; NAS 1.26:207923; No Copyright; Avail: CASI; A03, Hardcopy The linear aerospike engine is being developed for single-stage -to-orbit (SSTO) applications. The primary advantages of a linear aerospike engine over a conventional bell nozzle engine include altitude compensation, which provides enhanced performance, and lower vehicle weight resulting from the integration of the engine into the vehicle structure. A feature of this integration is the ability to provide thrust vector control (TVC) by differential throttling of the engine combustion elements, rather than the more conventional approach of gimballing the entire engine. An analysis of the X-33 ﬂight trajectories has shown that it is necessary to provide +/- 15% roll, pitch and yaw TVC authority with an optional capability of +/- 30% pitch at select times during the mission. The TVC performance requirements for X-33 engine became a major driver in the design of the engine control system. The thrust level of the X-33 engine as well as the amount of TVC are managed by a control system which consists of electronic, instrumentation, propellant valves, electro-mechanical actuators, spark igniters, and harnesses. The engine control system is responsible for the thrust control, mixture ratio control, thrust vector control, engine health monitoring, and communication to the vehicle during all operational modes of the engine (checkout, pre-start, start, main-stage, shutdown and post shutdown). The methodology for thrust vector control, the health monitoring approach which includes failure detection, isolation, and response, and the basic control system design are the topic of this paper. As an additional point of interest a brief description of the X-33 engine system will be included in this paper. Author X-33 Reusable Launch Vehicle; Control Systems Design; Aerospike Engines; Engine Control; Single Stage to Orbit Vehicles; Throttling; Thrust Control; Thrust Vector Control

19980174930 Lockheed Martin Corp., Palmdale, CA USA, Allied-Signal Aerospace Co., Torrance, CA USA, Rockwell International Corp., Canoga Park, CA USA, Goodrich (B. F.) Aerospace, Chula Vista, CA USA, Sverdrup Technology, Inc., Saint Louis, MO USA X-33, Phase 2 1998; In English; Original contains color illustrations Contract(s)/Grant(s): NCC8-115 Report No.(s): NASA/CR-1998-208192; NAS 1.26:208192; No Copyright; Avail: CASI; A07, Hardcopy In response to the Cooperative Agreement, Lockheed Martin Skunk Works has compiled an Annual Performance Report of the X-33/RLV Program. This report consists of individual reports from all industry team members, as well as NASA team centers. The ﬁrst milestone was hand delivered to NASA MSFC. The second year has been one of signiﬁcant accomplishment in which team members have demonstrated their ability to meet vital benchmarks while continuing on the technical adventure of the 20th century. Author (revised) X-33 Reusable Launch Vehicle; Reusable Spacecraft

19980169231 NASA Dryden Flight Research Center, Edwards, CA USA Wingless Flight: The Lifting Body Story Reed, R. Dale; Lister, Darlene, Editor; Huntley, J. D., Editor; 1997; In English Report No.(s): NASA/SP-4220; NAS 1.21:4220; No Copyright; Avail: CASI; A12, Hardcopy

105 Wingless Flight tells the story of the most unusual flying machines ever flown, the lifting bodies. It is my story about my friends and colleagues who committed a significant part of their lives in the 1960s and 1970s to prove that the concept was a viable one for use in spacecraft of the future. This story, filled with drama and adventure, is about the twelve-year period from 1963 to 1975 in which eight different lifting-body configurations flew. It is appropriate for me to write the story, since I was the engineer who first presented the idea of flight-testing the concept to others at the NASA Flight Research Center. Over those twelve years, I experienced the story as it unfolded day by day at that remote NASA facility northeast of los Angeles in the bleak Mojave Desert. Benefits from this effort immediately influenced the design and operational concepts of the winged NASA Shuttle Orbiter. However, the full benefits would not be realized until the 1990s when new spacecraft such as the X-33 and X-38 would fully employ the lifting-body concept. A lifting body is basically a wingless vehicle that flies due to the lift generated by the shape of its fuselage. Although both a lifting reentry vehicle and a ballistic capsule had been considered as options during the early stages of NASA’s space program, NASA initially opted to go with the capsule. A number of individuals were not content to close the book on the lifting-body concept. Researchers including Alfred Eggers at the NASA Ames Research Center conducted early wind-tunnel experiments, finding that half of a rounded nose-cone shape that was flat on top and rounded on the bottom could generate a lift-to-drag ratio of about 1.5 to 1. Eggers’ preliminary design sketch later resembled the basic M2 lifting-body design. At the NASA Langley Research Center, other researchers toyed with their own lifting-body shapes. Meanwhile, some of us aircraft-oriented researchers at the, NASA Flight Research Center at Edwards Air Force Base (AFB) in California were experiencing our own fascination with the lifting-body concept. A model-aircraft builder and private pilot on my own time, I found the lifting-body idea intriguing. I built a model based on Eggers’ design, tested it repeatedly, made modifications in its control and balance characteristics along the way, then eventually presented the concept to others at the Center, using a film of its flights that my wife, Donna and I had made with our 8-mm home camera. Author Aircraft Models; Flight Tests; Shapes; Nose Cones; Lifting Bodies; Wings

19980151078 NASA Lewis Research Center, Cleveland, OH USA Neural Network and Regression Approximations in High Speed Civil Transport Aircraft Design Optimization Patniak, Surya N.; Guptill, James D.; Hopkins, Dale A.; Lavelle, Thomas M.; Apr. 1998; In English Contract(s)/Grant(s): RTOP 523-22-13 Report No.(s): NASA/TM-1998-206316; NAS 1.15:206316; E-10872; No Copyright; Avail: CASI; A03, Hardcopy Nonlinear mathematical-programming-based design optimization can be an elegant method. However, the calculations required to generate the merit function, constraints, and their gradients, which are frequently required, can make the process computational intensive. The computational burden can be greatly reduced by using approximating analyzers derived from an original analyzer utilizing neural networks and linear regression methods. The experience gained from using both of these approximation methods in the design optimization of a high speed civil transport aircraft is the subject of this paper. The Langley Research Center’s Flight Optimization System was selected for the aircraft analysis. This software was exercised to generate a set of training data with which a neural network and a regression method were trained, thereby producing the two approximating analyzers. The derived analyzers were coupled to the Lewis Research Center’s CometBoards test bed to provide the optimization capability. With the combined software, both approximation methods were examined for use in aircraft design optimization, and both performed satisfactorily. The CPU time for solution of the problem, which had been measured in hours, was reduced to minutes with the neural network approximation and to seconds with the regression method. Instability encountered in the aircraft analysis software at certain design points was also eliminated. On the other hand, there were costs and difficulties associated with training the approximating analyzers. The CPU time required to generate the input-output pairs and to train the approximating analyzers was seven times that required for solution of the problem. Author Nonlinear Programming; Neural Nets; Aircraft Design; National Aerospace Plane Program; Civil Aviation; Supersonic Transports

19980151073 Alabama Univ., Huntsville, AL USA Affordable In-Space Transportation, Phase 2, An Advanced Concepts Project Nov. 26, 1996; In English; Technical Interchange Meeting, 16-17 Oct. 1996, Huntsville, AL, USA Contract(s)/Grant(s): NAS8-38609 Report No.(s): NASA/CR-97-205203; NAS 1.26:205203; UAH-5-25620; No Copyright; Avail: CASI; A04, Hardcopy The Affordable In-Space Transportation (AIST) program was established by the NASA Office of Space Access to improve transportation and lower the costs from Low Earth Orbit (LEO) to Geostationary Earth Orbit (GEO) and beyond (to Lunar orbit, Mars orbit, inner solar system missions, and return to LEO). A goal was established to identify and develop radically

106 innovative concepts for new upper stages for Reusable Launch Vehicles (RLV) and Highly Reusable Space Transportation (HRST) systems. New architectures and technologies are being identiﬁed which have the potential to meet a cost goal of $1,000 to $2,000 per pound for transportation to GEO and beyond for overall mission cost (including the cost to LEO). A Technical Interchange Meeting (ITM) was held on October 16 and 17, 1996 in Huntsville, Alabama to review previous studies, present advanced concepts and review technologies that could be used to meet the stated goals. The TIM was managed by NASA-Mar-shaU Space Flight Center (MSFC) Advanced Concepts Office with Mr. Alan Adams providing TIM coordination. Mr. John C. Manidns of NASA Headquarters provided overall sponsorship. The University of Alabama in Huntsville (UAH) Propulsion Research Center hosted the TM at the UAH Research Center. Dr. Clark Hawk, Center Director, was the principal investigator. Technical support was provided by Christensen Associates. Approximately 70 attendees were present at the meeting. This Executive Summary provides a record of the key discussions and results of the TIM in a summary format. It incorporates the response to the following basic issues of the TPA, which addressed the following questions: 1. What are the cost drivers and how can they be reduced? 2. What are the operational issues and their impact on cost? What is the current Technology Readiness Level (TRL) and what will it take to reach TRL 6? 4. What are the key enabling technologies and sequence for their accomplishment? 5. What is the proposed implementation time frame? See Appendix A for the TIM Agenda and Appendix C for the AIST Program Terms of Reference. Author Space Transportation System; Cost Reduction; Coordination; Reusable Launch Vehicles

19980107885 NASA Langley Research Center, Hampton, VA USA Thermal Structures Technology Development for Reusable Launch Vehicle Cryogenic Propellant Tanks Johnson, Theodore F.; Natividad, Roderick; Rivers, H. Kevin; Smith, Russell; 1998; In English; 3rd Space Technology and Application International Forum: Conference on Next Generation Launch Systems, 25-29 Jan. 1998, Albuquerque, NM, USA Report No.(s): NASA/TM-1998-207375; NAS 1.15:207375; No Copyright; Avail: CASI; A03, Hardcopy Analytical and experimental studies conducted at the NASA Langley Research Center for investigating integrated cryogenic propellant tank systems for a Reusable Launch Vehicle are described. The cryogenic tanks are investigated as an integrated tank system. An integrated tank system includes the tank wall, cryogenic insulation, Thermal Protection System (TPS) attachment sub-structure, and TPS. Analysis codes are used to size the thicknesses of cryogenic insulation and TPS insulation for thermal loads, and to predict tank buckling strengths at various ring frame spacings. The unique test facilities developed for the testing of cryogenic tank components are described. Testing at cryogenic and high-temperatures veriﬁes the integrity of materials, design concepts, manufacturing processes, and thermal/structural analyses. Test specimens ranging from the element level to the subcomponent level are subjected to projected vehicle operational mechanical loads and temperatures. The analytical and experimental studies described in this paper provide a portion of the basic information required for the development of light-weight reusable cryogenic propellant tanks. Author Reusable Launch Vehicles; Cryogenic Fluid Storage; Structural Engineering; Structural Analysis; Thermal Protection; Propellant Tanks

19980073196 Lockheed Sanders, Inc., Nashua, NH USA X-33/RLV System Health Management/ Vehicle Health Management Garbos, Raymond J.; Mouyos, William; 1998; In English; Structures, Structural Dynnamics, and Materials Conference and Adaptive Structures Forum, 20-23 Apr. 1998, Long Beach, CA, USA Contract(s)/Grant(s): NCC8-115 Report No.(s): NASA/CR-1998-208136; NAS 1.26:208136; AIAA Paper 98-1928; No Copyright; Avail: CASI; A02, Hardcopy To reduce operations cost, the RLV must include the following elements: highly reliable, robust subsystems designed for simple repair access with a simplified servicing infrastructure and incorporating expedited decision making about faults and anomalies. A key component for the Single Stage to Orbit (SSTO) RLV System used to meet these objectives is System Health Management (SHM). SHM deals with the vehicle component- Vehicle Health Management (VHM), the ground processing associated with the fleet (GVHM) and the Ground Infrastructure Health Management (GIHM). The objective is to provide an automated collection and paperless health decision, maintenance and logistics system. Many critical technologies are necessary to make the SHM (and more specifically VHM) practical, reliable and cost effective. Sanders is leading the design, development and integration of the SHM system for RLV and X-33 SHM (a sub-scale, sub-orbit Advanced Technology

107 Demonstrator). This paper will present the X-33 SHM design which forms the baseline for RLV SHM. This paper will also discuss other applications of these technologies. Author X-33 Reusable Launch Vehicle; Decision Making; Smart Structures; Cost Reduction; Management Systems

19980069722 Lockheed Sanders, Inc., Nashua, NH USA Reusable Cryogenic Tank VHM Using Fiber Optic Distributed Sensing Technology Bodan-Sanders, Patricia; Bouvier, Carl; 1998; In English; Structures, Structural Dynamics and Materials Conference and Adaptive Structures Forum, 20-23 Apr. 1998, Long Beach, CA, USA Contract(s)/Grant(s): NCC8-115 Report No.(s): NASA/CR-1998-208132; NAS 1.26:208132; AIAA Paper 98-1929; No Copyright; Avail: CASI; A02, Hardcopy The reusable oxygen and hydrogen tanks are key systems for both the X-33 (sub-scale, sub-orbital technology demonstrator) and the commercial Reusable Launch Vehicle (RLV). The backbone of the X-33 Reusable Cryogenic Tank Vehicle Health Management (VHM) system lies in the optical network of distributed strain temperature and hydrogen sensors. This network of ﬁber sensors will create a global strain and temperature map for monitoring the health of the tank structure, cryogenic insulation, and Thermal Protection System. Lockheed Martin (Sanders and LMMSS) and NASA Langley have developed this sensor technology for the X-33 and have addressed several technical issues such as ﬁber bonding and laser performance in this harsh environment. Author X-33 Reusable Launch Vehicle; Cryogenic Fluid Storage; Fuel Tanks; Temperature Sensors; Fluid Management; Thermal Insulation; Fiber Optics; Cryogenic Rocket Propellants

19980055201 Boeing Commercial Airplane Co., Seattle, WA USA Aircraft Emission Scenarios Projected in Year 2015 for the NASA Technology Concept Aircraft (TCA) High Speed Civil Transport Baughcum, Steven L.; Henderson, Stephen C.; Mar. 1998; In English Contract(s)/Grant(s): NAS1-20220; RTOP 537-09-23-02 Report No.(s): NASA/CR-1998-207635; NAS 1.26:207635; No Copyright; Avail: CASI; A03, Hardcopy This report describes the development of a three-dimensional database of aircraft fuel burn and emissions (fuel burned, NOx, CO, and hydrocarbons) from projected fleets of high speed civil transports (HSCTs) on a universal airline network. Inventories for 500 and 1000 HSCT fleets, as well as the concurrent subsonic fleets, were calculated. The HSCT scenarios are calculated using the NASA technology concept airplane (TCA) and update an earlier report. These emissions inventories are available for use by atmospheric scientists conducting the Atmospheric Effects of Stratospheric Aircraft (AESA) modeling studies. Fuel burned and emissions of nitrogen oxides (NOx as NO2), carbon monoxide, and hydrocarbons have been calculated on a 1 degree latitude x 1 degree longitude x 1 kilometer pressure altitude grid and delivered to NASA as electronic files. Author Aircraft Fuels; Emission; Supersonic Transports; Atmospheric Effects; Airline Operations; Ozonosphere; X-30 Vehicle

19980055123 NASA Dryden Flight Research Center, Edwards, CA USA Development of the X-33 Aerodynamic Uncertainty Model Cobleigh, Brent R.; Apr. 1998; In English Contract(s)/Grant(s): RTOP 242-33-02-00-23 Report No.(s): NASA/TP-1998-206544; NAS 1.60:206544; H-2221; No Copyright; Avail: CASI; A03, Hardcopy An aerodynamic uncertainty model for the X-33 single-stage-to-orbit demonstrator aircraft has been developed at NASA Dryden Flight Research Center. The model is based on comparisons of historical flight test estimates to preflight wind-tunnel and analysis code predictions of vehicle aerodynamics documented during six lifting-body aircraft and the Space Shuttle Orbiter flight programs. The lifting-body and Orbiter data were used to define an appropriate uncertainty magnitude in the subsonic and supersonic flight regions, and the Orbiter data were used to extend the database to hypersonic Mach numbers. The uncertainty data consist of increments or percentage variations in the important aerodynamic coefficients and derivatives as a function of Mach number along a nominal trajectory. The uncertainty models will be used to perform linear analysis of the X-33 flight control system and Monte Carlo mission simulation studies. Because the X-33 aerodynamic uncertainty model

108 was developed exclusively using historical data rather than X-33 speciﬁc characteristics, the model may be useful for other lifting-body studies. Author Aerodynamic Characteristics; X-33 Reusable Launch Vehicle; Monte Carlo Method; Lifting Bodies; Hypersonics; Flight Control; Space Shuttle Orbiters

19980053187 NASA Marshall Space Flight Center, Huntsville, AL USA Analysis of Aerospike Plume Induced Base-Heating Environment Wang, Ten-See; 1998; In English; 7th Joint Thermophysics and Heat Transfer Conference, 15-18 Jun. 1998, Albuquerque, NM, USA Report No.(s): NASA/TM-1998-208139; NAS 1.15:208139; AIAA Paper 98-2469; No Copyright; Avail: CASI; A02, Hardcopy Computational analysis is conducted to study the effect of an aerospike engine plume on X-33 base-heating environment during ascent flight. To properly account for the effect of forebody and aftbody flowfield such as shocks and to allow for potential plume-induced flow-separation, thermo-flowfield of trajectory points is computed. The computational methodology is based on a three-dimensional finite-difference, viscous flow, chemically reacting, pressure-base computational fluid dynamics formulation, and a three-dimensional, finite-volume, spectral-line based weighted-sum-of-gray-gases radiation absorption model computational heat transfer formulation. The predicted convective and radiative base-heat fluxes are presented. Author Aerospike Engines; Plumes; Base Heating; Boundary Layer Separation; Convective Heat Transfer; Gray Gas; Heat Transfer; Separated Flow; Three Dimensional Flow; X-33 Reusable Launch Vehicle

19980048417 NASA Marshall Space Flight Center, Huntsville, AL USA Tether Transportation System Study Bangham, M. E.; Lorenzini, E.; Vestal, L.; Mar. 1998; In English Contract(s)/Grant(s): NAS8-39400 Report No.(s): NASA/TP-1998-206959; NAS 1.60:206959; M-853; No Copyright; Avail: CASI; A05, Hardcopy The projected traffic to geostationary earth orbit (GEO) is expected to increase over the next few decades. At the same time, the cost of delivering payloads from the Earth’s surface to low earth orbit (LEO) is projected to decrease, thanks in part to the Reusable Launch Vehicle (RLV). A comparable reduction in the cost of delivering payloads from LEO to GEO is sought. The use of in-space tethers, eliminating the requirement for traditional chemical upper stages and thereby reducing the launch mass, has been identiﬁed as such an alternative. Spinning tethers are excellent kinetic energy storage devices for providing the large delta vee’s required for LEO to GEO transfer. A single-stage system for transferring payloads from LEO to GEO was proposed some years ago. The study results presented here contain the ﬁrst detailed analyses of this proposal, its extension to a two-stage system, and the likely implementation of the operational system. Author Low Earth Orbits; Geosynchronous Orbits; Tethering; Reusable Launch Vehicles; Payloads; Costs

19980047114 NASA Langley Research Center, Hampton, VA USA Hyper-X Wind Tunnel Program McClinton, C. R.; Holland, S. D.; Rock, K. E.; Engelund, W. C.; Voland, R. T.; Huebner, L. D.; Roger, R. C.; 1998; In English; Aerospace Sciences Meeting and Exhibit, 12-15 Jan. 1998, Reno, NV, USA Report No.(s): NASA/TM-1998-207317; NAS 1.15:207317; AIAA Paper 98-0553; No Copyright; Avail: CASI; A03, Hardcopy This paper provides an overview of NASA’s focused hypersonic technology program, called the Hyper-X Program. The Hyper-X Program, a joint NASA Langley and Dryden program, is designed to move hypersonic, air breathing vehicle technology from the laboratory environment to the flight environment, the last stage preceding prototype development. The Hyper-X research vehicle will provide the first ever opportunity to obtain data on an airframe integrated scramjet (supersonic combustion ramjet) propulsion system at true flight conditions and the first opportunity for flight validation of experimental wind tunnel, numerical and analytical methods used for design of these vehicles. A substantial portion of the program is experimentally based, both for database development and performance validation. The program is now concentrating on Mach 7 vehicle development, verification and validation and flight test risk reduction. This paper concentrates on the aerodynamic

109 and propulsion experimental programs. Wind tunnel testing of the flight engine and complete airframe integrated scramjet configuration flow-path is expected in 1998 and 1999, respectively, and flight test is planned for 2000. Author Supersonic Combustion Ramjet Engines; Research Vehicles; Hypersonic Vehicles; Flight Tests; Transatmospheric Vehicles

19980046639 NASA Langley Research Center, Hampton, VA USA Computational/Experimental Aeroheating Predictions for X-33 Phase 2 Vehicle Hamilton, H. Harris,II; Weilmuenster, K. James; Horvath, Thomas J.; Berry, Scott A.; 1998; In English; Aerospace Sciences Meeting and Exhibit, 12-15 Jan. 1998, Reno, NV, USA Report No.(s): NASA/TM-1998-207290; NAS 1.15:207290; AIAA Paper 98-0869; No Copyright; Avail: CASI; A03, Hardcopy Laminar and turbulent heating-rate calculations from an ‘engineering’ code and laminar calculations from a ‘benchmark’ Navier-Stokes code are compared with experimental wind-tunnel data obtained on several candidate configurations for the X-33 Phase 2 flight vehicle. The experimental data were obtained at a Mach number of 6 and a freestream Reynolds number ranging from 1 to 8 x 10(exp 6)/ft. Comparisons are presented along the windward symmetry plane and in a circumferential direction around the body at several axial stations at angles of attack from 20 to 40 deg. The experimental results include both laminar and turbulent How. For the highest angle of attack some of the measured heating data exhibited a ’non-laminar‘ behavior which caused the heating to increase above the laminar level long before ‘classical’ transition to turbulent flow was observed. This trend was not observed at the lower angles of attack. When the flow was laminar, both codes predicted the heating along the windward symmetry plane reasonably well but under-predicted the heating in the chine region. When the flow was turbulent the LATCH code accurately predicted the measured heating rates. Both codes were used to calculate heating rates over the X-33 vehicle at the peak heating point on the design trajectory and they were found to be in very good agreement over most of the vehicle windward surface. Author Computational Fluid Dynamics; Navier-Stokes Equation; Numerical Flow Visualization; Direct Numerical Simulation; X-33 Reusable Launch Vehicle; Reynolds Number; Transition Flow; Turbulent Flow; Hypersonic Boundary Layer; Hypersonic Vehicles

19980041407 Rockwell International Corp., Canoga Park, CA USA Operationally efficient propulsion system study (OEPSS) data book, Volume 10, Air Augmented Rocket Afterburning Farhangi, Shahram; Trent, Donnie, Editor; Oct. 30, 1992; In English Contract(s)/Grant(s): NAS10-11568 Report No.(s): NASA/CR-92-207458; NAS 1.26:207458; RI/RD90-149-10-VOL-10; CDR-91-099; No Copyright; Avail: CASI; A03, Hardcopy A study was directed towards assessing viability and effectiveness of an air augmented ejector/rocket. Successful thrust augmentation could potentially reduce a multi-stage vehicle to a single stage-to-orbit vehicle (SSTO) and, thereby, eliminate the associated ground support facility infrastructure and ground processing required by the eliminated stage. The results of this preliminary study indicate that an air augmented ejector/rocket propulsion system is viable. However, uncertainties resulting from simpliﬁed approach and assumptions must be resolved by further investigations. Author Thrust Augmentation; Rocket Engines; Advanced Launch System (Sts); Propulsion System Performance; Mach Number; Ejectors

19980041384 Rockwell International Corp., Canoga Park, CA USA Operationally Efficient Propulsion System Study (OEPSS): OEPSS Video Script Wong, George S.; Waldrop, Glen S.; Trent, Donnie, Editor; Sep. 30, 1992; In English Contract(s)/Grant(s): NAS10-11568 Report No.(s): NASA/CR-92-207441; NAS 1.26:207441; No Copyright; Avail: CASI; A05, Hardcopy The OEPSS video film, along with the OEPSS Databooks, provides a data base of current launch experience that will be useful for design of future expendable and reusable launch systems. The focus is on the launch processing of propulsion systems. A brief 15-minute overview of the OEPSS study results is found at the beginning of the film. The remainder of the film discusses in more detail: current ground operations at the Kennedy Space Center; typical operations issues and problems; critical operations technologies; and efficiency of booster and space propulsion systems. The impact of system architecture on

110 the launch site and its facility infrastucture is emphasized. Finally, a particularly valuable analytical tool, developed during the OEPSS study, that will provide for the ’ﬁrst time‘ a quantitative measure of operations efficiency for a propulsion system is described. Author Propulsion System Performance; Video Tapes; Spacecraft Launching

19980037670 NASA Langley Research Center, Hampton, VA USA Hypersonic Boundary-Layer Transition for X-33 Phase 2 Vehicle Thompson, Richard A.; Hamilton, Harris H., II; Berry, Scott A.; Horvath, Thomas J.; Nowak, Robert J.; 1998; In English; Aerospace Sciences Meeting and Exhibit, 12-15 Jan. 1998, Reno, NV, USA; Original contains color illustrations Report No.(s): NASA/TM-1998-207316; NAS 1.15:207316; AIAA Paper 98-0867; No Copyright; Avail: CASI; A03, Hardcopy A status review of the experimental and computational work performed to support the X-33 program in the area of hypersonic boundary-layer transition is presented. Global transition fronts are visualized using thermographic phosphor measurements. Results are used to derive transition correlations for ’smooth body‘ and discrete roughness data and a computational tool is developed to predict transition onset for X-33 using these results. The X-33 thermal protection system appears to be conservatively designed for transition effects based on these studies. Additional study is needed to address concerns related to surface waviness. A discussion of future test plans is included. Author Boundary Layer Transition; Hypersonics; Flow Distribution; Wind Tunnel Tests; Finite Volume Method; Flow Visualization

19980036345 Alabama Univ., Huntsville, AL USA Study Acoustic Emissions from Composites Walker, James; Workman,Gary; Jan. 1998; In English; Original contains color illustrations Contract(s)/Grant(s): NAS8-38609 Report No.(s): NASA/CR-1998-207483; NAS 1.26:207483; No Copyright; Avail: CASI; A06, Hardcopy The purpose of this work will be to develop techniques for monitoring the acoustic emissions from carbon epoxy composite structures at cryogenic temperatures. Performance of transducers at temperatures ranging from ambient to cryogenic and the characteristics of acoustic emission from composite structures will be studied and documented. This entire effort is directed towards characterization of structures used in NASA propulsion programs such as the X-33. Author Epoxy Resins; Acoustic Emission; Cryogenic Temperature; Composite Structures

19980020931 Albany Univ., Albany, NY USA Advanced High Brilliance X-Ray Source Gibson, Walter M.; 1998; In English Contract(s)/Grant(s): NAS8-39926 Report No.(s): NASA/CR-1998-207486; NAS 1.26:207486; No Copyright; Avail: CASI; A03, Hardcopy The possibility to dramatically increase the efficiency of laboratory based protein structure measurements through the use of polycapillary X-ray optics was investigated. This project initiated April 1, 1993 and concluded December 31, 1996 (including a no cost extension from June 31, 1996). This is a ﬁnal report of the project. The basis for the project is the ability to collect X-rays from divergent electron bombardment laboratory X-ray sources and redirect them into quasiparallel or convergent (focused) beams. For example, a 0.1 radian (approx. 6 deg) portion of a divergent beam collected by a polycapillary collimator and transformed into a quasiparallel beam of 3 millradian (0.2 deg) could give a gain of 6(exp 2)/0.2(exp 2) x T for the intensity of a diffracted beam from a crystal with a 0.2 deg diffraction width. T is the transmission efficiency of the polycapillary diffraction optic, and for T=0.5, the gain would be 36/0.04 x O.5=45. In practice, the effective collection angle will depend on the source spot size, the input focal length of the optic (usually limited by the source spot-to-window distance on the x-ray tube) and the size of the crystal relative to the output diameter of the optic. The transmission efficiency, T, depends on the characteristics (fractional open area, surface roughness, shape and channel diameter) of the polycapillary optic and is typically in the range 0.2-0.4. These effects could substantially reduce the expected efficiency gain. During the course of this study, the possibility to use a weakly focused beam (0.5 deg convergence) was suggested which could give an additional 10-20 X efficiency gain for small samples . Weakly focused beams from double focusing mirrors are frequently used for macromolecular crystallography studies. Furthermore the crystals are typically oscillated by as much as

111 2 deg during each X-ray exposure in order to increase the reciprocal space (number of crystal planes) sampled and use of a slightly convergent beam could, in principle, provide a similar sampling beneﬁt without oscillation. Although more problematic, because of complications in analyzing the diffraction patterns, it was also suggested that even more extreme beam convergence might be used to give another order of magnitude intensity gain and even smaller focused spot size which could make it possible to study smaller protein crystals than can be studied using standard laboratory based X-ray diffraction systems. This project represents the ﬁrst systematic investigation of these possibilities. As initially proposed, the contract included requirements for design, purchase, evaluation and delivery of three polycapillary lenses to the Laboratory for Structural Biology at MSFC and demonstration of such optics at MSFC for selected protein crystal diffraction applications. Author X Ray Sources; Effıciency; Proteins; Chemical Analysis

19980019978 San Jose State Univ., CA USA Development of Processing Techniques for Advanced Thermal Protection Materials Selvaduray, Guna; Cox, Michael; Srinivasan, Vijayakumar; Nov. 30, 1997; In English Contract(s)/Grant(s): NAG2-848 Report No.(s): NASA/CR-1998-207440; NAS 1.26:207440; No Copyright; Avail: CASI; A02, Hardcopy Thermal Protection Materials Branch (TPMB) has been involved in various research programs to improve the properties and structural integrity of the existing aerospace high temperature materials. Specimens from various research programs were brought into the analytical laboratory for the purpose of obtaining and reﬁning the material characterization. The analytical laboratory in TPMB has many different instruments which were utilized to determine the physical and chemical characteristics of materials. Some of the instruments that were utilized by the SJSU students are: Scanning Electron Microscopy (SEM), Energy Dispersive X-ray analysis (EDX), X-ray Diffraction Spectrometer (XRD), Fourier Transform-Infrared Spectroscopy (FTIR), Ultra Violet Spectroscopy/Visible Spectroscopy (UV/VIS), Particle Size Analyzer (PSA), and Inductively Coupled Plasma Atomic Emission Spectrometer (ICP-AES). The above mentioned analytical instruments were utilized in the material characterization process of the specimens from research programs such as: aerogel ceramics (I) and (II), X-33 Blankets, ARC-Jet specimens, QUICFIX specimens and gas permeability of lightweight ceramic ablators. In addition to analytical instruments in the analytical laboratory at TPMB, there are several on-going experiments. One particular experiment allows the measurement of permeability of ceramic ablators. From these measurements, physical characteristics of the ceramic ablators can be derived. Derived from text Thermal Protection; Infrared Spectroscopy; Refractory Materials; Aircraft Construction Materials; Ablative Materials; Infrared Spectra; Plasma Radiation; X Ray Analysis

19980019278 NASA Langley Research Center, Hampton, VA USA Computational/Experimental Aeroheating Predictions for X-33, Phase 2, Vehicle Hamilton, H. Harris, II; Weilmuenster, K. James; Horvath, Thomas J.; Berry, Scott A.; 1998; In English; 36th Aerospace Sciences Meeting and Exhibit, 12-15 Jan. 1998, Reno, NV, USA Report No.(s): NASA/TM-1998-207315; NAS 1.15:207315; AIAA Paper 98-0869; No Copyright; Avail: CASI; A03, Hardcopy Laminar and turbulent heating-rate calculations from an ’engineering‘ code and laminar calculations from a ’benchmark‘ Navier-Stokes code are compared with experimental wind-tunnel data obtained on several candidate configurations for the X-33 Phase 2 flight vehicle. The experimental data were obtained at a Mach number of 6 and a freestream Reynolds number ranging from 1 to 8 x 10(exp 6)/ft. Comparisons are presented along the windward symmetry plane and in a circumferential direction around the body at several axial stations at angles of attack from 20 to 40 deg. The experimental results include both laminar and turbulent flow. For the highest angle of attack some of the measured heating data exhibited a ’non-laminar‘ behavior which caused the heating to increase above the laminar level long before ’classical‘ transition to turbulent flow was observed. This trend was not observed at the lower angles of attack. When the flow was laminar, both codes predicted the heating along the windward symmetry plane reasonably well but under-predicted the heating in the chine region. When the flow was turbulent the LATCH code accurately predicted the measured heating rates. Both codes were used to calculate heating rates over the X-33 vehicle at the peak heating point on the design trajectory and they were found to be in very good agreement over most of the vehicle windward surface. Author Aerodynamic Heating; X-33 Reusable Launch Vehicle; Wind Tunnel Tests; Angle of Attack; Turbulent Flow; Laminar Flow; Navier-Stokes Equation; Computational Fluid Dynamics

112 19980015220 NERAC, Inc., Tolland, CT USA, NASA, Washington, DC USA National Aerospace Plane Thermal Development. (Latest Citations from the Aerospace Database) Mar. 1996; In English; Page count unavailable. Report No.(s): NASA/TM-96-206845; NAS 1.15:206845; PB96-864210; Copyright; Avail: National Technical Information Service (NTIS) The bibliography contains citations concerning thermal properties of the National Aerospace Plane (NASP). Analysis of thermal stress, and methods for determining thermal effects on the plane’s supersonic structure are discussed. The citations also review temperature extremes that the vehicle is likely to encounter. (Contains 50-250 citations and includes a subject term index and title list.) NTIS Bibliographies; Thermal Stresses; Stress Analysis; Aerospace Planes

19980010422 NASA, Washington, DC USA X-33 Development History Butrica, Andrew J.; Dec. 15, 1997; In English Contract(s)/Grant(s): NASw-97005 Report No.(s): NASA/CR-97-206438; NAS 1.26:206438; No Copyright; Avail: CASI; A02, Hardcopy The problem of dealing with various types of proprietary documents, whether from the Lockheed Martin, the Skunk Works, McDonnell Douglas, Rockwell, and other corporations extant or extinct, remains unresolved. The computerized archive ﬁnding aid has over 100 records at present. These records consist of X-33 photographs, press releases, media clippings, and the small number of X-33 project records collected to date. Author Computer Storage Devices; Computer Systems Programs; Data Processing

19980008580 NASA Dryden Flight Research Center, Edwards, CA USA Design and Calibration of the X-33 Flush Airdata Sensing (FADS) System Whitmore, Stephen A.; Cobleigh, Brent R.; Haering, Edward A.; Jan. 1998; In English; AIAA Aerospace Sciences Meeting and Exhibit, 12-15 Jan. 1998, Reno, NV, USA Contract(s)/Grant(s): RTOP 242-33-02-00-23 Report No.(s): NASA/TM-98-206540; H-2219; NAS 1.15:206540; AIAA Paper 98-0201; No Copyright; Avail: CASI; A03, Hardcopy This paper presents the design of the X-33 Flush Airdata Sensing (FADS) system. The X-33 FADS uses a matrix of pressure orifices on the vehicle nose to estimate airdata parameters. The system is designed with dual-redundant measurement hardware, which produces two independent measurement paths. Airdata parameters that correspond to the measurement path with the minimum fit error are selected as the output values. This method enables a single sensor failure to occur with minimal degrading of the system performance. The paper shows the X-33 FADS architecture, derives the estimating algorithms, and demonstrates a mathematical analysis of the FADS system stability. Preliminary aerodynamic calibrations are also presented here. The calibration parameters, the position error coefficient (epsilon), and flow correction terms for the angle of attack (delta alpha), and angle of sideslip (delta beta) are derived from wind tunnel data. Statistical accuracy of’ the calibration is evaluated by comparing the wind tunnel reference conditions to the airdata parameters estimated. This comparison is accomplished by applying the calibrated FADS algorithm to the sensed wind tunnel pressures. When the resulting accuracy estimates are compared to accuracy requirements for the X-33 airdata, the FADS system meets these requirements. Author X-33 Reusable Launch Vehicle; Data Systems; Systems Integration; Calibrating; Remote Sensing; Systems Engineering

19980003836 Alabama Univ., Huntsville, AL USA Affordable In-Space Transportation Phase 2: An Advanced Concepts Project 1996; In English; Technical Interchange Meeting, 16-17 Oct. 1996, USA Contract(s)/Grant(s): NAS8-38609 Report No.(s): NASA/CR-97-206469; NAS 1.26:206469; No Copyright; Avail: CASI; A04, Hardcopy The Affordable In-Space Transportation (AIST) program was established by the NASA Office of Space Access to improve transportation and lower the costs from Low Earth Orbit (LEO) to Geostationary Earth Orbit (GEO) and beyond (to Lunar orbit, Mars orbit, inner solar system missions, and return to LEO). A goal was established to identify and develop radically

113 innovative concepts for new upper stages for Reusable Launch Vehicles (RLV) and Highly Reusable Space Transportation (HRST) systems. New architectures and technologies are being identiﬁed which have the potential to meet a cost goal of $1,000 to $2,000 per pound for transportation to GEO and beyond for overall mission cost (including the cost to LEO). A Technical Interchange Meeting (TTM) was held on October 16 and 17, 1996 in Huntsville, Alabama to review previous studies, present advanced concepts and review technologies that could be used to meet the stated goals. The TIN4 was managed by NASA-Marshall Space Flight Center (MSFC) Advanced Concepts Office with Mr. Alan Adams providing TIM coordination. Mr. John C. Mankins of NASA Headquarters provided overall sponsorship. The University of Alabama in Huntsville (UAH) Propulsion Research Center hosted the TIM at the UAH Research Center. Dr. Clark Hawk, Center Director, was the principal investigator. Technical support was provided by Christensen Associates. Approximately 70 attendees were present at the meeting. This Executive Summary provides a record of the key discussions and results of the TIN4 in a summary for-mat. It incorporates the response to the following basic issues of the TDVL which addressed the following questions: 1. What are the cost drivers and how can they be reduced? 2. What are the operational issues and their impact on cost? 3. What is the current technology readiness level (TRL) and what will it take to reach TRL 6? 4. What are the key enabling technologies and sequence for their accomplishment? 5 . What is the proposed implementation time frame? See Appendix A for the TIM Agenda and Appendix C for the AIST Program Terms of Reference. Derived from text Space Transportation System; Low Earth Orbits; Cost Reduction; Geosynchronous Orbits; Reusable Launch Vehicles; Technology Assessment; Lunar Orbits; Costs

19980000968 NASA, Washington, DC USA National Aerospace Plane Thermal Development. (Latest citations from the Aerospace Database) Jun. 1997; In English Report No.(s): NASA/TM-97-113072; NAS 1.15:113072; PB97-860860; No Copyright; Avail: National Technical Information Service (NTIS) The bibliography contains citations concerning thermal properties of the National Aerospace Plane (NASP). Analysis of thermal stress, and methods for determining thermal effects on the plane’s supersonic structure are discussed. The citations also review temperature extremes that the vehicle is likely to encounter. NTIS National Aerospace Plane Program; Thermodynamic Properties; Aerospace Planes

19970040845 State of New Mexico, Las Cruces, NM USA Business Plan for the Southwest Regional Spaceport: Executive Summary Jul. 1997; In English; Original contains color illustrations Contract(s)/Grant(s): NAGw-5131 Report No.(s): NASA/CR-97-206023; NAS 1.26:206023; No Copyright; Avail: CASI; A03, Hardcopy A proposal for a commercial, full-service launch, tracking, and recovery complex for Reusable Launch Vehicles in New Mexico is presented. Vision, mission, business deﬁnition, competitive advantages, and business approach are formulated. Management plan and team structure are detailed, and anticipated market is described. Finance and marketing plans are presented. Financial analysis is performed. Derived from text Launching Bases; Proposals

19970040679 Space America, Inc., Huntsville, AL USA Highly Reusable Space Transportation (HRST) Baseline Concepts and Analysis: Rocket/RBCC Options, Part 2, A Comparative Study Woodcock, Gordon; Oct. 29, 1997; In English Contract(s)/Grant(s): W-7-PP-03050(1F); NASA Order H-28681-D Report No.(s): NASA/CR-97-112565; NAS 1.26:112565; No Copyright; Avail: CASI; A06, Hardcopy This study is an extension of a previous effort by the Principal Investigator to develop baseline data to support comparative analysis of Highly Reusable Space Transportation (HRST) concepts. The analyses presented herin develop baseline data bases for two two-stage-to-orbit (TSTO) concepts: (1) Assisted horizontal take-off all rocket (assisted HTOHL); and (2) Assisted vertical take-off rocket based combined cycle (RBCC). The study objectives were to: (1) Provide conﬁguration deﬁnitions and illustrations for assisted HTOHL and assisted RBCC; (2) Develop a rationalization approach and

114 compare these concepts with the HRST reference; and (3) Analyze TSTO conﬁgurations which try to maintain SSTO beneﬁts while reducing inert weight sensitivity. Author Space Transportation; Reusable Spacecraft; Single Stage to Orbit Vehicles; Data Bases; Vertical Takeoff

19970040461 Lockheed Martin Corp., Palmdale, CA USA X-33 Phase 2 1997; In English; Original contains color illustrations Contract(s)/Grant(s): NCC8-115 Report No.(s): NASA-CR-205695; NAS 1.26:205695; No Copyright; Avail: CASI; A06, Hardcopy In response to Clause 17 of the Cooperative Agreement NCC8-115, Lockheed Martin Skunk Works has compiled an Annual Performance Report of the X-33/RLV Program. This report consists of individual reports from all industry team members, as well as NASA team centers. Contract award was announced on July 2, 1996 and the ﬁrst milestone was hand delivered to NASA MSFC on July 17, 1996. The ﬁrst year has been one of growth and progress as all team members staffed up and embarked on the technical adventure of the 20th century... the ultimate goal..aSingle Stage to Orbit (SSTO) Reuseable Launch Vehicle (RLV). Derived from text X-33 Reusable Launch Vehicle; Single Stage to Orbit Vehicles; Launch Vehicles; Reusable Spacecraft

19970040167 NASA Ames Research Center, Moffett Field, CA USA Near-Optimal Re-Entry Trajectories for Reusable Launch Vehicles Chou, H.-C.; Ardema, M. D.; Bowles, J. V.; American Inst. of Aeronautics and Astronautics Journal; 1997; In English Contract(s)/Grant(s): NCC2-5217; NCC2-5069 Report No.(s): NASA/CR-97-112611; NAS 1.26:112611; AIAA Paper 97-3582; Copyright; Avail: CASI; A03, Hardcopy A near-optimal guidance law for the descent trajectory for earth orbit re-entry of a fully reusable single-stage-to-orbit pure rocket launch vehicle is derived. A methodology is developed to investigate using both bank angle and altitude as control variables and selecting parameters that maximize various performance functions. The method is based on the energy-state model of the aircraft equations of motion. The major task of this paper is to obtain optimal re-entry trajectories under a variety of performance goals: minimum time, minimum surface temperature, minimum heating, and maximum heading change; four classes of trajectories were investigated: no banking, optimal left turn banking, optimal right turn banking, and optimal bank chattering. The cost function is in general a weighted sum of all performance goals. In particular, the trade-off between minimizing heat load into the vehicle and maximizing cross range distance is investigated. The results show that the optimization methodology can be used to derive a wide variety of near-optimal trajectories. Author Reusable Launch Vehicles; Reentry Trajectories; Descent Trajectories; Equations of Motion; Heating; Time Temperature Parameter; Surface Temperature

19970037803 Pennsylvania State Univ., University Park, PA USA An Experimental Study of Characteristic Combustion-Driven Flow for CFD Validation Santoro, Robert J.; Oct. 1997; In English Contract(s)/Grant(s): NAS8-38862 Report No.(s): NASA/CR-97-205735; NAS 1.26:205735; PSU-NAS8-38862-2; No Copyright; Avail: CASI; A07, Hardcopy A series of uni-element rocket injector studies were completed to provide benchmark quality data needed to validate computational fluid dynamic models. A shear coaxial injector geometry was selected as the primary injector for study using gaseous hydrogen/oxygen and gaseous hydrogen/liquid oxygen propellants. Emphasis was placed on the use of nonintrusive diagnostic techniques to characterize the flowfields inside an optically-accessible rocket chamber. Measurements of the velocity and species fields were obtained using laser velocimetry and Raman spectroscopy, respectively. Qualitative flame shape information was also obtained using laser-induced fluorescence excited from OH radicals and laser light scattering studies of aluminum oxide particle seeded combusting flows. The gaseous hydrogen/liquid oxygen propellant studies for the shear coaxial injector focused on breakup mechanisms associated with the liquid oxygen jet under subcritical pressure conditions. Laser sheet illumination techniques were used to visualize the core region of the jet and a Phase Doppler Particle Analyzer was utilized for drop velocity, size and size distribution characterization. The results of these studies indicated that the shear coaxial geometry configuration was a relatively poor injector in terms of mixing. The oxygen core was observed to

115 extend well downstream of the injector and a signiﬁcant fraction of the mixing occurred in the near nozzle region where measurements were not possible to obtain. Detailed velocity and species measurements were obtained to allow CFD model validation and this set of benchmark data represents the most comprehensive data set available to date. As an extension of the investigation, a series of gas/gas injector studies were conducted in support of the X-33 Reusable Launch Vehicle program. A Gas/Gas Injector Technology team was formed consisting of the Marshall Space Flight Center, the NASA Lewis Research Center, Rocketdyne and Penn State. Injector geometries studied under this task included shear and swirl coaxial conﬁgurations as well as an impinging jet injector. Author Optical Measurement; Flow Distribution; Nonintrusive Measurement; Thrust Chambers; Velocity Measurement; Injectors; Combustible Flow

19970029012 Analytic Services and Materials, Inc., Hampton, VA USA Bonding and Sealing Evaluations for Cryogenic Tanks Glass, David E.; Aug. 1997; In English Contract(s)/Grant(s): NAS1-96014; RTOP 242-33-01 Report No.(s): NASA-CR-201734; NAS 1.26:201734; ASM-LS05-97-01; No Copyright; Avail: CASI; A03, Hardcopy Several different cryogenic tank concepts are being considered for reusable launch vehicles (RLV’S) . Though different tank concepts are being considered, many will require that the cryogenic insulation be evacuated and be bonded to a structure. In this work, an attempt was made to evaluate the effectiveness of maintaining a vacuum on a specimen where foam or honeycomb core was encased within Gr/Ep. In addition to these tests, ﬂatwise adhesion pull off tests were performed at room temperature with PR 1664, EA 9394, FM-300, Crest 3170, and HT 435 adhesives. The materials bonded included Gr/Ep, Gr/BMI, Al, and stainless steel facesheets, and Ti honeycomb, Hexcel honeycomb, and Rohacell foam core materials. Author Bonding; Sealing; Cryogenic Fluid Storage; Adhesion Tests; Honeycomb Cores; Reusable Launch Vehicles; Bonded Joints

19970028584 Princeton Synergetics, Inc., NJ USA STARS: The Space Transportation Architecture Risk System Greenberg, Joel S.; Apr. 1997; In English Contract(s)/Grant(s): NASA Order H-27284D Report No.(s): NASA-CR-2051;87; NAS 1.26:205187; No Copyright; Avail: CASI; A04, Hardcopy Because of the need to perform comparisons between transportation systems that are likely to have signiﬁcantly different levels of risk, both because of differing degrees of freedom in achieving desired performance levels and their different states of development and utilization, an approach has been developed for performing early comparisons of transportation architectures explicitly taking into account quantitative measures of uncertainty and resulting risk. The approach considers the uncertainty associated with the achievement of technology goals, the effect that the achieved level of technology will have on transportation system performance and the relationship between transportation system performance/capability and the ability to accommodate variations in payload mass. The consequences of system performance are developed in terms of expected values and associated standard deviations of nonrecurring, recurring and the present value of transportation system life cycle cost. Typical results are presented to illustrate the application of the methodology. Author Space Transportation System; Risk; Life Cycle Costs; Monte Carlo Method

19970028362 Boeing Defense and Space Group, Canoga Park, CA USA Space Shuttle Main Engine Start with Off-Nominal Propellant Inlet Pressures Bradley, Michael; 1997; In English; 33rd Joint Propulsion, 6-9 Jul. 1997, Seattle, WA, USA Contract(s)/Grant(s): NCC8-45 Report No.(s): NASA-CR-205359; NAS 1.26:205359; AIAA Paper 97-2687; No Copyright; Avail: CASI; A03, Hardcopy This paper describes Rocketdyne’s successful analysis and demonstration of the Space Shuttle Main Engine (SSME) operation at off-nominal propellant inlet conditions during the Reusable Launch Vehicle (RLV) evaluation tests. The nominal inlet condition range is: 103 to 111 psia and 170.5 to 178 deg. R for the oxidizer and 43 to 47 psia and 37 to 40 deg. R for the fuel. The SSME start was successfully demonstrated with engine inlet pressures of 50 psia liquid oxygen (LOX) with

116 subcooled LOX at 160 deg R and 38 psia fuel at 38 deg. R. Four tests were used to incrementally modify the start sequence to demonstrate the ﬁnal goal. Author Space Shuttle Main Engine; Liquid Oxygen; Engine Inlets; Oxidizers; Mixing Ratios; Liquid Flow

19970028361 Boeing Defense and Space Group, Canoga Park, CA USA Space Shuttle Main Engine Off-Nominal Low Power Level Operation Bradley, Michael; 1997; In English; 33rd Joint Propulsion, 6-9 Jul. 1997, Seattle, WA, USA Contract(s)/Grant(s): NCC8-45 Report No.(s): NASA-CR-205360; NAS 1.26:205360; AIAA Paper 97-2685; Copyright; Avail: CASI; A03, Hardcopy This paper describes Rocketdyne’s successful analysis and demonstration of the Space Shuttle Main Engine (SSME) operation at off-nominal power levels during Reusable Launch Vehicle (RLV) evaluation tests. The nominal power level range for the SSME is from 65% rated power level (RPL) to 109% RPL. Off-nominal power levels incrementally demonstrated were: 17% RPL, 22% RPL, 27% RPL, 40% RPL, 45% RPL, and 50% RPL. Additional achievements during low power operation included: use of a hydrostatic bearing High Pressure Oxidizer Turbopump (HPOTP), nominal High Pressure Fuel Turbopump (HPFTP) first rotor critical speed operation, combustion stability at low power levels, and refined definition of nozzle flow separation heat loads. Author Space Shuttle Main Engine; Turbine Pumps; Separated Flow; Oxidizers; Nozzle Flow; Hydrostatic Pressure; Combustion Stability

19970028022 NASA Langley Research Center, Hampton, VA USA Systems Challenges for Hypersonic Vehicles Hunt, James L.; Laruelle, Gerard; Wagner, Alain; 1997; In English; Future Aerospace Technology in the Service of the Alliance, 14-16 Apr. 1997, Palaiseau, France Report No.(s): NASA-TM-112908; NAS 1.15:112908; AGARD-Paper-C37; No Copyright; Avail: CASI; A03, Hardcopy This paper examines the system challenges posed by fully reusable hypersonic cruise airplanes and access to space vehicles. Hydrocarbon and hydrogen fueled airplanes are considered with cruise speeds of Mach 5 and 10, respectively. The access to space matrix is examined. Airbreathing and rocket powered, single- and two-stage vehicles are considered. Reference vehicle architectures are presented. Major systems/subsystems challenges are described. Advanced, enhancing systems concepts as well as common system technologies are discussed. Author Hypersonic Vehicles; Hypersonics; Hydrogen Fuels; Hydrocarbon Fuels; Air Breathing Engines; Rocket Engines

19970027074 NASA Ames Research Center, Moffett Field, CA USA Evaluation of Thermal Control Coatings for Flexible Ceramic Thermal Protection Systems Kourtides, Demetrius; Carroll, Carol; Smith, Dane; Guzinski, Mike; Marschall, Jochen; Pallix, Joan; Ridge, Jerry; Tran, Duoc; Jul. 1997; In English Contract(s)/Grant(s): RTOP 242-20-01 Report No.(s): NASA-TM-112199; NAS 1.15:112199; A-976757; No Copyright; Avail: CASI; A03, Hardcopy This report summarizes the evaluation and testing of high emissivity protective coatings applied to ﬂexible insulations for the Reusable Launch Vehicle technology program. Ceramic coatings were evaluated for their thermal properties, durability, and potential for reuse. One of the major goals was to determine the mechanism by which these coated blanket surfaces become brittle and try to modify the coatings to reduce or eliminate embrittlement. Coatings were prepared from colloidal silica with a small percentage of either SiC or SiB6 as the emissivity agent. These coatings are referred to as gray C-9 and protective ceramic coating (PCC), respectively. The colloidal solutions were either brushed or sprayed onto advanced ﬂexible reusable surface insulation blankets. The blankets were instrumented with thermocouples and exposed to reentry heating conditions in the Ames Aeroheating Arc Jet Facility. Post-test samples were then characterized through impact testing, emissivity measurements, chemical analysis, and observation of changes in surface morphology. The results show that both coatings performed well in arc jet tests with backface temperatures slightly lower for the PCC coating than with gray C-9. Impact testing showed that the least extensive surface destruction was experienced on blankets with lower areal density coatings. Author Thermal Control Coatings; Thermal Protection; Reusable Launch Vehicles; Reentry Effects; Protective Coatings; Durability; Chemical Analysis; Ceramic Coatings; Brittleness; Aerodynamic Heating

117 19970026958 NASA Ames Research Center, Moffett Field, CA USA Surface Catalysis and Characterization of Proposed Candidate TPS for Access-to-Space Vehicles Stewart, David A.; Jul. 1997; In English; Original contains color illustrations Contract(s)/Grant(s): RTOP 242-30-01 Report No.(s): NASA-TM-112206; NAS 1.15:112206; A-976955; No Copyright; Avail: CASI; A04, Hardcopy Surface properties have been obtained on several classes of thermal protection systems (TPS) using data from both side-arm-reactor and arc-jet facilities. Thermochemical stability, optical properties, and coefficients for atom recombination were determined for candidate TPS proposed for single-stage-to-orbit vehicles. The systems included rigid fibrous insulations, blankets, reinforced carbon carbon, and metals. Test techniques, theories used to define arc-jet and side-arm-reactor flow, and material surface properties are described. Total hemispherical emittance and atom recombination coefficients for each candidate TPS are summarized in the form of polynomial and Arrhenius expressions. Author Thermal Protection; Catalysis; Surface Properties; Thermochemical Properties; Emittance; Atomic Recombination; Recombination Coeffıcient

19970024873 NASA Lewis Research Center, Cleveland, OH USA Comparison of High Aspect Ratio Cooling Channel Designs for a Rocket Combustion Chamber Wadel, Mary F.; Jun. 1997; In English; 33rd Joint Propulsion Conference and Exhibit, 6-9 Jul. 1997, Seattle, WA, USA Contract(s)/Grant(s): RTOP 242-20-06 Report No.(s): NASA-TM-107473; NAS 1.15:107473; E-10762; AIAA Paper 97-2913; No Copyright; Avail: CASI; A03, Hardcopy An analytical investigation on the effect of high aspect ratio (height/width) cooling channels, considering different coolant channel designs, on hot-gas-side wall temperature and coolant pressure drop for a liquid hydrogen cooled rocket combustion chamber, was performed. Coolant channel design elements considered were: length of combustion chamber in which high aspect ratio cooling was applied, number of coolant channels, and coolant channel shape. Seven coolant channel designs were investigated using a coupling of the Rocket Thermal Evaluation code and the Two-Dimensional Kinetics code. Initially, each coolant channel design was developed, without consideration for fabrication, to reduce the hot-gas-side wall temperature from a given conventional cooling channel baseline. These designs produced hot-gas-side wall temperature reductions up to 22 percent, with coolant pressure drop increases as low as 7.5 percent from the baseline. Fabrication constraints for milled channels were applied to the seven designs. These produced hot-gas-side wall temperature reductions of up to 20 percent, with coolant pressure drop increases as low as 2 percent. Using high aspect ratio cooling channels for the entire length of the combustion chamber had no additional beneﬁt on hot-gas-side wall temperature over using high aspect ratio cooling channels only in the throat region, but increased coolant pressure drop 33 percent. Independent of coolant channel shape, high aspect ratio cooling was able to reduce the hot-gas-side wall temperature by at least 8 percent, with as low as a 2 percent increase in coolant pressure drop. The design with the highest overall beneﬁt to hot-gas-side wall temperature and minimal coolant pressure drop cooling can now be done in relatively short periods of time with multiple iterations. Author Combustion Chambers; High Aspect Ratio; Cooling; Channels; Computer Programs; Design

19970024861 NASA Lewis Research Center, Cleveland, OH USA Performance Tests of a Liquid Hydrogen Propellant Densification Ground System for the X33/RLV Tomsik, Thomas M.; Jun. 1997; In English; 33rd Joint Propulsion Conference and Exhibit, 6-9 Jul. 1997, Seattle, WA, USA Contract(s)/Grant(s): RTOP 242-33-01 Report No.(s): NASA-TM-107469; NAS 1.15:107469; AIAA Paper 97-2976; E-10737; No Copyright; Avail: CASI; A03, Hardcopy A concept for improving the performance of propulsion systems in expendable and single-stage-to-orbit (SSTO) launch vehicles much like the X33/RLV has been identified. The approach is to utilize densified cryogenic liquid hydrogen (LH2) and liquid oxygen (LOX) propellants to fuel the propulsion stage. The primary benefit for using this relatively high specific impulse densified propellant mixture is the subsequent reduction of the launch vehicle gross lift-off weight. Production of densified propellants however requires specialized equipment to actively subcool both the liquid oxygen and liquid hydrogen to temperatures below their normal boiling point. A propellant densification unit based on an external thermodynamic vent principle which operates at subatmospheric pressure and supercold temperatures provides a means for the LH2 and LOX densification process to occur. To demonstrate the production concept for the densification of the liquid hydrogen propellant, a system comprised of a multistage gaseous hydrogen compressor, LH2 recirculation pumps and a cryogenic LH2 heat

118 exchanger was designed, built and tested at the NASA Lewis Research Center (LeRC). This paper presents the design configuration of the LH2 propellant densification production hardware, analytical details and results of performance testing conducted with the hydrogen densifier Ground Support Equipment (GSE). Author Propulsion System Performance; Liquid Hydrogen; Densification; Cryogenic Rocket Propellants; Performance Tests; Design Analysis

19970023998 GPS Solutions, Inc., Carson City, NV USA Reusable Launch Vehicle Research Brinker, Stan; Crisalli, Dave; French, Jim; Henry, John; Quisenberry, Bob; Scottoline, Charles; Shell, Dale; Sullivan, Greg; May 31, 1997; In English Contract(s)/Grant(s): NAS8-97212 Report No.(s): NASA-CR-204526; NAS 1.26:204526; No Copyright; Avail: CASI; A24, Hardcopy NASA is engaged in the development of technologies and experimental Reusable Launch Vehicles (RLV) for the purpose of signiﬁcantly reducing the cost of space access through the commercialization of space launch. This involves the unprecedented integration of often conﬂicting business and technical requirements and activities. The following efforts shall enhance NASA’s ability to successfully enable low cost space access while meeting National needs. Author Reusable Launch Vehicles; Cost Reduction; Spacecraft Launching

19970017261 Santa Clara Univ., CA USA Minimum Heating Re-Entry Trajectories for Advanced Hypersonic Launch Vehicles Windhorst, Robert; Feb. 05, 1997; In English Contract(s)/Grant(s): NCC2-5165 Report No.(s): NASA-CR-204339; NAS 1.26:204339; No Copyright; Avail: CASI; A04, Hardcopy Optimal re-entry trajectories are generated for reusable launch vehicles which minimize: (1) the heat absorbed at the vehicle surface, (2) the lower surface temperature, and (3) the heat absorbed by the internal structure. The approach uses the energy state approximation technique and a ﬁnite control volume heat transfer code coupled to a ﬂight path integration code. These trajectories are compared to the optimal re-entry trajectory minimizing the integrated convective heat rate to determine which trajectory produces the minimum internal structural temperatures for a given thermal protection system. Three different thermal protection systems are considered: tile, blanket, and metallic. Author Reentry Trajectories; Reusable Launch Vehicles; Hypersonic Vehicles; Convective Heat Transfer

19970016839 NASA Ames Research Center, Moffett Field, CA USA Near-Optimal Operation of Dual-Fuel Launch Vehicles Ardema, M. D.; Chou, H. C.; Bowles, J. V.; 1996; In English; Atmospheric Flight Mechanics, 29-31 Jul. 1996, San Diego, CA, USA Contract(s)/Grant(s): NCC2-5165 Report No.(s): NASA-CR-204358; NAS 1.26:204358; AIAA Paper 96-3397; Copyright; Avail: CASI; A03, Hardcopy A near-optimal guidance law for the ascent trajectory from earth surface to earth orbit of a fully reusable single-stage-to-orbit pure rocket launch vehicle is derived. Of interest are both the optimal operation of the propulsion system and the optimal ﬂight path. A methodology is developed to investigate the optimal throttle switching of dual-fuel engines. The method is based on selecting propulsion system modes and parameters that maximize a certain performance function. This function is derived from consideration of the energy-state model of the aircraft equations of motion. Because the density of liquid hydrogen is relatively low, the sensitivity of perturbations in volume need to be taken into consideration as well as weight sensitivity. The cost functional is a weighted sum of fuel mass and volume; the weighting factor is chosen to minimize vehicle empty weight for a given payload mass and volume in orbit. Author Ascent Trajectories; Single Stage to Orbit Vehicles; Propulsion System Performance; Flight Paths; Guidance (Motion); Algorithms; Trajectory Optimization; Mach Number

119 19970015639 NASA Ames Research Center, Moffett Field, CA USA Director’s Discretionary Fund Report for Fiscal Year 1996 Mar. 1997; In English Contract(s)/Grant(s): RTOP 274-53-71 Report No.(s): NASA-TM-110445; NAS 1.15:110445; A-975819; No Copyright; Avail: CASI; A07, Hardcopy Topics covered include: Waterprooﬁng the Space Shuttle tiles, thermal protection system for Reusable Launch Vehicles, computer modeling of the thermal conductivity of cometary ice, effects of ozone depletion and ultraviolet radiation on plants, a novel telemetric biosensor to monitor blood pH on-line, ion mobility in polymer electrolytes for lithium-polymer batteries, a microwave-pumped far infrared photoconductor, and a new method for measuring cloud liquid vapor using near infrared remote sensing. Also included: laser-spectroscopic instrument for turbulence measurement, remote sensing of aircraft contrails using a ﬁeld portable imaging interferometer, development of a silicon-micromachined gas chromatography system for determination of planetary surface composition, planar Doppler velocimetry, chaos in interstellar chemistry, and a limited pressure cycle engine for high-speed output. Derived from text Thermal Protection; Ozone Depletion; Remote Sensing; Climate Change; Photoconductors; Comets; Turbulent Flow; Semiconductor Lasers; Gas Chromatography; Interstellar Chemistry; Tilt Rotor Aircraft

19970014941 NASA Langley Research Center, Hampton, VA USA Multidisciplinary Approach to Aerospike Nozzle Design Korte, J. J.; Salas, A. O.; Dunn, H. J.; Alexandrov, N. M.; Follett, W. W.; Orient, G. E.; Hadid, A. H.; Feb. 1997; In English Contract(s)/Grant(s): RTOP 522-31-41-02 Report No.(s): NASA-TM-110326; NAS 1.15:110326; No Copyright; Avail: CASI; A03, Hardcopy A model of a linear aerospike rocket nozzle that consists of coupled aerodynamic and structural analyses has been developed. A nonlinear computational fluid dynamics code is used to calculate the aerodynamic thrust, and a three-dimensional finite-element model is used to determine the structural response and weight. The model will be used to demonstrate multidisciplinary design optimization (MDO) capabilities for relevant engine concepts, assess performance of various MDO approaches, and provide a guide for future application development. In this study, the MDO problem is formulated using the multidisciplinary feasible (MDF) strategy. The results for the MDF formulation are presented with comparisons against separate aerodynamic and structural optimized designs. Significant improvements are demonstrated by using a multidisciplinary approach in comparison with the single-discipline design strategy. Author Nozzle Design; Engine Design; Multidisciplinary Design Optimization; Finite Element Method; Rocket Nozzles

19970014807 Georgia Inst. of Tech., Atlanta, GA USA Supercharged Ejector Ramjet with Maglifter Launch Assist (for the Argus Concept) Olds, John R.; Apr. 08, 1997; In English Contract(s)/Grant(s): NCC8-108 Report No.(s): NASA-CR-204070; NAS 1.26:204070; No Copyright; Avail: CASI; A03, Hardcopy Main propulsion system consists of two 330 klb-class LOX/LH2 Supercharged Ejector RamJet (SERJ) RBCC engines. Engines are capable of multi-mode operation including supercharged ejector, fan-ramjet, ramjet, and pure rocket modes. A low thrust ‘fan-only’ mode is available for powered landing, go-around, taxi, etc. (5 minutes at full throttle). Baseline Maglifter launch assist system consists of a magnetically levitated sled and track system that accelerates the vehicle to approximately 800 fps. At an average acceleration of 1-g, the Maglifter track will be approximately 10,000 ft. long. The baseline system assumes an exposed, horizontal track constructed at a sea-level launch site. Power for the Maglifter is drawn from a large, nearby energy storage system (capacitor-like). The Argus launch vehicle is a moderate lift-to-drag ratio fuselage (circular cross section), highly reusable launch vehicle. It operates autonomously (or remotely). There is no crew cabin. It is capable of delivering and returning a 20 klb payload or 6 passengers to a 100 nmi. circular low earth orbit, due east from KSC (about 10 klb to Space Station Alpha). The fan-derived propulsion modes allow improved aborts and landing operations. Without a payload, the vehicle is capable of transcontinental self-ferry. Derived from text Propulsion; Reusable Launch Vehicles; Ramjet Engines; Ejectors

120 19970012889 NASA Langley Research Center, Hampton, VA USA Effects of Stiffening and Mechanical Load on Thermal Buckling of Stiffened Cylindrical Shells Johnson, Theodore F.; Card, Michael F.; Apr. 12, 1995; In English; 36th Structures, Structural Dynamics and Materials Conference, 10-12 Apr. 1995, New Orleans, LA, USA Report No.(s): NASA-TM-112006; NAS 1.15:112006; AIAA Paper 95-1317; No Copyright; Avail: CASI; A02, Hardcopy A study of thermal buckling of stiffened cylindrical shells with the proportions of a preliminary supersonic transport fuselage design (1970) is presented. The buckling analysis is performed using an axisymmetric shell-of-revolution code, BOSOR4. The effects of combined mechanical (axial loading) and thermal loading (heated skins) are investigated. Results indicate that the location of longitudinal eccentric stiffening has a very large effect on the thermal buckling strength of longitudinally stiffened shells, and on longitudinally stiffened shells with rings. Author Thermal Buckling; Cylindrical Shells; Stiffening; Mechanical Properties; Applications Programs (Computers); Axial Loads; Transatmospheric Vehicles; Digital Simulation; Computerized Simulation

19970012468 NASA Ames Research Center, Moffett Field, CA USA Materials Testing on the DC-X and DC-XA Smith, Dane; Carroll, Carol; Marschall, Jochen; Pallix, Joan; Jan. 1997; In English; Original contains color illustrations Contract(s)/Grant(s): RTOP 242-30-01 Report No.(s): NASA-TM-110430; NAS 1.15:110430; A-975493; No Copyright; Avail: CASI; A03, Hardcopy Flight testing of thermal protection materials has been carried out over a two year period on the base heat shield of the Delta Clipper (DC-X and DC-XA), as well on a body flap. The purpose was to use the vehicle as a test bed for materials and more efficient repair or maintenance processes which would be potentially useful for application on new entry vehicles (i.e., X-33, RLV, planetary probes), as well as on the existing space shuttle orbiters. Panels containing Thermal Protection Systems (TPS) and/or structural materials were constructed either at NASAAmes Research Center or at McDonnell Douglas Aerospace (MDA) and attached between two of the four thrusters in the base heat shield of the DC-X or DC-XA. Three different panels were flown on DC-X flights 6, 7, and 8. A total of 7 panels were flown on DC-XA flights 1, 2, and 3. The panels constructed at Ames contained a variety of ceramic TPS including flexible blankets, tiles with high emissivity coatings, lightweight ceramic ablators and other ceramic composites. The MDS test panels consisted primarily of a variety of metallic composites. This report focuses on the ceramic TPS test results. Author Thermal Protection; Delta Clipper; Heat Shielding; Ceramic Matrix Composites; Ablative Materials

19970012335 Georgia Inst. of Tech., Atlanta, GA USA Technology and Risk Assessment Using a Robust Design Simulation Methodology, Phase 1, Definition and Initial Implementation of RDS Mavris, Dimitri N.; 1997; In English Report No.(s): NASA-CR-203308; NAS 1.26:203308; No Copyright; Avail: CASI; A02, Hardcopy This year we were able to implement the Robust Design Methodology developed at Aerospace System Design Laboratory (ASDL) into our aerospace engineering graduate design curriculum by having this year’s design team execute the methodology for their team project. The project was focused on the preliminary design of an High Speed Civil Transport (HSCT) concept, and more specifically, the development of a design configuration that is insensitive to variability induced by the economic environment it is operated in. This work is summarized in the proceedings of this year’s ASDL external advisory board symposium held at Georgia Tech in May 1996 (proceedings have already been mailed to sponsor) and a 300 page final report submitted by the team which will be mailed. Derived from text Aerospace Engineering; Aircraft Design; Aerospace Planes; Design to Cost

19970011757 Society of Automotive Engineers, Inc., Warrendale, PA USA Results of a Rocket-Based Combined-Cycle SSTO Design Using Parametric MDO Methods Olds, John R.; 1994; ISSN 0148-7191; In English; Aerospace Atlantic, 18-22 Apr. 1994, Dayton, OH, USA Contract(s)/Grant(s): NCC1-168; NAGw-1331 Report No.(s): NASA-CR-203224; NAS 1.26:203224; SAE-TP-941165; Copyright; Avail: CASI; A03, Hardcopy This paper reports the results of the second phase of a research project to characterize and optimize the design of an

121 advanced launch vehicle for human access to low earth orbit. The vehicle makes use of Rocket-Based Combined-Cycle (RBCC) propulsion - a concept combining operating modes of an ejector, ramjet, scramjet, and rocket in a single engine. This research builds on previous work focused on advanced multiple mode propulsion concepts and advanced conical acceleration-class Single-Stage-To-Orbit (SSTO) launch vehicles. Three systems level design variables of interest were optimized using Multidisciplinary Design Optimization (MDO) techniques. Speciﬁcally, Taguchi’s method of robust design was used to identify a combination of variables that minimize the vehicle sensitivity to unpredictable changes in engine weights and performance. In addition, a second-order Response Surface Method (RSM) was used to approximate the design space and predict the minimum dry weight vehicle. The optimized vehicle results (weights, dimensions, performance) are favorably compared with other SSTO designs including rocket and airbreathing concepts. Author Design Analysis; Launch Vehicles; Low Earth Orbits

19970009763 NASA Johnson Space Center, Houston, TX USA A Collection of Technical Papers A Collection of Technical Papers from the 6th Space Logistics Symposium; 1995; In English; 6th Space Logistics Symposium, 22-24 Feb. 1995, Houston, TX, USA Report No.(s): NASA-TM-112136; NAS 1.15:112136; Copyright; Avail: CASI; A10, Hardcopy Papers presented at the 6th Space Logistics Symposium covered such areas as: The International Space Station; The Hubble Space Telescope; Launch site computer simulation; Integrated logistics support; The Baikonur Cosmodrome; Probabalistic tools for high conﬁdence repair; A simple space station rescue vehicle; Integrated Traffic Model for the International Space Station; Packaging the maintenance shop; Leading edge software support; Storage information management system; Consolidated maintenance inventory logistics planning; Operation concepts for a single stage to orbit vehicle; Mission architecture for human lunar exploration; Logistics of a lunar based solar power satellite scenario; Just in time in space; NASA acquisitions/logistics; Effective transition management; Shuttle logistics; and Revitalized space operations through total quality control management. Derived from text International Space Station; Single Stage to Orbit Vehicles; Solar Power Satellites; Space Logistics; Logistics Management; Rescue Operations; Support Systems; Total Quality Management

19970009604 NASA Langley Research Center, Hampton, VA USA Aerodynamic Characteristics of a Vertical Takeoff Vertical Landing (VTVL) Single Stage to Orbit Vehicle from M(sub infinity) = 0.1 to 10 Woods, W. C.; Merski, N. R.; 1995; In English; 13th Applied Aerodynamics Conference, 19-22 Jun. 1995, San Diego, CA, USA Report No.(s): NASA-TM-112160; NAS 1.15:112160; AIAA Paper 95-1828; Copyright; Avail: CASI; A03, Hardcopy Future access to space studies have considered a wide variety of follow on options to augment/replace the Space Shuttle. One option is a vertical takeoff and landing single stage to orbit vehicle. Experimental aerodynamic characteristics have been obtained on a generic design representative of such a vehicle; i.e., spherical nose blunting, a forebody of revolution and flat surfaces on the afterbody to accommodate control surfaces. Data has been obtained at subsonic, supersonic, and hypersonic Mach numbers to aid in determining cross range capability and performance, stability and control characteristics. The baseline configuration is longitudinally and laterally unstable, but appears to have the control authority to provide trim and stability augmentation. At subsonic speeds and angles of attack above 30 deg large lateral forces occur at zero sideslip and the lateral/directional characteristics are not well behaved. Author Aerodynamic Characteristics; Single Stage to Orbit Vehicles; Afterbodies; Angle of Attack; Control Surfaces; Noses (Forebodies); Vertical Landing; Vertical Takeoff; Mach Number

19970008365 NASA Marshall Space Flight Center, Huntsville, AL USA Computational Aerodynamics Analysis of Future Launch Vehicle Configurations Vu, B.; McConnaughey, P.; Huddleston, D.; 1994; In English; 12th Applied Aerodynamics, 20-22 Jun. 1994, Colorado Springs, CO, USA Report No.(s): NASA-TM-112141; NAS 1.15:112141; AIAA Paper 94-1936; No Copyright; Avail: CASI; A02, Hardcopy An implicit finite volume high resolution scheme is applied for predicting three-dimensional, inviscid flow field over

122 several launch vehicles including the National Launch System (NLS), the Single-Stage-To-Orbit (SSTO) lifting bodies, and wing-body vehicles. Simulations about NLS conﬁgurations were used to help benchmark Computational Fluid Dynamics (CFD) capabilities at NASA/MSFC. Likewise, simulations about SSTO were used to provide aerodynamic data to structure group to calculate structural wing loading for preliminary conceptual designs. Author Computational Fluid Dynamics; Computerized Simulation; Digital Simulation; Single Stage to Orbit Vehicles; Three Dimensional Flow; Lifting Bodies; Launch Vehicles; Launch Vehicle Conﬁgurations; Finite Volume Method

19970006985 NASA Marshall Space Flight Center, Huntsville, AL USA Impact Characteristics of Candidate Materials for Single-Stage-to-Orbit (SSTO) Technology Nettles, Alan; 1995; In English; 40th International SAMPE Symposium and Exhibition, 8-11 May 1995, Anaheim, CA, USA Report No.(s): NASA-TM-111794; NAS 1.15:111794; No Copyright; Avail: CASI; A03, Hardcopy Four ﬁber/resin systems were compared for resistance to damage and damage tolerance. One toughened epoxy and three toughened bismaleimide (BMI) resins were used., all with IM7 carbon ﬁber reinforcement. A statistical design of experiments technique was used to evaluate the effects of impact energy, specimen thickness and tup diameter on the damage area and residual compression-after-impact (CAI) strength. Results showed that two of the BMI systems sustained relatively large damage areas yet had an excellent retention of CAI strength. Author Impact Strength; Impact Resistance; Polymer Matrix Composites; Bismaleimide; Polyimide Resins; Epoxy Resins; Impact Tests; Nondestructive Tests; Spacecraft Construction Materials

19970005470 NASA Langley Research Center, Hampton, VA USA Comparison of Two Multidisciplinary Optimization Strategies for Launch-Vehicle Design Braun, R. D.; Powell, R. W.; Lepsch, R. A.; Stanley, D. O.; Kroo, I. M.; Journal of Spacecraft and Rockets; Jun. 1995; Volume 33, No. 3; In English Report No.(s): NASA-TM-111822; NAS 1.15:111822; No Copyright; Avail: CASI; A02, Hardcopy The investigation focuses on development of a rapid multidisciplinary analysis and optimization capability for launch-vehicle design. Two multidisciplinary optimization strategies in which the analyses are integrated in different manners are implemented and evaluated for solution of a single-stage-to-orbit launch-vehicle design problem. Weights and sizing, propulsion, and trajectory issues are directly addressed in each optimization process. Additionally, the need to maintain a consistent vehicle model across the disciplines is discussed. Both solution strategies were shown to obtain similar solutions from two different starting points. These solutions suggests that a dual-fuel, single-stage-to-orbit vehicle with a dry weight of approximately 1.927 x 10(exp 5)lb, gross liftoff weight of 2.165 x 10(exp 6)lb, and length of 181 ft is attainable. A comparison of the two approaches demonstrates that treatment or disciplinary coupling has a direct effect on optimization convergence and the required computational effort. In comparison with the ﬁrst solution strategy, which is of the general form typically used within the launch vehicle design community at present, the second optimization approach is shown to he 3-4 times more computationally efficient. Author Design Analysis; Optimization; Launch Vehicles; Convergence; Single Stage to Orbit Vehicles

19970005454 NASA Marshall Space Flight Center, Huntsville, AL USA Advanced Low-Cost O2/H2 Engines for the SSTO Application Goracke, B. David; Levack, Daniel J. H.; Nixon, Robert F.; 1994; In English; 30th Joint Propulsion, 27-29 Jun. 1994, Indianapolis, IN, USA Report No.(s): NASA-TM-111897; NAS 1.15:111897; AIAA Paper 94-3317; Copyright; Avail: CASI; A02, Hardcopy The recent NASAAccess to Space study examined future Earth to orbit (ETO) transportation needs and fleets out to 2030. The baseline in the option 3 assessment was a single stage to orbit (SSTO) vehicle. A study was conducted to assess the use of new advanced low cost O2/H2 engines for this SSTO application. The study defined baseline configurations and ground rules and defined six engine cycles to explore engine performance. The cycles included an open cycle, and a series of closed cycles with varying abilities to extract energy from the propellants to power he turbomachinery. The cycles thus varied in the maximum chamber pressure they could reach and in their weights at any given chamber pressure. The weight of each cycle was calculated for two technology levels versus chamber pressure up to the power limit of the cycle. The performance in the SSTO mission was then modeled using the resulting engine weights and specific impulse performance using the Access to

123 Space option 3 vehicle. The results showed that new O2/H2 engines are viable and competitive candidates for the SSTO application using chamber pressures of 4,000 psi. Author Single Stage to Orbit Vehicles; Thermodynamic Cycles; Speciﬁc Impulse; Engine Design; Propulsion System Conﬁgurations; Closed Cycles; Liquid Propellant Rocket Engines

19970005361 National Aeronautics and Space Administration. Langley Research Center, Hampton, VA USA Development of Metallic Thermal Protection Systems for the Reusable Launch Vehicle Blosser, Max L.; Oct. 1996; In English; Space Technology and Applications International Forum, 26-30 Jan. 1997, Albuquerque, NM, USA Contract(s)/Grant(s): RTOP 242-20-02-01 Report No.(s): NASA-TM-110296; NAS 1.15:110296; No Copyright; Avail: CASI; A03, Hardcopy A reusable Thermal Protection System (TPS) that is not only lightweight, but durable, operable and cost effective is one of the technologies required by the Reusable Launch Vehicle (RLV) to achieve the goal of drastically reducing the cost of delivering payload to orbit. Metallic TPS is one of the systems being developed to meet this challenge. Current efforts involve improving the superalloy honeycomb TPS concept, which consists of a foil-gage metallic box encapsulating a low density ﬁbrous insulation, and evaluating it for RLV requirements. The superalloy honeycomb TPS concept is mechanically attached to the vehicle structure. Improvements include more efficient internal insulation, a simpler, lighter weight conﬁguration, and a quick-release fastener system for easier installation and removal. Evaluation includes thermal and structural analysis, fabrication and testing of both coupons and TPS panels under conditions simulating RLV environments. Coupons of metallic honeycomb sandwich, representative of the outer TPS surface, were subjected to low speed impact, hypervelocity impact, and rain erosion testing as well as subsequent arcjet exposure. Arrays of TPS panels have been subjected to radiant heating in a thermal/vacuum facility, aerodynamic heating in an arcjet facility and acoustic loading. Author Thermal Protection; Structural Analysis; Reusable Launch Vehicles; Measuring Instruments; Payloads; Loads (Forces); Hypervelocity Impact; Honeycomb Structures; Heat Resistant Alloys; Aerodynamic Heating; Acoustic Fatigue

19970005131 NASA Langley Research Center, Hampton, VA USA Airbreathing/Rocket Single-Stage-to-Orbit Design Matrix Hunt, James L.; 1995; In English; 6th International Aerospace Planes and Hypersonics Technologies, 3-7 Apr. 1995, Chattanooga, TN, USA Report No.(s): NASA-TM-112140; NAS 1.15:112140; AIAA Paper 95-6011; Copyright; Avail: CASI; A03, Hardcopy A deﬁnitive design/performance study was performed on a single-stage-to-orbit (SSTO) airbreathing propelled orbital vehicle with rocket propulsion augmentation in the Access to Space activities during 1993. A credible reference design was established, but by no means an optimum. The results supported the viability of SSTO airbreathing/rocket vehicles for operational scenarios and indicated compelling reasons to continue to explore the design matrix. This paper will (1) summarize the Access to Space design activity from the SSTO airbreathing/rocket perspective, (2) present an airbreathing/rocket SSTO design matrix established for continued optimization of the design space, and (3) focus on the compelling reasons for airbreathing vehicles in Access to Space scenarios. Author Single Stage to Orbit Vehicles; Air Breathing Engines; Systems Integration; System Effectiveness

19970005122 NASA Langley Research Center, Hampton, VA USA Subsonic-to-Hypersonic Aerodynamic Characteristics for a Winged, Circular-Body, Single-Stage-to-Orbit Spacecraft Configuration Phillips, W. P.; Engelund, W. C.; 1995; In English; Applied Aerodynamics, 19-22 Jun. 1995, San Diego, CA, USA Report No.(s): NASA-TM-111926; NAS 1.15:111926; AIAA Paper 95-1848; Copyright; Avail: CASI; A03, Hardcopy Experimental aerodynamic characteristics were obtained for a generic, winged, circular-body, single-stage-to-orbit spacecraft configuration. The baseline configuration was longitudinally stable and trimmable at almost all Mach numbers from 0.15 to 10.0--with the exception occurring at low supersonic speeds. Landing speed and subsonic-to-hypersonic longitudinal stability and control appear to be within design guidelines. Lateral-directional instabilities found over the entire speed range, however, create a problem area for this configuration. Longitudinal aerodynamic predictions made utilizing the Aerodynamic

124 Preliminary Analysis System (APAS) were in qualitative, often quantitative agreement with experimental values. Author Aerodynamic Characteristics; Spacecraft Conﬁgurations; Subsonic Speed; Supersonic Speed; Hypersonics; Single Stage to Orbit Vehicles; Longitudinal Stability; Controllability

19970005058 NASA Langley Research Center, Hampton, VA USA Dual-Fuel Propulsion in Single-Stage Advanced Manned Launch System Vehicle Lepsch, Roger A., Jr.; Stanley, Douglas O.; Unal, Resit; Journal of Spacecraft and Rockets; Jun. 1995; Volume 32, Number 3; In English; 29th Joint Propulsion Conference, 28-30 Jun. 1993, Monterey, CA, USA Report No.(s): NASA-TM-111903; NAS 1.15:111903; AIAA Paper 93-2275; Copyright; Avail: CASI; A03, Hardcopy As part of the USA Advanced Manned Launch System study to determine a follow-on, or complement, to the Space Shuttle, a reusable single-stage-to-orbit concept utilizing dual-fuel rocket propulsion has been examined. Several dual-fuel propulsion concepts were investigated. These include: a separate-engine concept combining Russian RD-170 kerosene-fueled engines with space shuttle main engine-derivative engines: the kerosene- and hydrogen-fueled Russian RD-701 engine; and a dual-fuel, dual-expander engine. Analysis to determine vehicle weight and size characteristics was performed using conceptual-level design techniques. A response-surface methodology for multidisciplinary design was utilized to optimize the dual-fuel vehicles with respect to several important propulsion-system and vehicle design parameters, in order to achieve minimum empty weight. The tools and methods employed in the analysis process are also summarized. In comparison with a reference hydrogen- fueled single-stage vehicle, results showed that the dual-fuel vehicles were from 10 to 30% lower in empty weight for the same payload capability, with the dual-expander engine types showing the greatest potential. Author Engine Design; Design Analysis; Space Shuttle Main Engine; Propulsion System Conﬁgurations; Single Stage to Orbit Vehicles; Weight Reduction

19970003025 Maryland Univ., College Park, MD USA Effect of Aeropropulsive Interactions and Design Sensitivities on Optimal Hypersonic Ascent Trajectories Lovell, T. Alan; Schmidt, David K.; [1994]; In English Contract(s)/Grant(s): NAG1-1540 Report No.(s): NASA-CR-202601; NAS 1.26:202601; AIAA Paper 94-3524; Copyright; Avail: CASI; A03, Hardcopy In this paper, trajectory optimization is used as a tool to better understand the performance characteristics of hypersonic Single-Stage-To-Orbit(SSTO) vehicle exhibiting significant aeropropulsive interactions. The energy state approximation is used to determine a straightforward method of determining the scramjet-powered phase of the mission. The energy state method is then used to generate fuel-optimal trajectory results over this portion of the mission for a vehicle configuration of this class. The fuel-optimal unconstrained trajectory is marked by low-altitude acceleration, while a dynamic pressure- constrained trajectory is seen to ride the dynamic pressure constraint for the entire scramjet mission phase. The significance of aeropropulsive interactions in affecting vehicle performance is also investigated. Author Ascent Trajectories; Trajectory Optimization; Single Stage to Orbit Vehicles; Supersonic Combustion Ramjet Engines; Propulsion System Performance; Fuel Consumption

19970001742 North Carolina State Univ., Raleigh, NC USA System Sensitivity Analysis Applied to the Conceptual Design of a Dual-Fuel Rocket SSTO Olds, John R.; 1994; In English; 5th Symposium on Multidisciplinary Analysis and Optimization, 7-9 Sep. 1994, Panama City Beach, Fl, USA Contract(s)/Grant(s): NCC1-185 Report No.(s): NASA-CR-202599; NAS 1.26:202599; AIAA Paper 94-4339; Copyright; Avail: CASI; A03, Hardcopy This paper reports the results of initial efforts to apply the System Sensitivity Analysis (SSA) optimization method to the conceptual design of a single-stage-to-orbit (SSTO) launch vehicle. SSA is an efficient, calculus-based MDO technique for generating sensitivity derivatives in a highly multidisciplinary design environment. The method has been successfully applied to conceptual aircraft design and has been proven to have advantages over traditional direct optimization methods. The method is applied to the optimization of an advanced, piloted SSTO design similar to vehicles currently being analyzed by NASA as possible replacements for the Space Shuttle. Powered by a derivative of the Russian RD-701 rocket engine, the vehicle employs a combination of hydrocarbon, hydrogen, and oxygen propellants. Three primary disciplines are included in the

125 design - propulsion, performance, and weights & sizing. A complete, converged vehicle analysis depends on the use of three standalone conceptual analysis computer codes. Efforts to minimize vehicle dry (empty) weight are reported in this paper. The problem consists of six system-level design variables and one system-level constraint. Using SSA in a ‘manual’ fashion to generate gradient information, six system-level iterations were performed from each of two different starting points. The results showed a good pattern of convergence for both starting points. A discussion of the advantages and disadvantages of the method, possible areas of improvement, and future work is included. Author Single Stage to Orbit Vehicles; Multidisciplinary Design Optimization; Systems Analysis; Design Analysis

19970001416 Lockheed Martin Engineering and Sciences Co., Hampton, VA USA Computational Study of a McDonnell Douglas Single-Stage-to-Orbit Vehicle Concept for Aerodynamic Analysis Prabhu, Ramadas K.; Sep. 1996; In English Contract(s)/Grant(s): NAS1-19000; RTOP 242-20-80-02 Report No.(s): NASA-CR-201606; NAS 1.26:201606; No Copyright; Avail: CASI; A03, Hardcopy This paper presents the results of a computational flow analysis of the McDonnell Douglas single-stage-to-orbit vehicle concept designated as the 24U. This study was made to determine the aerodynamic characteristics of the vehicle with and without body flaps over an angle of attack range of 20-40 deg. Computations were made at a flight Mach number of 20 at 200,000 ft. altitude with equilibrium air, and a Mach number of 6 with CF4 gas. The software package FELISA (Finite Element Langley imperial College Sawansea Ames) was used for all the computations. The FELISA software consists of unstructured surface and volume grid generators, and inviscid flow solvers with (1) perfect gas option for subsonic, transonic, and low supersonic speeds, and (2) perfect gas, equilibrium air, and CF4 options for hypersonic speeds. The hypersonic flow solvers with equilibrium air and CF4 options were used in the present studies. Results are compared with other computational results and hypersonic CF4 tunnel test data. Author Unstructured Grids (Mathematics); Single Stage to Orbit Vehicles; Inviscid Flow; Hypersonic Speed; Hypersonic Flow; Finite Element Method; Computational Grids; Applications Programs (Computers); Aerodynamic Characteristics

19970001260 NASA, Huntsville, AL USA Acoustic Emission Monitoring of the DC-XA Composite Liquid Hydrogen Tank During Structural Testing Wilkerson, C.; Oct. 1996; In English Report No.(s): NASA-TM-108520; NAS 1.15:108520; No Copyright; Avail: CASI; A03, Hardcopy The results of acoustic emission (AE) monitoring of the DC-XA composite liquid hydrogen tank are presented in this report. The tank was subjected to pressurization, tensile, and compressive loads at ambient temperatures and also while full of liquid nitrogen. The tank was also pressurized with liquid hydrogen. AE was used to monitor the tank for signs of structural defects developing during the test. Author Acoustic Emission; Liquid Hydrogen; Fuel Tanks; Nondestructive Tests; Temperature Effects; Structural Members

19960055049 NASA Langley Research Center, Hampton, VA USA The 1995 NASA High-Speed Research Program Sonic Boom Workshop, Volume 1 Baize, Daniel G., Editor; 1995 NASA High-Speed Research Program Sonic Boom Workshop. Volume 1; Jul. 1996; In English, 12-13 Sep. 1995, Hampton, VA, USA Contract(s)/Grant(s): RTOP 537-07-21-21 Report No.(s): NASA-CP-3335-Vol-1; L-17572-Vol-1; NAS 1.55:3335-Vol-1; No Copyright; Avail: CASI; A17, Hardcopy The High-Speed Research Program and NASA Langley Research Center sponsored the NASA High-Speed Research Program Sonic Boom Workshop on September 12-13, 1995. The workshop was designed to bring together NASAs scientists and engineers and their counterparts in industry, other Government agencies, and academia working together in the sonic boom element of NASAs High-Speed Research Program. Speciﬁc objectives of this workshop were to (1) report the progress and status of research in sonic boom propagation, acceptability, and design; (2) promote and disseminate this technology within the appropriate technical communities; (3) help promote synergy among the scientists working in the Program; and (4) identify technology pacing the development of viable reduced-boom High-Speed Civil Transport concepts. The Workshop included these sessions: Session 1 - Sonic Boom Propagation (Theoretical); Session 2 - Sonic Boom Propagation

126 (Experimental); and Session 3 - Acceptability Studies - Human and Animal. Author Aerospace Planes; Sonic Booms; High Speed; Sound Propagation

19960053992 NASA Langley Research Center, Hampton, VA USA Reusable Launch Vehicle Technology Program Freeman, Delma C., Jr.; Talay, Theodore A.; Austin, R. Eugene; 1996; In English; 47th International Astronautical Congress, 7-11 Oct. 1996, Beijing, China Report No.(s): NASA-TM-110473; NAS 1.15:110473; IAF-96-V.4.01; Copyright; Avail: CASI; A03, Hardcopy Industry/NASA Reusable Launch Vehicle (RLV) Technology Program efforts are underway to design, test, and develop technologies and concepts for viable commercial launch systems that also satisfy national needs at acceptable recurring costs. Significant progress has been made in understanding the technical challenges of fully reusable launch systems and the accompanying management and operational approaches for achieving a low-cost program. This paper reviews the current status of the Reusable Launch Vehicle Technology Program including the DC-XA, X-33 and X-34 flight systems and associated technology programs. It addresses the specific technologies being tested that address the technical and operability challenges of reusable launch systems including reusable cryogenic propellant tanks, composite structures, thermal protection systems, improved propulsion, and subsystem operability enhancements. The recently concluded DC-XA test program demonstrated some of these technologies in ground and flight tests. Contracts were awarded recently for both the X-33 and X-34 flight demonstrator systems. The Orbital Sciences Corporation X-34 flight test vehicle will demonstrate an air-launched reusable vehicle capable of flight to speeds of Mach 8. The Lockheed-Martin X-33 flight test vehicle will expand the test envelope for critical technologies to flight speeds of Mach 15. A propulsion program to test the X-33 linear aerospike rocket engine using a NASA SR-71 high speed aircraft as a test bed is also discussed. The paper also describes the management and operational approaches that address the challenge of new cost-effective, reusable launch vehicle systems. Author Propulsion System Performance; Technologies; Single Stage to Orbit Vehicles; Composite Structures; Cryogenic Fluid Storage; Propellant Tanks

19960049683 Lockheed Martin Corp., Hampton, VA USA A study of facilities and fixtures for testing of a high speed civil transport wing component Cerro, J. A.; Vause, R. F.; Bowman, L. M.; Jensen, J. K.; Martin, C. J., Jr.; Stockwell, A. E.; Waters, W. A., Jr.; Jul. 1996; In English Contract(s)/Grant(s): NAS1-19000; RTOP 537-06-34 Report No.(s): NASA-CR-198352; NAS 1.26:198352; No Copyright; Avail: CASI; A04, Hardcopy A study was performed to determine the feasibility of testing a large-scale High Speed Civil Transport wing component in the Structures and Materials Testing Laboratory in Building 1148 at NASA Langley Research Center. The report includes a survey of the electrical and hydraulic resources and identifies the backing structure and floor hard points which would be available for reacting the test loads. The backing structure analysis uses a new finite element model of the floor and backstop support system in the Structures Laboratory. Information on the data acquisition system and the thermal power requirements is also presented. The study identified the hardware that would be required to test a typical component, including the number and arrangement of hydraulic actuators required to simulate expected flight loads. Load introduction and reaction structure concepts were analyzed to investigate the effects of experimentally induced boundary conditions. Author X-30 Vehicle; Wings; Aerodynamic Loads; Test Chambers; Test Facilities; Finite Element Method; Loads (Forces); Supersonic Transports; Support Systems

19960047476 Boeing Commercial Airplane Co., Seattle, WA USA Materials Research for High Speed Civil Transport and Generic Hypersonics-Metals Durability Schulz, Paul; Hoffman, Daniel; Apr. 1996; In English Contract(s)/Grant(s): NAS1-20013; RTOP 537-06-20-05 Report No.(s): NASA-CR-198326; NAS 1.26:198326; No Copyright; Avail: CASI; A05, Hardcopy This report covers a portion of an ongoing investigation of the durability of titanium alloys for the High Speed Civil Transport (HSCT). Candidate alloys need to possess an acceptable combination of properties including strength and toughness as well as fatigue and corrosion resistance when subjected to the HSCT operational environment. These materials must also

127 be capable of being processed into required product forms while maintaining their properties. Processing operations being considered for this airplane include forming, welding, adhesive bonding, and superplastic forming with or without diffusion bonding. This program was designed to develop the material properties database required to lower the risk of using advanced titanium alloys on the HSCT. Author Aerospace Planes; Hypersonic Aircraft; Titanium Alloys; Supersonic Transports; Metal Bonding; Fatigue (Materials); Fracture Strength; Mechanical Properties; Durability

19960047140 NASA Langley Research Center, Hampton, VA USA Subsonic Aerodynamic Characteristics of a Circular Body Earth-to-Orbit Vehicle Lepsch, Roger A., Jr.; Ware, George M.; MacConochie, Ian O.; Jul. 1996; In English Contract(s)/Grant(s): RTOP 242-20-08-01 Report No.(s): NASA-TM-4726; L-17429; NAS 1.15:4726; No Copyright; Avail: CASI; A03, Hardcopy A test of a generic reusable earth-to-orbit transport was conducted in the 7- by 10-Foot high-speed tunnel at the Langley Research Center at Mach number 0.3. The model had a body with a circular cross section and a thick clipped delta wing as the major lifting surface. For directional control, three different vertical fin arrangements were investigated: a conventional aft-mounted center vertical fin, wingtip fins, and a nose-mounted vertical fin. The configuration was longitudinally stable about the estimated center-of-gravity position of 0.72 body length and had sufficient pitch-control authority for stable trim over a wide range of angle of attack, regardless of fin arrangement. The maximum trimmed lift/drag ratio for the aft center-fin configuration was less than 5, whereas the other configurations had values of above 6. The aft center-fin configuration was directionally stable for all angles of attack tested. The wingtip and nose fins were not intended to produce directional stability but to be active controllers for artificial stabilization. Small rolling-moment values resulted from yaw control of the nose fin. Large adverse rolling-moment increments resulted from tip-fin controller deflection above 13 deg angle of attack. Flow visualization indicated that the adverse rolling-moment increments were probably caused by the influence of the deflected tip-fin controller on wing flow separation. Author Wind Tunnel Tests; Delta Wings; Angle of Attack; Rolling Moments; Flow Visualization; Subsonic Flow; Wind Tunnel Models; Fins; Satellite Control; Directional Control; Lift Drag Ratio

19960047075 Vanderbilt Univ., Nashville, TN USA Dynamics and Control of Orbiting Space Structures NASA Advanced Design Program (ADP) Cruse, T. A.; Aug. 20, 1996; In English Contract(s)/Grant(s): NAGw-4252 Report No.(s): NASA-CR-202061; NAS 1.26:202061; Rept-96-08-02; No Copyright; Avail: CASI; A03, Hardcopy The report summarizes the advanced design program in the mechanical engineering department at Vanderbilt University for the academic years 1994-1995 and 1995-1996. Approximately 100 students participated in the two years of the subject grant funding. The NASA-oriented design projects that were selected included lightweight hydrogen propellant tank for the reusable launch vehicle, a thermal barrier coating test facility, a piezoelectric motor for space antenna control, and a lightweight satellite for automated materials processing. The NASA supported advanced design program (ADP) has been a success and a number of graduates are working in aerospace and are doing design. Author Spacecraft Structures; Reusable Launch Vehicles; Spacecraft Design

19960045833 NASA Marshall Space Flight Center, Huntsville, AL USA The Reusable Launch Vehicle Technology Program and the X-33 Advanced Technology Demonstrator Cook, Stephen A.; Apr. 04, 1995; In English; 6th AIAA Aerospace Plane and Hypersonics Technology Conference, 4 Apr. 1995, Chattanooga, TN, USA Report No.(s): NASA-TM-111868; NAS 1.15:111868; No Copyright; Avail: CASI; A02, Hardcopy The goal of the Reusable Launch Vehicle (RLV) technology program is formulated, and the primary objectives of RLV are listed. RLV technology program implementation phases are outlined. X-33 advanced technology demonstrator is described. Program management is addressed. CASI Technologies; Single Stage to Orbit Vehicles; X-33 Reusable Launch Vehicle; Advanced Launch System (Sts)

128 19960045205 NASA Langley Research Center, Hampton, VA USA Design and Analysis of a Two-Stage-to-Orbit Airbreathing Hypersonic Vehicle Concept Lockwood, Mary Kae; Hunt, James L.; Kabis, Hanee; Moses, Paul; Pao, Jenn-Louh; Yarrington, Phillip; Collier, Craig; Jul. 03, 1996; In English; 32nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 1-3 Jul. 1996, Lake Buena Vista, FL, USA; Original contains color illustrations Report No.(s): NASA-CR-201784; NAS 1.26:201784; AIAA Paper 96-2890; Copyright; Avail: CASI; A03, Hardcopy An initial design of a second generation two-stage-to-orbit vehicle, with an airbreathing first stage, capable of delivering a 2000 lb payload to orbit has been completed. Two low speed propulsion systems were considered for the first stage vehicle for Mach 0 to 3 operation, an ejector ramjet and an advanced air turbo-ramjet engine. A dual mode ramjet/scramjet was used above Mach 3. The airframe structure is composed of cold integral-tank graphite-epoxy stiffened panels with an adhesively-bonded thermal protection system. Active cooling through aluminum heat exchanger panels is used in the engine. The near optimal staging Mach number to minimize vehicle dry weight and take-off gross weight for this mission was determined to be Mach 8. Dry weights range from 67,000 to 69,000 lbs and take-off gross weights range from 119,000 to 131,000 lbs, depending on the low speed propulsion system. Three-stage-to-orbit configurations were also considered, with the first stage being a platform for a Mach .8 launch, and the second and third stages being similar to the first and second stages of the two-stage-to-orbit vehicle. A 10,000 lb reduction in TOGW resulted for the combined 2nd and 3rd stages of the three-stage-to-orbit vehicle, as compared to the two-stage-to-orbit vehicle. This relatively small reduction probably does not warrant the added complexity for the air launch in this case. Author Reusable Launch Vehicles; Air Breathing Engines; Graphite-Epoxy Composites; Hypersonic Vehicles; Propulsion System Configurations; Supersonic Combustion Ramjet Engines

19960042873 Lockheed Martin Missiles and Space, Huntsville, AL USA Advanced Transportation System Studies Technical Area 2 (TA-2) Heavy Lift Launch Vehicle Development Contract, Volume 2, Technical Results Jul. 1995; In English Contract(s)/Grant(s): NAS8-39208 Report No.(s): NASA-CR-201127-Vol-2; NAS 1.26:201127-Vol-2; LMSC-P038190-Vol-2; No Copyright; Avail: CASI; A99, Hardcopy The sections in this report include: Single Stage to Orbit (SSTO) Design Ground-rules; Operations Issues and Lessons Learned; Vertical-Takeoff/Landing Versus Vertical-Takeoff/Horizontal-Landing; SSTO Design Results; SSTO Simulation Results; SSTO Assessment Results; SSTO Sizing Tool User’s Guide; SSto Turnaround Assessment Report; Ground Operations Assessment First Year Executive Summary; Health Management System Deﬁnition Study; Major TA-2 Presentations; First Lunar Outpost Heavy Lift Launch Vehicle Design and Assessment; and the section, Russian Propulsion Technology Assessment Reports. CASI Heavy Lift Launch Vehicles; Single Stage to Orbit Vehicles; Propulsion; Vertical Takeoff; Vertical Landing; Technology Assessment

19960041001 Lockheed Martin Astronautics, Huntsville, AL USA Advanced transportation system studies. Technical area 2: Heavy lift launch vehicle development, Volume 2, Technical Results Jul. 1995; In English Contract(s)/Grant(s): NAS8-39208 Report No.(s): NASA-CR-201126; NAS 1.26:201126; LMSC-P038190; No Copyright; Avail: CASI; A17, Hardcopy Sections 10 to 13 of the Advanced Transportation System Studies final report are included in this volume. Section 10 contains a copy of an executive summary that was prepared by Lockheed Space Operations Company (LSOC) to document their support to the TA-2 contract during the first-year period of performance of the contract, May 1992 through May 1993. LSOC participated on the TA-2 contract as part of the concurrent engineering launch system definition team, and provided outstanding heavy lift launch vehicle (HLLV) ground operations requirements and concept assessments for Lockheed Missiles and Space Company (LMSC) through an intercompany work transfer as well as providing specific HLLV ground operations assessments at the direction of NASA KSC through KSC funding that was routed to the TA-2 contract. Section 11 contains a copy of a vehicle-independent, launch system health management requirements assessment. The purpose of the assessment was to define both health management requirements and the associated interfaces between a generic advanced transportation

129 system launch vehicle and all related elements of the entire transportation system, including the ground segment. Section 12 presents the major TA-2 presentations provided to summarize the significant results and conclusions that were developed over the course of the contract. Finally, Section 13 presents the design and assessment report on the first lunar outpost heavy lift launch vehicle. CASI Heavy Lift Launch Vehicles; Single Stage to Orbit Vehicles; Launch Vehicle Configurations; Advanced Launch System (Sts); Recoverable Spacecraft; Spacecraft Design

19960040942 Georgia Inst. of Tech., Atlanta, GA USA Options for flight testing rocket-based combined-cycle (RBCC) engines Olds, John; Jul. 03, 1996; In English; 32nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 1-3 Jul. 1996, Atlanta, GA, USA Contract(s)/Grant(s): NAG8-1202 Report No.(s): NASA-CR-201897; NAS 1.26:201897; AIAA Paper 96-2688; Copyright; Avail: CASI; A03, Hardcopy While NASA’s current next-generation launch vehicle research has largely focused on advanced all-rocket single-stage- to-orbit vehicles (i.e. the X-33 and it’s RLV operational follow-on), some attention is being given to advanced propulsion concepts suitable for ‘next-generation-and-a-half’ vehicles. Rocket-based combined-cycle (RBCC) engines combining rocket and airbreathing elements are one candidate concept. Preliminary RBCC engine development was undertaken by the USA in the 1960’s. However, additional ground and flight research is required to bring the engine to technological maturity. This paper presents two options for flight testing early versions of the RBCC ejector scramjet engine. The first option mounts a single RBCC engine module to the X-34 air-launched technology testbed for test flights up to about Mach 6.4. The second option links RBCC engine testing to the simultaneous development of a small-payload (220 lb.) two-stage-to-orbit operational vehicle in the Bantam payload class. This launcher/testbed concept has been dubbed the W vehicle. The W vehicle can also serve as an early ejector ramjet RBCC launcher (albeit at a lower payload). To complement current RBCC ground testing efforts, both flight test engines will use earth-storable propellants for their RBCC rocket primaries and hydrocarbon fuel for their airbreathing modes. Performance and vehicle sizing results are presented for both options. Author Engine Tests; Flight Tests; Supersonic Combustion Ramjet Engines; Rocket Engines; X-34 Reusable Launch Vehicle

19960028975 Lockheed Martin Missiles and Space, Huntsville, AL USA Advanced transportation system studies technical area 2 (TA-2): Heavy lift launch vehicle development, volume 3, Program Cost estimates McCurry, J. B.; Jul. 1995; In English Contract(s)/Grant(s): NAS8-39208 Report No.(s): NASA-CR-201128; NAS 1.26:201128; LMSC P038190; No Copyright; Avail: CASI; A04, Hardcopy The purpose of the TA-2 contract was to provide advanced launch vehicle concept definition and analysis to assist NASA in the identification of future launch vehicle requirements. Contracted analysis activities included vehicle sizing and performance analysis, subsystem concept definition, propulsion subsystem definition (foreign and domestic), ground operations and facilities analysis, and life cycle cost estimation. The basic period of performance of the TA-2 contract was from May 1992 through May 1993. No-cost extensions were exercised on the contract from June 1993 through July 1995. This document is part of the final report for the TA-2 contract. The final report consists of three volumes: Volume 1 is the Executive Summary, Volume 2 is Technical Results, and Volume 3 is Program Cost Estimates. The document-at-hand, Volume 3, provides a work breakdown structure dictionary, user’s guide for the parametric life cycle cost estimation tool, and final report developed by ECON, Inc., under subcontract to Lockheed Martin on TA-2 for the analysis of heavy lift launch vehicle concepts. Derived from text Space Transportation System; Launch Vehicles; Life Cycle Costs; Spacecraft Design; Cost Analysis

19960028972 Lockheed Martin Missiles and Space, Huntsville, AL USA Advanced transportation system studies technical area 2(TA-2): Heavy lift launch vehicle development, volume 1, Executive summary McCurry, J.; Jul. 1995; In English Contract(s)/Grant(s): NAS8-39208 Report No.(s): NASA-CR-201124; NAS 1.26:201124; LMSC PC308190; No Copyright; Avail: CASI; A04, Hardcopy

130 The purpose of the TA-2 contract was to provide advanced launch vehicle concept definition and analysis to assist NASA in the identification of future launch vehicle requirements. Contracted analysis activities included vehicle sizing and performance analysis, subsystem concept definition, propulsion subsystem definition (foreign and domestic), ground operations and facilities analysis, and life cycle cost estimation. This document is part of the final report for the TA-2 contract. The final report consists of three volumes: Volume 1 is the Executive Summary, Volume 2 is Technical Results, and Volume 3 is Program Cost Estimates. The document-at-hand, Volume 1, provides a summary description of the technical activities that were performed over the entire contract duration, covering three distinct launch vehicle definition activities: heavy-lift (300,000 pounds injected mass to low Earth orbit) launch vehicles for the First Lunar Outpost (FLO), medium-lift (50,000-80,000 pounds injected mass to low Earth orbit) launch vehicles, and single-stage-to-orbit (SSTO) launch vehicles (25,000 pounds injected mass to a Space Station orbit). Derived from text Launch Vehicles; Life Cycle Costs; Space Transportation; Low Earth Orbits; Space Stations; Heavy Lift Launch Vehicles

19960027977 Lockheed Martin Missiles and Space, Huntsville, AL USA Advanced Transportation System Studies Technical Area 2 (TA-2) Heavy Lift Launch Vehicle Development Contract, Volume 2, Technical Results Jul. 1995; In English Contract(s)/Grant(s): NAS8-39208 Report No.(s): NASA-CR-201125-Vol-2; NAS 1.26:201125-Vol-2; LMSC-P038190; No Copyright; Avail: CASI; A17, Hardcopy The purpose of the Advanced Transportation System Studies (ATSS) Technical Area 2 (TA-2) Heavy Lift Launch Vehicle Development contract was to provide advanced launch vehicle concept definition and analysis to assist NASA in the identification of future launch vehicle requirements. Contracted analysis activities included vehicle sizing and performance analysis, subsystem concept definition, propulsion subsystem definition (foreign and domestic), ground operations and facilities analysis, and life cycle cost estimation. This document is Volume 2 of the final report for the contract. It provides documentation of selected technical results from various TA-2 analysis activities, including a detailed narrative description of the SSTO concept assessment results, a user’s guide for the associated SSTO sizing tools, an SSTO turnaround assessment report, an executive summary of the ground operations assessments performed during the first year of the contract, a configuration-independent vehicle health management system requirements report, a copy of all major TA-2 contract presentations, a copy of the FLO launch vehicle final report, and references to Pratt & Whitney’s TA-2 sponsored final reports regarding the identification of Russian main propulsion technologies. Derived from text Heavy Lift Launch Vehicles; Single Stage to Orbit Vehicles; Space Transportation; Spacecraft Design

19960027549 Lockheed Martin Astronautics, Huntsville, AL USA Transportation Systems Analyses (TSA): Highly Reusable Space Transportation (HRST). A preliminary programmatic assessment Duffey, Jack; Lowrey, Alan; May 1996; In English Contract(s)/Grant(s): NAS8-39209 Report No.(s): NASA-CR-201069; NAS 1.26:201069; No Copyright; Avail: CASI; A03, Hardcopy This report overviews the strategic implications of the Highly Reusable Space Transportation (HRST) program. The analysis postulates the anticipated HRST market (window is 2006-30, with a 2015 focus). Next the analysis speculates on market ‘price of entry’ for several potential markets. HRST is envisioned as a NASA overlay to either the STS modernization or the on-going RLV initiative. Three NASA options are reviewed. An example HRST program (MagLifter + RBCC RLV) is assessed in terms of ﬁnancial/political issues. The merits of HRST-vs-RLV are brieﬂy examined. Finally, a Small Launch Vehicle (SLV) HRST application is reviewed. Author Space Transportation System; Market Research; Reusable Launch Vehicles; Reusable Spacecraft

19960021025 NASA Marshall Space Flight Center, Huntsville,AL USA, American Inst. of Aeronautics and Astronautics, Washington, DC USA Rocket-Induced Magnetohydrodynamic Ejector: A Single-Stage-to-Orbit Advanced Propulsion Concept Cole, John; Campbell, Jonathan; Robertson, Anthony; Sep. 1995; In English; AIAA 1995 Space Programs and Technologies Conference, 26-28 Sep. 1995, Huntsville, AL, USA Report No.(s): NASA-TM-111435; NAS 1.15:111435; AIAA Paper 95-4079; Copyright; Avail: CASI; A03, Hardcopy

131 During the atmospheric boost phase of a rocket trajectory, magnetohydrodynamic (MHD) principles can be utilized to augment the thrust by several hundred percent without the input of additional energy. The concept is an MHD implementation of a thermodynamic ejector. Some ejector history is described and some test data showing the impressive thrust augmentation capabilities of thermodynamic ejectors are provided. A momentum and energy balance is used to derive the equations to predict the MHD ejector performance. Results of these equations are compared with the test data and then applied to a speciﬁc performance example. The rocket-induced MHD ejector (RIME) engine is described and a status of the technology and availability of the engine components is provided. A top level vehicle sizing analysis is performed by scaling existing MHD designs to the required ﬂight vehicle levels. The vehicle can achieve orbit using conservative technology. Modest improvements are suggested using recently developed technologies, such as superconducting magnets, which can improve predicted performance well beyond those expected for current single-stage-to-orbit (SSTO) designs. Author Magnetohydrodynamics; Single Stage to Orbit Vehicles; Rocket Engines; Ejectors; Thrust Augmentation

19960020534 NASA, Hampton, VA USA Development of a Pressure Box to Evaluate Reusable-Launch-Vehicle Cryogenic-Tank Panels Ambur, Damodar R.; Sikora, Joseph; Maguire, James F.; Winn, Peter M.; Apr. 1996; In English; AIAA/ASME/ASCE/AHS/ ASC 37th Structures, Structural Dynamics, and Materials Conference, 15-17 Apr. 1996, Salt Lake City, UT, USA Report No.(s): NASA-TM-111452; NAS 1.15:111452; AIAA Paper 96-1640; Copyright; Avail: CASI; A03, Hardcopy A cryogenic pressure-box test machine has been designed and is being developed to test full-scale reusable launch vehicle cryogenic-tank panels. This machine is equipped with an internal pressurization system, a cryogenic cooling system, and a heating system to simulate the mechanical and thermal loading conditions that are representative of a reusable launch vehicle mission profile. The cryogenic cooling system uses liquid helium and liquid nitrogen to simulate liquid hydrogen and liquid oxygen tank internal temperatures. A quartz lamp heating system is used for heating the external surface of the test panels to simulate cryogenic-tank external surface temperatures during re-entry of the launch vehicle. The pressurization system uses gaseous helium and is designed to be controlled independently of the cooling system. The tensile loads in the axial direction of the test panel are simulated by means of hydraulic actuators and a load control system. The hoop loads in the test panel are reacted by load-calibrated turnbuckles attached to the skin and frame elements of the test panel. The load distribution in the skin and frames can be adjusted to correspond to the tank structure by using these turnbuckles. The seal between the test panel and the cryogenic pressure box is made from a reinforced Teflon material which can withstand pressures greater than 52 psig at cryogenic temperatures. Analytical results and tests on prototype test components indicate that most of the cryogenic-tank loading conditions that occur in flight can be simulated in the cryogenic pressure-box test machine. Author Reusable Launch Vehicles; Cryogenic Fluid Storage; Fuel Tanks; Storage Tanks; Panels; Pressure Vessels; Load Tests; Thermal Stresses; Tensile Stress; Seals (Stoppers)

19960017710 Rockwell International Corp., Canoga Park, CA USA Alternate Propulsion Subsystem Concepts Tripropellant Comparison Study Levack, Daniel; Oct. 1995; In English Contract(s)/Grant(s): NAS8-39210 Report No.(s): NASA-CR-200703; NAS 1.26:200703; 1-1-PP-02147; No Copyright; Avail: CASI; A10, Hardcopy A study was conducted under MSFC contract NAS8-39210 to compare tripropellant and bipropellant engine configurations for the SSTO mission. The objective was to produce an ‘apples-to-apples’ comparison to isolate the effects of design implementation, designing company, year of design, or technologies included from the basic tripropellant/bipropellant comparison. Consequently, identical technologies were included (e.g., jet pumps) and the same design groundrules and practices were used. Engine power cycles were examined as were turbomachinery/preburner arrangements for each cycle. The bipropellant approach and two tripropellant approaches were separately optimized in terms of operating parameters: exit pressures, mixture ratios, thrust splits, etc. This briefing presents the results of the study including engine weights for both tripropellant and bipropellant engines; dry vehicle weight performance for a range of engine chamber pressures; discusses the basis for the results; examines vehicle performance due to engine cycles and the margin characteristics of various cycles; and identifies technologies with significant payoffs for this application. Author Propulsion System Performance; Liquid Propellant Rocket Engines; Propulsion System Configurations; Design Analysis

132 19960017671 NASA Langley Research Center, Hampton, VA USA, Applied Engineering Technologies Ltd., Westborough, MA USA Development of a Pressure Box to Evaluate Reusable-Launch-Vehicle Cryogenic-Tank Panels Ambur, Damodar R.; Sikora, Joseph; Maguire, James F.; Winn, Peter M.; 1996; In English; AIAA/ASME/ASCE/AHS/ASC 37th Structures, Structural Dynamics, and Materials Conference, 15-17 Apr. 1996, Salt Lake City, UT, USA Report No.(s): NASA-TM-111406; NAS 1.15:111406; AIAA Paper 96-1640; No Copyright; Avail: CASI; A03, Hardcopy A cryogenic pressure-box test machine has been designed and is being developed to test full-scale reusable-launch-vehicle cryogenic-tank panels. This machine is equipped with an internal pressurization system, a cryogenic cooling system, and a heating system to simulate the mechanical and thermal loading conditions that are representative of a reusable-launch-vehicle mission profile. The cryogenic cooling system uses liquid helium and liquid nitrogen to simulate liquid hydrogen and liquid oxygen tank internal temperatures. A quartz lamp heating system is used for heating the external surface of the test panels to simulate cryogenic-tank external surface temperatures during re-entry of the launch vehicle. The pressurization system uses gaseous helium and is designed to be controlled independently of the cooling system. The tensile loads in the axial direction of the test panel are simulated by means of hydraulic actuators and a load control system. The hoop loads in the test panel are reacted by load-calibrated turnbuckles attached to the skin and frame elements of the test panel. The load distribution in the skin and frames can be adjusted to correspond to the tank structure by using these turnbuckles. The seal between the test panel and the cryogenic pressure box is made from a reinforced Teflon material which can withstand pressures greater than 52 psig at cryogenic temperatures. Analytical results and tests on prototype test components indicate that most of the cryogenic-tank loading conditions that occur in flight can be simulated in the cryogenic pressure-box test machine. Author (AIAA) Reusable Launch Vehicles; Storage Tanks; Cryogenic Fluid Storage; Load Tests; Panels; Teflon (Trademark)

19960015940 NASA Ames Research Center, Moffett Field, CA USA Surface Catalytic Efficiency of Advanced Carbon Carbon Candidate Thermal Protection Materials for SSTO Vehicles Stewart, David A.; Feb. 01, 1996; In English Contract(s)/Grant(s): RTOP 242-20-01 Report No.(s): NASA-TM-110383; A-961133; NAS 1.15:110383; No Copyright; Avail: CASI; A03, Hardcopy The catalytic efficiency (atom recombination coefficients) for advanced ceramic thermal protection systems was calculated using arc-jet data. Coefficients for both oxygen and nitrogen atom recombination on the surfaces of these systems were obtained to temperatures of 1650 K. Optical and chemical stability of the candidate systems to the high energy hypersonic ﬂow was also demonstrated during these tests. Author Thermal Protection; Recombination Coeffıcient; Oxygen Atoms; Nitrogen Atoms; Carbon-Carbon Composites; Surface Properties; Catalytic Activity

19960015880 Smith Advanced Technology, Inc., Huntsville,AL USA Advanced Propulsion System Test and Tell Capability Horn, Albert E.; Jan. 1996; In English Contract(s)/Grant(s): NAS3-27580; RTOP 244-01-00 Report No.(s): NASA-CR-198415; E-9979; NAS 1.26:198415; No Copyright; Avail: CASI; A03, Hardcopy The feasibility of a ‘test and tell’ capability that performs real-time (or near real-time) diagnostic and prognostic health monitoring for liquid rocket propulsion systems, utilizing a synthesized mathematical modeling expert system technology, was investigated. Such a system would have the potential for reducing operational costs of a launch vehicle propulsion system by reducing the post-firing analysis time and personnel resources needed to determine component health and subsystem replacement requirements. A candidate architecture was defined consisting of propulsion system and subsystem models executing on a flight-avionics based platform. Related work was analyzed, supporting technologies such as NASA Lewis Research Center’s (LeRC’s) ‘feature extraction’ were studied, and the proposed commercial modeling package, Abductive Information Models (AIM), was investigated. The space shuttle main engine (SSME) was selected as a representative propulsion system for development of actual models. Example line replaceable units (LRU’s) and their appropriate sensor measurements were defined. A sample SSME subsystem, the high pressure fuel turbopump (HPFTP), was selected, models developed, and abductive polynomials were synthesized with the AIM tool using training data from actual SSME test firings provided by NASA LeRC. The models were then run with additional test data inputs representing both nominal data and off-nominal data. Evaluation of the results led to the conclusion that the ‘test and tell’ (and eventually ‘land and tell’ for

133 reusable launch vehicles) capability was in fact feasible, and could be developed into a commercially-viable product. Derived from text Expert Systems; Systems Health Monitoring; Reusable Launch Vehicles; Failure Analysis; Liquid Propellant Rocket Engines; Mathematical Models; Test Firing

19960014875 Dayton Univ., OH USA O/S analysis of conceptual space vehicles, Part 1 Ebeling, Charles E.; Dec. 31, 1995; In English Contract(s)/Grant(s): NAG1-1327 Report No.(s): NASA-CR-199465; NAS 1.26:199465; No Copyright; Avail: CASI; A05, Hardcopy The application of recently developed computer models in determining operational capabilities and support requirements during the conceptual design of proposed space systems is discussed. The models used are the reliability and maintainability (R&M) model, the maintenance simulation model, and the operations and support (O&S) cost model. In the process of applying these models, the R&M and O&S cost models were updated. The more signiﬁcant enhancements include (1) improved R&M equations for the tank subsystems, (2) the ability to allocate schedule maintenance by subsystem, (3) redeﬁned spares calculations, (4) computing a weighted average of the working days and mission days per month, (5) the use of a position manning factor, and (6) the incorporation into the O&S model of new formulas for computing depot and organizational recurring and nonrecurring training costs and documentation costs, and depot support equipment costs. The case study used is based upon a winged, single-stage, vertical-takeoff vehicle (SSV) designed to deliver to the Space Station Freedom (SSF) a 25,000 lb payload including passengers without a crew. CASI Cost Analysis; Maintainability; Spacecraft Reliability; Support Systems; Spacecraft Design; Operations Research; Reliability Analysis; Life Cycle Costs; Spacecraft Maintenance

19960014829 NASA Marshall Space Flight Center, Huntsville, AL USA Design of launch systems using continuous improvement process Brown, Richard W.; Sep. 28, 1995 Report No.(s): NASA-TM-111266; NAS 1.15:111266; Copyright; Avail: CASI; A03, Hardcopy The purpose of this paper is to identify a systematic process for improving ground operations for future launch systems. This approach is based on the Total Quality Management (TQM) continuous improvement process. While the continuous improvement process is normally identiﬁed with making incremental changes to an existing system, it can be used on new systems if they use past experience as a knowledge base. In the case of the Reusable Launch Vehicle (RLV), the Space Shuttle operations provide many lessons. The TQM methodology used for this paper will be borrowed from the USA Air Force ‘Quality Air Force’ Program. There is a general overview of the continuous improvement process, with concentration on the formulation phase. During this phase critical analyses are conducted to determine the strategy and goals for the remaining development process. These analyses include analyzing the mission from the customers point of view, developing an operations concept for the future, assessing current capabilities and determining the gap to be closed between current capabilities and future needs and requirements. A brief analyses of the RLV, relative to the Space Shuttle, will be used to illustrate the concept. Using the continuous improvement design concept has many advantages. These include a customer oriented process which will develop a more marketable product and a better integration of operations and systems during the design phase. But, the use of TQM techniques will require changes, including more discipline in the design process and more emphasis on data gathering for operational systems. The beneﬁts will far outweigh the additional effort. Author Total Quality Management; Launching; Ground Operational Support System

19960014828 NASA Langley Research Center, Hampton, VA USA Optimal technology investment strategies for a reusable launch vehicle Moore, A. A.; Braun, R. D.; Powell, R. W.; Sep. 28, 1995 Report No.(s): NASA-TM-111267; AIAA Paper 95-3610; NAS 1.15:111267; Copyright; Avail: CASI; A02, Hardcopy Within the present budgetary environment, developing the technology that leads to an operationally efficient space transportation system with the required performance is a challenge. The present research focuses on a methodology to determine high payoff technology investment strategies. Research has been conducted at Langley Research Center in which design codes for the conceptual analysis of space transportation systems have been integrated in a multidisciplinary design

134 optimization approach. The current study integrates trajectory, propulsion, weights and sizing, and cost disciplines where the effect of technology maturation on the development cost of a single stage to orbit reusable launch vehicle is examined. Results show that the technology investment prior to full-scale development has a signiﬁcant economic payoff. The design optimization process is used to determine strategic allocations of limited technology funding to maximize the economic payoff. Author Multidisciplinary Design Optimization; Recoverable Launch Vehicles; Single Stage to Orbit Vehicles; Technology Assessment; Technology Utilization; NASA Space Programs; Applications Programs (Computers); Computer Aided Design; Design Analysis

19960013899 National Academy of Sciences - National Research Council, Washington, DC USA Reusable launch vehicle: Technology development and test program Jan. 01, 1995 Report No.(s): NASA-CR-200228; LC-95-73106; NAS 1.26:200228; NAS 1.26:200228; Copyright; Avail: CASI; A05, Hardcopy The National Aeronautics and Space Administration (NASA) requested that the National Research Council (NRC) assess the Reusable Launch Vehicle (RLV) technology development and test programs in the most critical component technologies. At a time when discretionary government spending is under close scrutiny, the RLV program is designed to reduce the cost of access to space through a combination of robust vehicles and a streamlined infrastructure. Routine access to space has obvious benefits for space science, national security, commercial technologies, and the further exploration of space. Because of technological challenges, knowledgeable people disagree about the feasibility of a single-stage-to-orbit (SSTO) vehicle. The purpose of the RLV program proposed by NASA and industry contractors is to investigate the status of existing technology and to identify and advance key technology areas required for development and validation of an SSTO vehicle. This report does not address the feasibility of an SSTO vehicle, nor does it revisit the roles and responsibilities assigned to NASA by the National Transportation Policy. Instead, the report sets forth the NRC committee’s findings and recommendations regarding the RLV technology development and test program in the critical areas of propulsion, a reusable cryogenic tank system (RCTS), primary vehicle structure, and a thermal protection system (TPS). Derived from text Reusable Launch Vehicles; NASA Programs; Cryogenic Fluid Storage; Thermal Protection; Composite Structures; Propulsion System Configurations; Spacecraft Design; Spacecraft Configurations

19960012497 NASA Langley Research Center, Hampton, VA, USA Analytical comparison of three stiffened panel concepts Maloney, Jill M.; Wu, K. Chauncey; Robinson, James C.; Dec 1, 1995; In English Contract(s)/Grant(s): RTOP 242-10-01-01 Report No.(s): NASA-TM-110165; NAS 1.15:110165; NIPS-96-08490; No Copyright; Avail: CASI; A03, Hardcopy Three stiffened panel concepts are evaluated to find optimized designs for integral stiffeners in the barrels of Reusable Launch Vehicle fuel tanks. The three panel concepts considered are a T-stiffened panel, a panel with one blade stiffener centered between each pair of T-stiffeners, and a panel with two blade stiffeners equally spaced between each pair of T-stiffeners. The panels are optimized using PASCO for a range of compressive loads, and the computed areal weight for each panel is used to compare the concepts and predict tank weights. The areal weight of the T-stiffened panel with one blade is up to seven-percent lower than the other panel concepts. Two tank construction methods are compared for a representative tank design with three barrels. In the first method, 45-degree circumferential sections of a barrel are each designed to carry the same maximum load in the barrel. In the second method, each barrel section is designed for the maximum load in that section. Representative tanks designed with the first method are over 250 lb heavier than tanks designed using the second method. Optimized panel designs and areal weights are also computed for variation of the nominal panel length and skin thickness. Author Compression Loads; Design Analysis; Fuel Tanks; Panels; Reinforced Plates; Reusable Launch Vehicles; Structural Members

19960007723 North Carolina State Univ., Raleigh, NC, USA Highly reusable space transportation: Approaches for reducing ETO launch costs to $100 - $200 per pound of payload Olds, John R.; Jan 1, 1995; In English Report No.(s): NASA-CR-199561; NAS 1.15:199561; NIPS-95-05542; No Copyright; Avail: CASI; A04, Hardcopy

135 The Commercial Space Transportation Study (CSTS) suggests that considerable market expansion in earth-to-orbit transportation would take place if current launch prices could be reduced to around $400 per pound of payload. If these low prices can be achieved, annual payload delivered to low earth orbit (LEO) is predicted to reach 6.7 million pounds. The primary market growth will occur in communications, government missions, and civil transportation. By establishing a cost target of $100-$200 per pound of payload for a new launch system, the Highly Reusable Space Transportation (HRST) program has clearly set its sights on removing the current restriction on market growth imposed by today’s high launch costs. In particular, achieving the goal of $100-$200 per pound of payload will require signiﬁcant coordinated efforts in (1) marketing strategy development, (2) business planning, (3) system operational strategy, (4) vehicle technical design, and (5) vehicle maintenance strategy. Derived from text Cost Reduction; Low Earth Orbits; Payloads; Reusable Launch Vehicles; Space Commercialization; Space Transportation; Spacecraft Launching

19960003439 NASA Marshall Space Flight Center, Huntsville, AL, USA A guidance and control assessment of three vertical landing options for RLV Gallaher, M.; Coughlin, D.; Krupp, D; Sep 1, 1995; In English Report No.(s): NASA-TM-108500; NAS 1.15:108500; NIPS-95-05719; No Copyright; Avail: CASI; A04, Hardcopy The National Aeronautics and Space Administration is considering a vertical lander as a candidate concept for a single-stage-to-orbit reusable launch vehicle (RLV). Three strategies for guiding and controlling the inversion of a reentering RLV from a nose-ﬁrst attitude to a vertical landing attitude are suggested. Each option is simulated from a common reentry state to touchdown, using a common guidance algorithm and different controllers. Results demonstrate the characteristics that typify and distinguish each concept and help to identify peculiar problems, level of guidance and control sophistication required, feasibility concerns, and areas in which stringent subsystem requirements will be imposed by guidance and control. Author Command and Control; Command Guidance; Control Systems Design; Controllers; Single Stage to Orbit Vehicles; Space Transportation; Spacecraft Control; Technology Assessment; Vertical Landing

19960003337 Eloret Corp., Palo Alto, CA, USA Hypersonic flows as related to the National Aerospace Plane Kussoy, Marvin; Huang, George; Menter, Florian; May 9, 1995; In English Contract(s)/Grant(s): NCC2-452 Report No.(s): NASA-CR-199365; NAS 1.26:199365; No Copyright; Avail: CASI; A03, Hardcopy The object of Cooperative Agreement NCC2-452 was to identify, develop, and document reliable turbulence models for incorporation into CFD codes, which would then subsequently be incorporated into numerical design procedures for the NASP and any other hypersonic vehicles. In a two-pronged effort, consisting of an experimental and a theoretical approach, several key features of flows over complex vehicles were identified, and test bodies were designed which were composed of simple geometric shapes over which these flow features were measured. The experiments were conducted in the 3.5’ Hypersonic Wind Tunnel at NASA Ames Research Center, at nominal Mach numbers from 7 to 8.3 and Re/m from 4.9 x 10(exp 6) to 5.8 x 10(exp 6). Boundary layers approaching the interaction region were 2.5 to 3.7 cm thick. Surface and flow field measurements were conducted, and the initial boundary conditions were experimentally documented. Author Aerospace Planes; Boundary Layers; Computational Fluid Dynamics; Flow Distribution; Flow Measurement; Hypersonic Flow; Hypersonic Speed; Shock Wave Interaction; Turbulence Models; Wind Tunnel Tests

19960003127 NASA Langley Research Center, Hampton, VA, USA Single-stage-to-orbit: Meeting the challenge Freeman, Delma C., Jr.; Talay, Theodore A.; Austin, Robert Eugene; Oct 6, 1995; In English; 46th International Astronautical Congress, 2-6 Oct. 1995, Oslo, Norway Report No.(s): NASA-TM-111127; NIPS-95-05489; NAS 1.15:111127; IAF-95-V-5-07; No Copyright; Avail: CASI; A03, Hardcopy There has been and continues to be signiﬁcant discussion about the viability of fully reusable, single-stage-to-orbit (SSTO) concepts for delivery of payloads to orbit. Often, these discussions have focused in detail on performance and technology requirements relating to the technical feasibility of the concept, with only broad generalizations on how the SSTO

136 will achieve its economic goals of greatly reduced vehicle ground and flight operations costs. With the current industry and NASA Reusable Launch Vehicle Technology Program efforts underway to mature and demonstrate technologies leading to a viable commercial launch system that also satisfies national needs, achieving acceptable recurring costs becomes a significant challenge. This paper reviews the current status of the Reusable Launch Vehicle Technology Program including the DC-XA, X-33, and X-34 flight systems and associated technology programs. The paper also examines lessons learned from the recently completed DC-X reusable rocket demonstrator program. It examines how these technologies and flight systems address the technical and operability challenges of SSTO whose solutions are necessary to reduce costs. The paper also discusses the management and operational approaches that address the challenge of a new cost-effective, reusable launch vehicle system. Author Cost Effectiveness; Flight Operations; Single Stage to Orbit Vehicles; Space Transportation; Technology Assessment

19960003035 Joint Inst. for Advancement of Flight Sciences, Hampton, VA, USA Abort performance for a winged-body single-stage to orbit vehicle Lyon, Jeffery A.; Aug 1, 1995; In English Contract(s)/Grant(s): NCC1-104; RTOP 242-80-01-02 Report No.(s): NASA-CR-4690; NAS 1.26:4690; No Copyright; Avail: CASI; A05, Hardcopy Optimal control theory is employed to determine the performance of abort to orbit (ATO) and return to launch site (RTLS) maneuvers for a single-stage to orbit vehicle. The vehicle conﬁguration examined is a seven engine, winged-body vehicle, that lifts-off vertically and lands horizontally. The abort maneuvers occur as the vehicle ascends to orbit and are initiated when the vehicle suffers an engine failure. The optimal control problems are numerically solved in discretized form via a nonlinear programming (NLP) algorithm. A description highlighting the attributes of this NLP method is provided. ATO maneuver results show that the vehicle is capable of ascending to orbit with a single engine failure at lift-off. Two engine out ATO maneuvers are not possible from the launch pad, but are possible after launch when the thrust to weight ratio becomes sufficiently large. Results show that single engine out RTLS maneuvers can be made for up to 180 seconds after lift-off and that there are scenarios for which RTLS maneuvers should be performed instead of ATP maneuvers. Author Abort Trajectories; Aerodynamic Characteristics; Engine Failure; Multiengine Vehicles; Single Stage to Orbit Vehicles; Winged Vehicles

19960002345 NASA Lewis Research Center, Cleveland, OH, USA Human lunar mission capabilities using SSTO, ISRU and LOX-augmented NTR technologies: A preliminary assessment Borowski, Stanley K.; Oct 1, 1995; In English; 31st Joint Propulsion Conference and Exhibit, 10-12 Jul. 1995, San Diego, CA, USA Contract(s)/Grant(s): RTOP 242-10-01 Report No.(s): NASA-TM-107095; NAS 1.15:107095; E-9973; AIAA PAPER 95-2631; No Copyright; Avail: CASI; A03, Hardcopy The feasibility of conducting human missions to the Moon is examined assuming the use of three ‘high leverage’ technologies: (1) a single-stage-to-orbit (SSTO) launch vehicle, (2) ‘in-situ’ resource utilization (ISRU)--specifically ‘lunar-derived’ liquid oxygen (LUNOX), and (3) LOX-augmented nuclear thermal rocket (LANTR) propulsion. Lunar transportation system elements consisting of a LANTR-powered lunar transfer vehicle (LTV) and a chemical propulsion lunar landing/Earth return vehicle (LERV) are configured to fit within the ‘compact’ dimensions of the SSTO cargo bay (diameter: 4.6 m/length: 9.0 m) while satisfying an initial mass in low Earth orbit (IMLEO) limit of approximately 60 t (3 SSTO launches). Using approximately8tofLUNOX to ‘reoxidize’ the LERV for a ‘direct return’ flight to Earth reduces its size and mass allowing delivery to LEO on a single 20 t SSTO launch. Similarly, the LANTR engine’s ability to operate at any oxygen/ hydrogen mixture ratio from 0 to 7 with high specific impulse (approximately 940 to 515 s) is exploited to reduce hydrogen tank volume, thereby improving packaging of the LANTR LTV’s ‘propulsion’ and ‘propellant modules’. Expendable and reusable, piloted and cargo missions and vehicle designs are presented along with estimates of LUNOX production required to support the different mission modes. Concluding remarks address the issue of lunar transportation system costs from the launch vehicle perspective. Author Chemical Propulsion; Low Earth Orbits; Lunar Landing; Moon; Nuclear Rocket Engines; Propulsion; Rocket Engines; Single Stage to Orbit Vehicles

137 19960000793 Santa Clara Univ., CA, USA Advances in hypersonic vehicle synthesis with application to studies of advanced thermal protection system Ardema, Mark D.; Sep 30, 1995; In English Contract(s)/Grant(s): NCC2-5069 Report No.(s): NASA-CR-199333; NAS 1.26:199333; No Copyright; Avail: CASI; A04, Hardcopy This report summarizes the work entitled ‘Advances in Hypersonic Vehicle Synthesis with Application to Studies of Advanced Thermal Protection Systems.’ The effort was in two areas: (1) development of advanced methods of trajectory and propulsion system optimization; and (2) development of advanced methods of structural weight estimation. The majority of the effort was spent in the trajectory area. Derived from text Engine Design; Hypersonic Vehicles; Propulsion System Conﬁgurations; Single Stage to Orbit Vehicles; Supersonic Combustion Ramjet Engines; Thermal Protection; Trajectory Optimization

19950028302 NASA Ames Research Center, Moffett Field, CA, USA Air-breathing aerospace plane development essential: Hypersonic propulsion flight tests Mehta, Unmeel B.; ESA, Proceedings of the 2nd European Symposium on Aerothermodynamics for Space Vehicles; Feb 1, 1995; In English Report No.(s): NASA-TM-108857-REV; NAS 1.15:108857-REV; Copyright; Avail: CASI; A03, Hardcopy Hypersonic airbreathing propulsion utilizing scramjets can change transatmospheric accelerators for low earth-to-orbit and return transportation. The value and limitation of ground tests, of flight tests, and of computations are presented, and scramjet development requirements are discussed. It is proposed that near full-scale hypersonic propulsion flight tests are essential for developing computational design technology so that it can be used for designing this system. In order to determine how these objectives should be achieved, some lessons learned from past programs are presented. A conceptual two-stage-to-orbit (TSTO) prototype/experimental aerospace plane is recommended as a means of providing access-to-space and for conducting flight tests. Author Aerospace Planes; Engine Tests; Flight Tests; Full Scale Tests; Hypersonic Speed; Spacecraft Propulsion; Supersonic Combustion Ramjet Engines

19950026746 NASA, Washington, DC, USA Shuttle to Space Station. Heart assist implant. Hubble update. X-30 mock-up Aug 1, 1992; In English Report No.(s): NASA-TM-110837; NONP-NASA-VT-95-63907; No Copyright; Avail: CASI; B01, Videotape-Beta; V01, Videotape-VHS Shuttle to Space Station, Heart Assist Implant, Hubble Update, and X-30 Mockup are the four parts that are discussed in this video. The ﬁrst part, Shuttle to Space Station, is focussed on the construction and function of the Space Station Freedom. While part two, Heart Assist Implant, discusses a newly developed electromechanical device that helps to reduce heart attack by using electric shocks. Interviews with the co-inventor and patients are also included. Brief introduction to Hubble Telescope, problem behind its poor image quality (mirror aberration), and the plan to correct this problem are the three issues that are discussed in part three, Hubble Update. The last part, part four, reviews the X-30 Mockup designed by the staff and students of Mississippi State University. CASI Cardiovascular System; Heart Diseases; Hubble Space Telescope; Space Station Freedom; X-30 Vehicle

19950024291 NASA Marshall Space Flight Center, Huntsville, AL, USA Transonic aerodynamic characteristics of a proposed wing-body reusable launch vehicle concept Springer, A. M.; Mar 1, 1995; In English Report No.(s): NASA-TM-108489; NAS 1.15:108489; No Copyright; Avail: CASI; A06, Hardcopy A proposed wing-body reusable launch vehicle was tested in the NASA Marshall Space Flight Center’s 14 x 14-inch trisonic wind tunnel during the winter of 1994. This test resulted in the vehicle’s subsonic and transonic, Mach 0.3 to 1.96, longitudinal and lateral aerodynamic characteristics. The effects of control surface deﬂections on the basic vehicle’s

138 aerodynamics, including a body flap, elevons, ailerons, and tip fins, are presented. Author Aerodynamic Characteristics; Aerodynamic Forces; Aerodynamic Stability; Angle of Attack; Body-Wing Configurations; Control Surfaces; Lateral Stability; Longitudinal Stability; Mach Number; Sideslip; Single Stage to Orbit Vehicles; Spacecraft Stability; Subsonic Speed; Transonic Speed; Wind Tunnel Tests

19950021839 Rockwell International Corp., Downey, CA, USA Task 4 supporting technology. Densification requirements definition and test objectives. Propellant densification requirements definition Lak, Tibor; Weeks, D. P.; May 30, 1995; In English Contract(s)/Grant(s): NCC8-79 Report No.(s): NASA-CR-198629; NAS 1.26:198629; SSD95D0241; No Copyright; Avail: CASI; A03, Hardcopy The primary challenge of the X-33 CAN is to build and test a prototype LO2 and LH2 densification ground support equipment (GSE) unit, and perform tank thermodynamic testing within the 15 month phase 1 period. The LO2 and LH2 propellant densification system will be scaled for the IPTD LO2 and LH2 tank configurations. The IPTD tanks were selected for the propellant technology demonstration because of the potential benefits to the phase 1 plan: tanks will be built in time to support thermodynamic testing; minimum cost; minimum schedule risk; future testing at MSFC will build on phase 1 data base; and densification system will be available to support X-33 and RLV engine test at IPTD. The objective of the task 1 effort is to define the preliminary requirements of the propellant densification GSE and tank recirculation system. The key densification system design parameters to be established in Task 1 are: recirculation flow rate; heat exchanger inlet temperature; heat exchanger outlet temperature; maximum heat rejection rate; vent flow rate (GN2 and GH2); densification time; and tank pressure level. Derived from text Densification; Heat Exchangers; Liquid Hydrogen; Liquid Oxygen; Liquid Rocket Propellants; Propellant Storage; Reusable Launch Vehicles

19950021838 Rockwell International Corp., Downey, CA, USA Task 4 supporting technology. Part 2: Detailed test plan for thermal seals. Thermal seals evaluation, improvement and test. CAN8-1, Reusable Launch Vehicle (RLV), advanced technology demonstrator: X-33. Leading edge and seals thermal protection system technology demonstration Hogenson, P. A.; Lu, Tina; May 30, 1995; In English Contract(s)/Grant(s): NCC8-79 Report No.(s): NASA-CR-198628; NAS 1.26:198628; SSD95D0240-PT-2; No Copyright; Avail: CASI; A03, Hardcopy The objective is to develop the advanced thermal seals to a technology readiness level (TRL) of 6 to support the rapid turnaround time and low maintenance requirements of the X-33 and the future reusable launch vehicle (RLV). This program is divided into three subtasks: (1) orbiter thermal seals operation history review; (2) material, process, and design improvement; and (3) fabrication and evaluation of the advanced thermal seals. CASI Compression Tests; High Temperature Tests; Reentry Shielding; Reusable Launch Vehicles; Sealers; Technology Assessment; Thermal Insulation; Thermal Protection

19950021837 Rockwell International Corp., Downey, CA, USA Task 4 supporting technology. Part 1: Detailed test plan for leading edge tile development. Leading edge material development and testing Hogenson, P. A.; Staszak, Paul; Hinkle, Karrie; May 30, 1995; In English Contract(s)/Grant(s): NCC8-79 Report No.(s): NASA-CR-198633; NAS 1.26:198633; SSD95D0239-PT-1; No Copyright; Avail: CASI; A03, Hardcopy This task develops two alternative candidate tile materials for leading edge applications: coated alumina enhanced thermal barrier (AETB) tile and silicone impregnated reusable ceramic ablator (SIRCA) tile. Upon reentry of the X-33/RLV space vehicle, the leading edges experience the highest heating rates and temperatures. The wing leading edge and nose cap experience peak temperatures in the range 2000 to 2700 F. Replacing reinforced carbon-carbon (RCC) with tile-based thermal protection system (TPS) materials is the primary objective. Weight, complexity, coating impact damage, and repairability are among the problems that this tile technology development addresses. The following subtasks will be performed in this

139 development effort: tile coating development; SIRCA tile development; robustness testing of tiles; tile repair development; tile operations/processing; tile leading edge conﬁguration; and life cycle testing. CASI Ablative Materials; Aluminum Oxides; Ceramics; Coatings; Leading Edges; Reentry Shielding; Reusable Heat Shielding; Reusable Launch Vehicles; Silicones; Thermal Protection; Tiles

19950020825 GPS Solutions, Inc., Carson City, NV, USA Reusable launch vehicle development research Jan 16, 1995; In English Contract(s)/Grant(s): NAS8-40257 Report No.(s): NASA-CR-196614; NAS 1.26:196614; No Copyright; Avail: CASI; A03, Hardcopy NASA has generated a program approach for a SSTO reusable launch vehicle technology (RLV) development which includes a follow-on to the Ballistic Missile Defense Organization’s (BMDO) successful DC-X program, the DC-XA (Advanced). Also, a separate sub-scale ﬂight demonstrator, designated the X-33, will be built and ﬂight tested along with numerous ground based technologies programs. For this to be a successful effort, a balance between technical, schedule, and budgetary risks must be attained. The adoption of BMDO’s ‘fast track’ management practices will be a key element in the eventual success of NASA’s effort. Derived from text Ballistic Missiles; Missile Defense; NASA Space Programs; Reusable Launch Vehicles; Single Stage to Orbit Vehicles

19950019884 Vanderbilt Univ., Nashville, TN, USA The 1994 NASA/USRA/ADP Design Projects Cruse, Thomas; Richardson, Joseph; Tryon, Robert; JAN 1, 1994; In English; See also N95-26305 through N95-26310; 7 functional color pages Contract(s)/Grant(s): NASW-4435 Report No.(s): NASA-CR-197415; NAS 1.26:197415; No Copyright; Avail: CASI; A07, Hardcopy; 7 functional color pages The NASA/USRA/ADP Design Projects from Vanderbilt University, Department of Mechanical Engineering (1994) are enclosed in this final report. Design projects include: (1) Protein Crystal Growth, both facilities and methodology; (2) ACES Deployable Space Boom; (3) Hybrid Launch System designs for both manned and unmanned systems; (4) LH2 Fuel Tank design (SSTO); (5) SSTO design; and (6) Pressure Tank Feed System design. Experiment Design; Functional Design Specifications; Government/Industry Relations; Launch Vehicle Configurations; Mechanical Engineering; Pressure Vessel Design; Single Stage to Orbit Vehicles; Spacecraft Design; Structural Design; University Program

19950017766 NASA Lewis Research Center, Cleveland, OH, USA A summary of the slush hydrogen technology program for the National Aero-Space Plane Mcnelis, Nancy B.; Hardy, Terry L.; Whalen, Margaret V.; Kudlac, Maureen T.; Moran, Matthew E.; Tomsik, Thomas M.; Haberbusch, Mark S.; Apr 1, 1995; In English; Hypersonics Technologies Conference, 3-7 Apr. 1995, Chattanooga, TN, USA Contract(s)/Grant(s): RTOP 763-22-21 Report No.(s): NASA-TM-106863; E-9469; NAS 1.15:106863; AIAA PAPER 95-6056; No Copyright; Avail: CASI; A03, Hardcopy Slush hydrogen, a mixture of solid and liquid hydrogen, offers advantages of higher density (16 percent) and higher heat capacity (18 percent) than normal boiling point hydrogen. The combination of increased density and heat capacity of slush hydrogen provided a potential to decrease the gross takeoff weight of the National Aero-Space Plane (NASP) and therefore slush hydrogen was selected as the propellant. However, no large-scale data was available on the production, transfer and tank pressure control characteristics required to use slush hydrogen as a fuel. Extensive testing has been performed at the NASA Lewis Research Center K-Site and Small Scale Hydrogen Test Facility between 1990 and the present to provide a database for the use of slush hydrogen. This paper summarizes the results of this testing. Author Aerospace Planes; Fuel Tests; Hydrogen Fuels; National Aerospace Plane Program; Slush; Slush Hydrogen

140 19950016772 Rockwell International Corp., Downey, CA, USA, Northrop Grumman Corp., Los Angeles, CA, USA, North American Aviation, Inc., Tulsa, OK, USA, Hercules, Inc., Wilmington, DE, USA Structural arrangement trade study. Volume 3: Reusable Hydrogen Composite Tank System (RHCTS) and Graphite Composite Primary Structures (GCPS). Addendum Mar 14, 1995; In English Contract(s)/Grant(s): NCC8-39; NCC1-193 Report No.(s): NASA-CR-197687; NAS 1.26:197687; SSD95D0069-VOL-3; No Copyright; Avail: CASI; A12, Hardcopy This volume is the third of a 3 volume set that addresses the structural trade study plan that will identify the most suitable structural configuration for an SSTO winged vehicle capable of delivering 25,000 lbs to a 220 nm circular orbit at 51.6 deg inclination. The most suitable Reusable Hydrogen Composite Tank System (RHCTS), and Graphite Composite Tank System (GCPS) composite materials for intertank, wing and thrust structures are identified. Vehicle resizing charts, selection criteria and back-up charts, parametric costing approach and the finite element method analysis are discussed. Derived from text Composite Structures; Dynamic Structural Analysis; Finite Element Method; Graphite; Hydrogen Fuels; Single Stage to Orbit Vehicles; Spacecraft Structures

19950016771 Rockwell International Corp., Downey, CA, USA, Northrop Grumman Corp., Los Angeles, CA, USA, North American Aviation, Inc., Tulsa, OK, USA, Hercules, Inc., Wilmington, DE, USA Structural Arrangement Trade Study. Volume 1: Reusable Hydrogen Composite Tank System (RHCTS) and Graphite Composite Primary Structures (GCPS). Executive summary Mar 14, 1995; In English Contract(s)/Grant(s): NCC8-39; NCC1-193 Report No.(s): NASA-CR-197685; NAS 1.26:197685; SSD95D0069-VOL-1; No Copyright; Avail: CASI; A03, Hardcopy This volume is the first of a three volume set that discusses the structural arrangement trade study plan that will identify the most suitable configuration for an SSTO winged vehicle capable of delivering 25,000 lbs to a 220 nm circular orbit at 51.6 deg inclination. The Reusable Hydrogen Composite Tank System (RHCTS), and Graphite Composite Primary Structures most suitable for intertank, wing and thrust structures are identified. This executive summary presents the trade study process, the selection process, requirements used, analysis performed and data generated. Conclusions and recommendations are also presented. Derived from text Composite Structures; Graphite; Hydrogen Fuels; Project Planning; Single Stage to Orbit Vehicles; Spacecraft Structures

19950010487 NASA Langley Research Center, Hampton, VA, USA Constrained minimization of smooth functions using a genetic algorithm Moerder, Daniel D.; Pamadi, Bandu N.; Nov 1, 1994; In English Contract(s)/Grant(s): RTOP 506-59-61-01 Report No.(s): NASA-TP-3329; L-17221; NAS 1.60-3329; No Copyright; Avail: CASI; A03, Hardcopy The use of genetic algorithms for minimization of differentiable functions that are subject to differentiable constraints is considered. A technique is demonstrated for converting the solution of the necessary conditions for a constrained minimum into an unconstrained function minimization. This technique is extended as a global constrained optimization algorithm. The theory is applied to calculating minimum-fuel ascent control settings for an energy state model of an aerospace plane. Author Aerospace Planes; Genetic Algorithms; Lagrange Multipliers; Mathematical Models; Minima; Optimization

19950009268 NASA Lewis Research Center, Cleveland, OH, USA The rationale/beneﬁts of nuclear thermal rocket propulsion for NASA’s lunar space transportation system Borowski, Stanley K.; Sep 1, 1994; In English; 27th Joint Propulsion Conference, 24-26 Jun. 1991, Sacramento, CA, USA Contract(s)/Grant(s): RTOP 232-01-06 Report No.(s): NASA-TM-106739; E-9142; NAS 1.15:106739; AIAA PAPER 91-2052; No Copyright; Avail: CASI; A03, Hardcopy The solid core nuclear thermal rocket (NTR) represents the next major evolutionary step in propulsion technology. With its attractive operating characteristics, which include high speciﬁc impulse (approximately 850-1000 s) and engine thrust-to-weight (approximately 4-20), the NTR can form the basis for an efficient lunar space transportation system (LTS)

141 capable of supporting both piloted and cargo missions. Studies conducted at the NASA Lewis Research Center indicate that an NTR-based LTS could transport a fully-fueled, cargo-laden, lunar excursion vehicle to the Moon, and return it to low Earth orbit (LEO) after mission completion, for less initial mass in LEO than an aerobraked chemical system of the type studied by NASA during its ‘90-Day Study.’ The all-propulsive NTR-powered LTS would also be ‘fully reusable’ and would have a ‘return payload’ mass fraction of approximately 23 percent--twice that of the ‘partially reusable’ aerobraked chemical system. Two NTR technology options are examined--one derived from the graphite-moderated reactor concept developed by NASA and the AEC under the Rover/NERVA (Nuclear Engine for Rocket Vehicle Application) programs, and a second concept, the Particle Bed Reactor (PBR). The paper also summarizes NASA’s lunar outpost scenario, compares relative performance provided by different LTS concepts, and discusses important operational issues (e.g., reusability, engine ‘end-of life’ disposal, etc.) associated with using this important propulsion technology. Author Lunar Bases; Lunar Orbital Rendezvous; Mission Planning; Nuclear Engine For Rocket Vehicles; Nuclear Propulsion; Propulsion System Conﬁgurations; Reusable Launch Vehicles; Space Transportation System; Thrust-Weight Ratio

19950008507 NASA Ames Research Center, Moffett Field, CA, USA Air-breathing aerospace plane development essential: Hypersonic propulsion flight tests Mehta, Unmeel B.; Nov 1, 1994; In English; 2nd European Symposium on Aerothermodynamics for Space Vehicles, ESTEC, 21-23 Nov. 1994, Noordwijk, Netherlands Contract(s)/Grant(s): RTOP 505-59-53 Report No.(s): NASA-TM-108857; A-95017; NAS 1.15:108857; No Copyright; Avail: CASI; A03, Hardcopy Hypersonic air-breathing propulsion utilizing scramjets can fundamentally change transatmospheric accelerators for low earth-to-orbit and return transportation. The value and limitations of ground tests, of flight tests, and of computations are presented, and scramjet development requirements are discussed. It is proposed that near full-scale hypersonic propulsion flight tests are essential for developing a prototype hypersonic propulsion system and for developing computational-design technology so that it can be used for designing this system. In order to determine how these objectives should be achieved, some lessons learned from past programs are presented. A conceptual two-stage-to-orbit (TSTO) prototype/experimental aerospace plane is recommended as a means of providing access-to-space and for conducting flight tests. A road map for achieving these objectives is also presented. Author Aerospace Planes; Flight Tests; Hypersonic Flight; Space Transportation; Supersonic Combustion Ramjet Engines

19950008387 Rockwell International Corp., Downey, CA, USA Milestone 4: Test plan for Reusable Hydrogen Composite Tank System (RHCTS). Task 3: Composite tank materials Greenberg, H. S.; Nov 4, 1994; In English Contract(s)/Grant(s): NCC8-39 Report No.(s): NASA-CR-197036; NAS 1.26:197036; SD94D0329; No Copyright; Avail: CASI; A04, Hardcopy This document is the detailed test plan for the series of tests enumerated in the preceding section. The purpose of this plan is to present the test objectives, test parameters and procedures, expected performance and data analysis plans, criteria for success, test schedules, and related safety provisions and to describe the test articles, test instrumentation, and test facility requirements. Initial testing will be performed to screen four composite materials for suitability for SSTO LH2 tank loads and environmental conditions. The laminates for this testing will be fabricated by fiber placement, which is the manufacturing approach identified as baseline for the tank wall. Even though hand layup will be involved in fabricating many of the internal structural members of the tank, no hand-layup laminates will be evaluated in the screening or subsequent characterization testing. This decision is based on the understanding that mechanical properties measured for hand-layup material should be at least equivalent to properties measured for fiber-placed material, so that the latter should provide no less than a conservative approximation of the former. A single material will be downselected from these screening tests. This material will be subsequently characterized for impact-damage tolerance and durability under conditions of mechanical and thermal cycling, and to establish a preliminary design database to support ongoing analysis. Next, testing will be performed on critical structural elements fabricated from the selected material. Finally, the 8-foot diameter tank article, containing the critical structural features of the full-scale tank, will be fabricated by fiber placement and tested to verify its structural integrity and LH2 containment. Derived from text Fatigue Tests; Impact Tests; Laminates; Liquid Hydrogen; Performance Tests; Propellant Tanks; Single Stage to Orbit Vehicles; Thermal Cycling Tests

142 19950008386 Rockwell International Corp., Downey, CA, USA Milestone 4: Thrust structure concepts and IHM screening Graphite Composite Primary Structure (GCPS) Greenberg, H. S.; Nov 4, 1994; In English Contract(s)/Grant(s): NCC1-193 Report No.(s): NASA-CR-197053; NAS 1.26:197053; SSD94D0328; No Copyright; Avail: CASI; A03, Hardcopy The first part of the task was to select up to three promising thrust structure constructions and to select materials for screening tests. Part of the nondestructive evaluation and inspection (NDE/I) and integrated health management (IHM) task is to acquire and develop NDE/I sensor technologies and to integrate those sensors into the full scale test articles which will be produced under the TA2 program. Review of the anticipated fault modes and the available sensor technology data indicates that three sensor technologies should be assessed for the in-situ monitoring of the composite primary structure elements. These are: ultrasonics (dry contact), acoustic emissions, and fiber optics (embedded or attached). In fact, a combination of sensor technologies will be needed to detect and evaluate the fault modes; not only do sensor technology have specific capabilities and applicability, but the three Gr/Ep primary structures being demonstrated under the TA2 effort have differing requirements based on their respective failure modes and designs. Derived from text Composite Structures; Graphite-Epoxy Composites; Nondestructive Tests; Single Stage to Orbit Vehicles; Spacecraft Structures; Structural Design

19950008380 Rockwell International Corp., Downey, CA, USA Requirements report for SSTO vertical take-off and horizontal landing vehicle Greenberg, H. S.; Nov 4, 1994; In English Contract(s)/Grant(s): NCC1-193; NCC2-9003; NCC8-39 Report No.(s): NASA-CR-197029; NAS 1.26:197029; SSD94D0217B; No Copyright; Avail: CASI; A08, Hardcopy This document describes the detailed design requirements and design criteria to support Structures/TPS Technology development for SSTO winged vehicle conﬁgurations that use vertical take-off and horizontal landing and delivers 25,000 lb payloads to a 220 nm circular orbit at an inclination of 51.6 degrees or 40,000 lb payloads to a 150 nm circular orbit at a 28.5 degree inclination. Derived from text Design Analysis; Launch Vehicle Conﬁgurations; Single Stage to Orbit Vehicles; Spacecraft Design; Structural Design; Thermal Protection

19950008107 Minnesota Univ., Minneapolis, MN, USA Access to space Jul 1, 1994; In English Contract(s)/Grant(s): NASW-4435 Report No.(s): NASA-CR-197153; NAS 1.26:197153; No Copyright; Avail: CASI; A07, Hardcopy The goal of this conceptual design was to devise a reusable, commercially viable, single-stage-to-orbit vehicle. The vehicle has the ability to deliver a 9100 kg (20,000 lb) payload to a low earth orbit of 433 km to 933 km (250 n.mi. - 450 n.mi.). The SSTO vehicle is 51 meters in length and has a gross takeoff mass of 680,400 kg (1,500,000 lb). The vehicle incorporates three RD-701 engines for the main propulsion system and two RL-10 engines for the orbital maneuvering system. The vehicle is designed for a three day stay on orbit with two crew members. Author Aerospace Planes; Launch Vehicle Conﬁgurations; Life Cycle Costs; Propulsion System Conﬁgurations; Single Stage to Orbit Vehicles; Space Commercialization; Space Transportation; Spacecraft Design; Structural Design; Transatmospheric Vehicles; Wind Tunnel Models

19950007252 NASA Lewis Research Center, Cleveland, OH, USA Computational analysis in support of the SSTO ﬂowpath test Duncan, Beverly S.; Trefny, Charles J.; Oct 1, 1994; In English Contract(s)/Grant(s): NAS3-27186; RTOP 505-70-59 Report No.(s): NASA-TM-106757; E-9179; NAS 1.15:106757; No Copyright; Avail: CASI; A03, Hardcopy A synergistic approach of combining computational methods and experimental measurements is used in the analysis of a hypersonic inlet. There are four major focal points within this study which examine the boundary layer growth on a

143 compression ramp upstream of the cowl lip of a scramjet inlet. Initially, the boundary layer growth on the NASP Concept Demonstrator Engine (CDE) is examined. The follow-up study determines the optimum diverter height required by the SSTO Flowpath test to best duplicate the CDE results. These ﬂow ﬁeld computations are then compared to the experimental measurements and the mass average Mach number is determined for this inlet. Author Boundary Layer Flow; Cowlings; Diverters; Engine Inlets; Hypersonic Inlets; Inlet Flow

19950006425 NASA Langley Research Center, Hampton, VA, USA Hypersonic vehicle model and control law development using H(infinity) and micron synthesis Gregory, Irene M.; Chowdhry, Rajiv S.; Mcminn, John D.; Shaughnessy, John D.; Oct 1, 1994; In English Contract(s)/Grant(s): RTOP 505-70-64-01 Report No.(s): NASA-TM-4562; L-17217; NAS 1.15:4562; No Copyright; Avail: CASI; A03, Hardcopy The control system design for a Single Stage To Orbit (SSTO) air breathing vehicle will be central to a successful mission because a precise ascent trajectory will preserve narrow payload margins. The air breathing propulsion system requires the vehicle to fly roughly halfway around the Earth through atmospheric turbulence. The turbulence, the high sensitivity of the propulsion system to inlet flow conditions, the relatively large uncertainty of the parameters characterizing the vehicle, and continuous acceleration make the problem especially challenging. Adequate stability margins must be provided without sacrificing payload mass since payload margins are critical. Therefore, a multivariable control theory capable of explicitly including both uncertainty and performance is needed. The H(infinity) controller in general provides good robustness but can result in conservative solutions for practical problems involving structured uncertainty. Structured singular value mu framework for analysis and synthesis is potentially much less conservative and hence more appropriate for problems with tight margins. An SSTO control system requires: highly accurate tracking of velocity and altitude commands while limiting angle-of-attack oscillations, minimized control power usage, and a stabilized vehicle when atmospheric turbulence and system uncertainty are present. The controller designs using H(infinity) and mu-synthesis procedures were compared. An integrated flight/propulsion dynamic mathematical model of a conical accelerator vehicle was linearized as the vehicle accelerated through Mach 8. Vehicle acceleration through the selected flight condition gives rise to parametric variation that was modeled as a structured uncertainty. The mu-analysis approach was used in the frequency domain to conduct controller analysis and was confirmed by time history plots. Results demonstrate the inherent advantages of the mu framework for this class of problems. Author (revised) Aerodynamic Characteristics; Control Systems Design; Dynamic Models; Hypersonic Vehicles; Mathematical Models; Propulsion System Configurations; Robustness (Mathematics); Single Stage to Orbit Vehicles

19950006252 California Univ., Los Angeles, CA, USA, Universities Space Research Association, Columbia, MD, USA Ion engine propelled Earth-Mars cycler with nuclear thermal propelled transfer vehicle, volume 2 Meyer, Rudolf X.; Baker, Myles; Melko, Joseph; JAN 1, 1994; In English Contract(s)/Grant(s): NASW-4435 Report No.(s): NASA-CR-197208; NAS 1.26:197208; No Copyright; Avail: CASI; A04, Hardcopy The goal of this project was to perform a preliminary design of a long term, reusable transportation system between earth and Mars which would be capable of providing both artiﬁcial gravity and shelter from solar ﬂare radiation. The heart of this system was assumed to be a Cycler spacecraft propelled by an ion propulsion system. The crew transfer vehicle was designed to be propelled by a nuclear-thermal propulsion system. Several Mars transportation system architectures and their associated space vehicles were designed. Author Ion Engines; Ion Propulsion; Manned Mars Missions; Manned Spacecraft; Nuclear Propulsion; Nuclear Rocket Engines; Reusable Spacecraft; Rotating Environments; Space Transportation; Spacecraft Design

19950006217 Vanderbilt Univ., Nashville, TN, USA, Universities Space Research Association, Columbia, MD, USA Various advanced design projects promoting engineering education JAN 1, 1994; In English; 6 functional color pages Contract(s)/Grant(s): NASW-4435 Report No.(s): NASA-CR-197170; NAS 1.26:197170; No Copyright; Avail: CASI; A07, Hardcopy; 6 functional color pages The Universities Space Research Association (USRA) Advanced Design Program (ADP) program promotes engineering

144 education in the ﬁeld of design by presenting students with challenging design projects drawn from actual NASA interests. In doing so, the program yields two very positive results. Firstly, the students gain a valuable experience that will prepare them for design problems with which they will be faced in their professional careers. Secondly, NASA is able to use the work done by students as an additional resource in meeting its own design objectives. The 1994 projects include: Universal Test Facility; Automated Protein Crystal Growth Facility; Stiffening of the ACES Deployable Space Boom; Launch System Design for Access to Space; LH2 Fuel Tank Design for SSTO Vehicle; and Feed System Design for a Reduced Pressure Tank. CASI Feed Systems; Fuel Tanks; Hybrid Rocket Engines; Launch Vehicles; Mechanical Engineering; Single Stage to Orbit Vehicles; Space Erectable Structures; Space Processing; Structural Design

19950006206 Lockheed Engineering and Sciences Co., Hampton, VA, USA An implementation of a chemical and thermal nonequilibrium flow solver on unstructured meshes and application to blunt bodies Prabhu, Ramadas K.; Aug 1, 1994; In English Contract(s)/Grant(s): NAS1-19000; RTOP 232-01-04-06 Report No.(s): NASA-CR-194967; NAS 1.26:194967; No Copyright; Avail: CASI; A04, Hardcopy This paper presents a nonequilibrium flow solver, implementation of the algorithm on unstructured meshes, and application to hypersonic flow past blunt bodies. Air is modeled as a mixture of five chemical species, namely O2, N2, O, NO, and N, having two temperatures namely translational and vibrational. The solution algorithm is a cell centered, point implicit upwind scheme that employs Roe’s flux difference splitting technique. Implementation of this algorithm on unstructured meshes is described. The computer code is applied to solve Mach 15 flow with and without a Type IV shock interference on a cylindrical body of 2.5mm radius representing a cowl lip. Adaptively generated meshes are employed, and the meshes are refined several times until the solution exhibits detailed flow features and surface pressure and heat flux distributions. Effects of a catalytic wall on surface heat flux distribution are studied. For the Mach 15 Type IV shock interference flow, present results showed a peak heat flux of 544 MW/m2 for a fully catalytic wall and 431 MW/m(exp 2) for a noncatalytic wall. Some of the results are compared with available computational data. Author Air Flow; Blunt Bodies; Computer Programs; Cylindrical Bodies; Grid Generation (Mathematics); Hypersonic Flow; Mach Number; Nonequilibrium Flow; Vibration

19950005762 Vigyan Research Associates, Inc., Hampton, VA, USA A simple analytical aerodynamic model of Langley Winged-Cone Aerospace Plane concept Pamadi, Bandu N.; Oct 1, 1994; In English Contract(s)/Grant(s): NAS1-19341; RTOP 232-01-04-05 Report No.(s): NASA-CR-194987; NAS 1.26:194987; No Copyright; Avail: CASI; A03, Hardcopy A simple three DOF analytical aerodynamic model of the Langley Winged-Coned Aerospace Plane concept is presented in a form suitable for simulation, trajectory optimization, and guidance and control studies. The analytical model is especially suitable for methods based on variational calculus. Analytical expressions are presented for lift, drag, and pitching moment coefficients from subsonic to hypersonic Mach numbers and angles of attack up to +/- 20 deg. This analytical model has break points at Mach numbers of 1.0, 1.4, 4.0, and 6.0. Across these Mach number break points, the lift, drag, and pitching moment coefficients are made continuous but their derivatives are not. There are no break points in angle of attack. The effect of control surface deﬂection is not considered. The present analytical model compares well with the APAS calculations and wind tunnel test data for most angles of attack and Mach numbers. Author (revised) Aerodynamic Coeffıcients; Aerodynamic Drag; Aerospace Planes; Angle of Attack; Hypersonic Aircraft; Lift; Mach Number; Mathematical Models; Pitching Moments

19950005532 Rockwell International Corp., Downey, CA, USA, Hercules, Inc., Wilmington, DE, USA Trade study plan for Reusable Hydrogen Composite Tank System (RHCTS) Greenberg, H. S.; Jul 29, 1994; In English Contract(s)/Grant(s): NCC8-39 Report No.(s): NASA-CR-196871; NAS 1.26:196871; SSD94D0219; RHCTS-TSP-1; No Copyright; Avail: CASI; A03, Hardcopy

145 This TA 1 document describes the trade study plan (with support from TA 2) that will identify the most suitable structural configuration for an SSTO winged vehicle capable of delivering 25,000 lbs to a 220 nm circular orbit at 51.6 degree inclination. The analysis uses information derived in the TA 2 study as identified within the study plan. In view of this, for convenience, the TA 2 study plan is included as an appendix to this document. Derived from text Composite Structures; Launch Vehicle Configurations; Liquid Hydrogen; Propellant Tanks; Single Stage to Orbit Vehicles; Spacecraft Design

19950005521 NASA Lyndon B. Johnson Space Center, Houston, TX, USA Liquid flyback booster pre-phase: A study assessment Peterson, W.; Ankney, W.; Bell, J.; Berning, M.; Bryant, L.; Bufkin, A.; Cain, L.; Caram, J.; Cockrell, B.; Curry, D., et al.; Sep 1, 1994; In English Report No.(s): NASA-TM-104801; S-778; NAS 1.15:104801; No Copyright; Avail: CASI; A08, Hardcopy The concept of a flyback booster has been around since early in the shuttle program. The original two-stage shuttle concepts used a manned flyback booster. These boosters were eliminated from the program for funding and size reasons. The current shuttle uses two Redesigned Solid Rocket Motors (RSRM’s), which are recovered and refurbished after each flight; this is one of the major cost factors of the program. Replacement options have been studied over the past ten years. The conclusion reached by the most recent study is that the liquid flyback booster (LFBB) is the only competitive option from a life-cycle cost perspective. The purpose of this study was to assess the feasibility and practicality of LFBB’s. The study provides an expansion of the recommendations made during the aforementioned study. The primary benefits are the potential for enhanced reusability and a reduction of recurring costs. The potential savings in vehicle turnaround could offset the up-front costs. Development of LFBB’s requires a commitment to the shuttle program for 20 to 30 years. LFBB’s also offer enhanced safety and abort capabilities. Currently, any failure of an RSRM can be considered catastrophic, since there are no intact abort capabilities during the burn of the RSRM’s. The performance goal of the LFBB’s was to lift a fully loaded orbiter under optimal conditions, so as not to be the limiting factor of the performance capability of the shuttle. In addition, a final benefit is the availability of growth paths for applications other than shuttle. Author (revised) Booster Recovery; Reusable Launch Vehicles; Solid Propellant Rocket Engines; Space Shuttle Boosters; Spacecraft Launching

19950005422 Rockwell International Corp., Downey, CA, USA Test Plans. Lightweight Durable TPS: Tasks 1,2,4,5, and 6 Greenberg, H. S.; Tu, Tina; Jul 29, 1994; In English Contract(s)/Grant(s): NCC2-9003 Report No.(s): NASA-CR-196867; NAS 1.26:196867; LDTPS-TP-1; SSD94D0218; No Copyright; Avail: CASI; A10, Hardcopy The objective of this task is to develop the fluted core flexible blankets, also referred to as the Tailorable Advanced Blanket Insulation (TABI), to a technology readiness level (TRL) of 6. This task is one of the six tasks under TA 3, Lightweight Durable TPS study, of the Single Stage to Orbit (SSTO) program. The purpose of this task is to develop a durable and low maintenance flexible TPS blanket material to be implemented on the SSTO vehicle. Author Ceramic Fibers; Composite Materials; Fabrication; Government/Industry Relations; Single Stage to Orbit Vehicles; Thermal Insulation

19950005421 Rockwell International Corp., Downey, CA, USA, Grumman Aerospace Corp., Bethpage, NY, USA Trade study plan for Graphite Composite Primary Structure (GCPS) Greenberg, H. S.; Jul 29, 1994; In English Contract(s)/Grant(s): NCC1-193 Report No.(s): NASA-CR-196868; NAS 1.26:196868; SSD94D0222; GCPS-TSP-1; No Copyright; Avail: CASI; A03, Hardcopy This TA 2 document (with support from TA 1) describes the trade study plan that will identify the most suitable structural conﬁguration for an SSTO winged vehicle capable of delivering 25,000 lbs to a 220 nm circular orbit at 51.6 degree inclination For this most suitable conﬁguration the structural attachment of the wing, and the most suitable GCPS composite materials

146 for intertank, wing, tail and thrust structure are identiﬁed. This trade study analysis uses extensive information derived in the TA 1 trade study plan and is identiﬁed within the study plan. In view of this, for convenience, the TA 1 study plan is included as an appendix to this document. Author Composite Structures; Fuel Tanks; Graphite; Single Stage to Orbit Vehicles; Thrust-Weight Ratio

19950005420 Rockwell International Corp., Downey, CA, USA, Northrop Corp., Los Angeles, CA, USA Test Plan. GCPS Task 4, subtask 4.2 thrust structure development Greenberg, H. S.; Sep 2, 1994; In English Contract(s)/Grant(s): NCC1-193 Report No.(s): NASA-CR-196869; NAS 1.26:196869; SSD94D0211; No Copyright; Avail: CASI; A08, Hardcopy The Single Stage To Orbit (SSTO) vehicle is designed to lift off from a vertical position, go into orbit, return to earth for a horizontal landing, and be reusable for the next mission. (NASA baseline only) In order to meet its performance goals, the SSTO relies on light weight structure and the use of 8 tri-propellant engines. These engines are mounted to the thrust structure. This test plan addresses selection of the material for this structure, and the integrity of the design through testing of elements and a full-scale subcomponent. This test plan supports the development of the design for an advanced composite thrust structure for a Single Stage to Orbit manned, heavy launch vehicle. The thrust structure is designed to transmit very high thrust loads from the engines to the rest of the vehicle (see Figure 1 ). The thrust structure will also be used for primary attachment of the twin vertical tails and possibly act as the aft attach point for the wing. The combination of high loading, high vibration, long service life and high acoustic environments will need to be evaluated by tests. To minimize design risk, a building block approach will be used. We will first screen materials to determine which materials show the most promise for this application. Factors in this screening will be the suitability of these materials for chosen design concepts, particularly concerning specific strength, environmental compatibility and applicability to fabrication processes. Next we will characterize two material systems that will be used in the design; the characterization will allow us to generate preliminary design data that will be used for the analysis. Element testing will be performed to evaluate critical structural locations under load. Final testing on the full scale test article will be performed to verify the design and to demonstrate predictability of the analysis. Additionally, risks associated with fabricating full scale thrust structures will be reduced through testing activities. One of the major concerns that stems from full scale fabrication is the realities of size and the associated complexities of handling, manufacturing, and assembly. The need exists to fabricate, assemble and test_representative joint specimens to achieve_confidence in the design and manufacturing technologies being proposed. Author Composite Structures; Design to Cost; Full Scale Tests; Propulsion System Configurations; Risk; Single Stage to Orbit Vehicles; Thrust-Weight Ratio

19950005410 Rockwell International Corp., Downey, CA, USA Integrated propulsion technology demonstrator. Program plan Aug 15, 1994; In English Contract(s)/Grant(s): NCC1-193 Report No.(s): NASA-CR-196870; NAS 1.26:196870; SSD94D0207; No Copyright; Avail: CASI; A05, Hardcopy NASA and Rockwell have embarked on a cooperative agreement to define, develop, fabricate, and operate an integrated propulsion technology demonstrator (IPTD) for the purpose of validating design, process, and technology improvements of launch vehicle propulsion systems. This program, a result of NRA8-11, Task Area 1 A, is jointly funded by both NASA and Rockwell and is sponsored by the Reusable Launch Vehicle office at NASA Marshall Space flight Center. This program plan provides to the joint NASA/Rockwell integrated propulsion technology demonstrator (IPTD) team a description of the activities within tasks / sub tasks and associated schedules required to successfully achieve program objectives. This document also defines the cost elements and manpower allocations for each sub task for purpose of program control. This plan is updated periodically by developing greater depth of direction for outyear tasks as the program matures. Updating is accomplished by adding revisions to existing pages or attaching page revisions to this plan. In either case, revisions will be identified by appropriate highlighting of the change, or specifying a revision page through the use of footnotes on the bottom right of each change page. Authorization for the change is provided by the principal investigators to maintain control of this program plan document and IPTD program activities. Author Project Planning; Propulsion System Configurations; Requirements; Rocket Engine Design; Schedules; Single Stage to Orbit Vehicles

147 19950005403 NASA Marshall Space Flight Center, Huntsville, AL, USA Damage tolerance of candidate thermoset composites for use on single stage to orbit vehicles Nettles, A. T.; Lance, D.; Hodge, A.; Sep 1, 1994; In English Report No.(s): NASA-TP-3506; M-758; NAS 1.60:3506; No Copyright; Avail: CASI; A03, Hardcopy Four ﬁber/resin systems were compared for resistance to damage and damage tolerance. One toughened epoxy and three toughened bismaleimide (BMI) resins were used, all with IM7 carbon ﬁber reinforcement. A statistical design of experiments technique was used to evaluate the effects of impact energy, specimen thickness, and impactor diameter on the damage area, as computed by C-scans, and residual compression-after-impact (CAI) strength. Results showed that two of the BMI systems sustained relatively large damage zones yet had an excellent retention of CAI strength. Author (revised) Bismaleimide; Carbon Fiber Reinforced Plastics; Epoxy Resins; Impact Damage; Impact Resistance; Matrix Materials; Polyimide Resins; Single Stage to Orbit Vehicles; Spacecraft Construction Materials

19950005396 Rockwell International Corp., Downey, CA, USA, Hercules, Inc., Wilmington, DE, USA Selection process for trade study: Reusable Hydrogen Composite Tank System (RHCTS) Greenberg, H. S.; Sep 2, 1994; In English Contract(s)/Grant(s): NCC8-39 Report No.(s): NASA-CR-196874; NAS 1.26:196874; SSD94D0209; No Copyright; Avail: CASI; A03, Hardcopy This document describes the selection process that will be used to identify the most suitable structural configuration option for an SSTO winged vehicle capable of delivering 25,000 lbs to a 220 nm circular orbit at 51.6 degree inclination. The most suitable RHCTS is within this configuration and will be the prototype design for subsequent design and analysis and the basis for the design and fabrication of a scale test article to be subjected to life cycle testing. The selection process for this TA 1 trade study is the same as that for the TA 2 trade study. As the trade study progresses additional insight may result in modifications to the selection criteria within in this process. Such modifications will result in an update of this document as appropriate. Author Circular Orbits; Design Analysis; Horizontal Spacecraft Landing; Hydrogen Fuels; Prototypes; Single Stage to Orbit Vehicles

19950005390 Rockwell International Corp., Downey, CA, USA, Grumman Aerospace Corp., Bethpage, NY, USA Selection process for trade study: Graphite Composite Primary Structure (GCPS) Greenberg, H. S.; Sep 2, 1994; In English Contract(s)/Grant(s): NCC1-193 Report No.(s): NASA-CR-196866; NAS 1.26:196866; SSD94D0210; No Copyright; Avail: CASI; A03, Hardcopy This TA 2 document describes the selection process that will be used to identify the most suitable structural configuration for an SSTO winged vehicle capable of delivering 25,000 lbs to a 220 nm circular orbit at 51.6 degree inclination. The most suitable unpressurized graphite composite structures and material selections is within this configuration and will be the prototype design for subsequent design and analysis and the basis for the design and fabrication of payload bay, wing, and thrust structure full scale test articles representing segments of the prototype structures. The selection process for this TA 2 trade study is the same as that for the TA 1 trade study. As the trade study progresses additional insight may result in modifications to the selection criteria within this process. Such modifications will result in an update of this document as appropriate. Author Composite Structures; Fabrication; Full Scale Tests; Graphite; Propulsion System Configurations; Single Stage to Orbit Vehicles; Structural Design; Thrust-Weight Ratio

19950005389 Rockwell International Corp., Downey, CA, USA, Hercules, Inc., Wilmington, DE, USA Test plan. Task 5, subtask 5.2: Early on-orbit TPSdebris impact tests Greenberg, H. S.; Jul 29, 1994; In English Contract(s)/Grant(s): NCC8-39 Report No.(s): NASA-CR-196873; NAS 1.26:196873; SSD94D0221; RHCTS-TP-1; No Copyright; Avail: CASI; A03, Hardcopy The limitation of damage to, and survival of, the cryogenic tankage during the on-orbit stay despite potential impact of orbital debris, may be a signiﬁcant discriminator in the RHCTS trade studies described in the TA-1 trade study plan (ref.

148 RHCTS-TSP-1) dated July 29, 1994. The objective of this early phase of an overall debris impact test program is to provide the data to support assessment of the relative suitability of integral and non integral tanks. Author Cryogenic Rocket Propellants; Damage; Fuel Tanks; Hydrogen Fuels; Impact Tests; Single Stage to Orbit Vehicles; Survival

19950005364 Rockwell International Corp., Downey, CA, USA, Northrop Corp., Los Angeles, CA, USA, Hercules Aerospace Co., Wilmington, DE, USA Requirements report for SSTO vertical take-off/horizontal landing vehicle Greenberg, H. S.; Jul 29, 1994; In English Contract(s)/Grant(s): NCC1-193; NCC2-9003; NCC8-39 Report No.(s): NASA-CR-196875; NAS 1.26:196875; SSD94D0217; SSTO-REQ-1; No Copyright; Avail: CASI; A04, Hardcopy This document describes the detailed design requirements and design criteria to support Structures/TPS Technology development for SSTO winged vehicle conﬁgurations that use vertical take-off and horizontal landing and deliver 25,000 lb payloads to a 220 nm circular orbit at an inclination of 51.6 degrees or 40,000 lb payloads to a 150 nm circular orbit at a 28.5 degree of inclination. This document will be updated on a timely basis as informatIon becomes available throughout the project. Author Design Analysis; Horizontal Spacecraft Landing; Reusable Launch Vehicles; Single Stage to Orbit Vehicles

19950005354 Rockwell International Corp., Downey, CA, USA, Northrop Corp., Los Angeles, CA, USA Test plan. GCPS task 7, subtask 7.1: IHM development Greenberg, H. S.; Sep 2, 1994; In English Contract(s)/Grant(s): NCC1-193 Report No.(s): NASA-CR-106876; NAS 1.26:196876; SSD940212; No Copyright; Avail: CASI; A03, Hardcopy The overall objective of Task 7 is to identify cost-effective life cycle integrated health management (IHM) approaches for a reusable launch vehicle’s primary structure. Acceptable IHM approaches must: eliminate and accommodate faults through robust designs, identify optimum inspection/maintenance periods, automate ground and on-board test and check-out, and accommodate and detect structural faults by providing wide and localized area sensor and test coverage as required. These requirements are elements of our targeted primary structure low cost operations approach using airline-like maintenance by exception philosophies. This development plan will follow an evolutionary path paving the way to the ultimate development of flight-quality production, operations, and vehicle systems. This effort will be focused on maturing the recommended sensor technologies required for localized and wide area health monitoring to a technology readiness level (TRL) of 6 and to establish flight ready system design requirements. The following is a brief list of IHM program objectives: design out faults by analyzing material properties, structural geometry, and load and environment variables and identify failure modes and damage tolerance requirements; design in system robustness while meeting performance objectives (weight limitations) of the reusable launch vehicle primary structure; establish structural integrity margins to preclude the need for test and checkout and predict optimum inspection/maintenance periods through life prediction analysis; identify optimum fault protection system concept definitions combining system robustness and integrity margins established above with cost effective health monitoring technologies; and use coupons, panels, and integrated full scale primary structure test articles to identify, evaluate, and characterize the preferred NDE/NDI/IHM sensor technologies that will be a part of the fault protection system. Derived from text Fault Tolerance; Nondestructive Tests; Single Stage to Orbit Vehicles; Spacecraft Maintenance; Spacecraft Reliability; Structural Reliability

19950005173 McDonnell-Douglas Aerospace, Huntington Beach, CA, USA Slush hydrogen technology program Cady, Edwin C.; Sep 9, 1994; In English Contract(s)/Grant(s): NAS3-25420; RTOP 763-22-2T Report No.(s): AD-A284706; NASA-CR-195353; E-9003; NAS 1.26:195353; MDC-94-H00684; No Copyright; Avail: CASI; A04, Hardcopy A slush hydrogen (SH2) technology facility (STF) was designed, fabricated, and assembled by a contractor team of McDonnell Douglas Aerospace (MDA), Martin Marietta Aerospace Group (MMAG), and Air Products and Chemicals, Inc.

149 (APCI). The STF consists of a slush generator which uses the freeze-thaw production process, a vacuum subsystem, a test tank which simulates the NASP vehicle, a triple point hydrogen receiver tank, a transfer subsystem, a sample bottle, a pressurization system, and a complete instrumentation and control subsystem. The STF was fabricated, checked-out, and made ready for testing under this contract. The actual SH2 testing was performed under the NASP consortium following NASP teaming. Pre-STF testing veriﬁed SH2 production methods, validated special SH2 instrumentation, and performed limited SH2 pressurization and expulsion tests which demonstrated the need for gaseous helium pre-pressurized of SH2 to control pressure collapse. The STF represents cutting-edge technology development by an effective Government-Industry team under very tight cost and schedule constraints. Author Cryogenics; Liquid Hydrogen; National Aerospace Plane Program; Research Facilities; Slush Hydrogen; Test Facilities

19950004141 NASA, Washington, DC, USA Aero-Space Plane: Flexible access to space Aug 1, 1991; In English Report No.(s): NASA-TM-109904; NONP-NASA-VT-94-23146; ASR-257; No Copyright; Avail: CASI; B01, Videotape- Beta; V01, Videotape-VHS The most recently designed X-30 (National Aerospace Plane) is described. The video feature also chronicles the development of the X-plane series, beginning with the X-1. CASI Aerospace Planes; National Aerospace Plane Program; X-31 Aircraft

19950002971 Rockwell International Corp., Downey, CA, USA Propulsion module design review. Milestone 8: Integrated Propulsion Technology Demonstrator Rogers, James P.; Vesperman, Curtis L.; Apr 17, 1995; In English Contract(s)/Grant(s): NCC8-47 Report No.(s): NASA-CR-197998; NAS 1.26:197998; SSD95D0173; No Copyright; Avail: CASI; A10, Hardcopy The propulsion module for the Integrated Propulsion Technology Demonstrator (IPTD) consists of main engine(s), and fluid, electrical, propulsion checkout and control systems (PCCS), and associated instrumentation representative of an X-33 main propulsion system (MPS). These systems are integrated within a structural support fixture which includes mounting provisions for the MPS feed lines and ancillary equipment. The weight of the main engines and associated thrust loads are reacted by a separated thrust structure mounted directly to the large structural beams (i.e., Norman Beams) of the Advanced Engine Test Facility (AETF). This document provides a summary of the design requirements, and associated design features of the major elements of the propulsion module. Significant design detail is provided for the hardware that will be procured, fabricated, assembled and delivered to MSFC in October 1995. Designs of hardware items due for fabrication at a later date (such as the thrust structure and tank modules) are presented only to the degree necessary to indicate critical interface requirements. This report also documents the findings of an independent peer of the design to ensure that major design oversights are caught early in this design and fabrication process. This peer review board consisted of senior propulsion, control, and software engineers at Rockwell who had avionics/ground control and data handling system or cryogenic design experience on the Saturn launch vehicle and/or the Shuttle. Derived from text Design Analysis; Liquid Propellant Rocket Engines; Propulsion System Configurations; Propulsion System Performance; Recoverable Launch Vehicles; Rocket Engine Design; Test Stands

19950002820 Rockwell International Corp., Downey, CA, USA Reusable launch vehicle reliability, maintainability and operability assessment Weber, T. F., Jr.; Mar 15, 1995; In English Contract(s)/Grant(s): NAS8-39210 Report No.(s): NASA-CR-197695; NAS 1.26:197695; SSD95D0123; No Copyright; Avail: CASI; A03, Hardcopy Rockwell’s Space Systems Division was funded to conduct thorough reliability, maintainability and operability (RM&O) assessments for 3 NASA conceptually-conceived SSTO vehicles: a Winged Body (WB001), a Vertical Lander (VL001) and, at some future time when input data becomes available, a lifting body. Rockwell was instructed to use their existing ‘generic’ RM&O analysis models (MAtrix, SIMtrix, and STARSIM), and apply them to the RLV vehicles of interest. The RM&O assessments were to encompass, as a minimum, the following parameters: MTBM (mean time between maintenance), MTBR

150 (mean time between removal), MTBF (mean time between failure), MTBCF (mean time before critical failure), MTTR (mean time to repair), probability of mission success, turnaround time, crew size per repair, manhours per repair, manhours per mission, ﬂight rate capability, availability, downtime per mission, manhours per processing ﬂow (scheduled and unscheduled), and facility utilization. Additionally, the MAtrix model was to be exercised for the existing Space Shuttle and its infrastructure in order to judge the ‘reasonableness’ of the SSTO RM&O estimates, and to provide a basis for comparing these SSTO estimates to ‘observed’ Shuttle experiences. Fully operable Reliability and Maintainability Models for WB001, VL001 and the Space Shuttle were electronically transmitted to MSFC on December 13, 1994. Derived from text Maintainability; Reliability; Single Stage to Orbit Vehicles; Spacecraft Maintenance

19950002293 Rockwell International Corp., Downey, CA, USA Lightweight durable TPS: Milestone 4 Fleming, M. M.; Greenberg, H. S.; Nov 15, 1994; In English Contract(s)/Grant(s): NCC2-9003 Report No.(s): NASA-CR-197103; NAS 1.26:197103; SSD94D0336; No Copyright; Avail: CASI; A03, Hardcopy This document, covering the TA3-Lightweight Durable Thermal Protection System, concerns the blanket assembly, the evaluation plan of the hook and loop (Velcro-type) fasteners, and a report relating the adverse environmental effects of the SSTO (Single Stage To Orbit) vehicle and their relationship to TPS components and locations. These adverse environmental effects concern only those part of the natural environment and not induced environmental effects, such as, vibro-acoustics, aerothermal heating, or aerodynamic loads. CASI Blankets; Chemical Analysis; Durability; Environment Effects; Fasteners; Materials Tests; Salt Spray Tests; Thermal Analysis; Thermal Protection

19940032138 NASA Lewis Research Center, Cleveland, OH, USA Hot dynamic test rig for measuring hypersonic engine seal flow and durability Miller, Jeffrey H.; Steinetz, Bruce M.; Sirocky, Paul J.; Kren, Lawrence A.; Feb 1, 1994; In English Report No.(s): NASA-TM-106507; E-8031-1; NAS 1.15:10657; No Copyright; Avail: CASI; A03, Hardcopy A test fixture for measuring the dynamic performance of candidate high-temperature engine seal concepts was developed. The test fixture was developed to evaluate seal concepts under development for advanced hypersonic engines, such as those being considered for the National Aerospace Plane (NASP). The fixture can measure dynamic seal leakage performance from room temperature up to 840 C and air pressure differentials of to 0.7 MPa. Performance of the seals can be measured while sealing against flat or engine-simulated distorted walls. In the fixture, two seals are preloaded against the sides of a 0.3 m long saber that slides transverse to the axis of the seals, simulating the scrubbing motion anticipated in these engines. The capabilities of this text fixture along with preliminary data showing the dependence of seal leakage performance on high temperature cycling are covered. Author (revised) Aerospace Planes; Dynamic Tests; Engine Design; Engine Tests; High Temperature; Hypersonic Flow; Hypersonics; Leakage; Propulsion System Performance; Seals (Stoppers); Temperature Effects; Thermal Cycling Tests

19940029244 NASA Lewis Research Center, Cleveland, OH, USA National aerospace plane Jul 1, 1990; In English Report No.(s): LERC-4002; NASA-TM-109843; NONP-NASA-VT-94-13533; No Copyright; Avail: CASI; B01, Videotape- Beta; V01, Videotape-VHS This video concentrates on materials being developed and tested at LeRC for possible use in NASP. CASI Aerospace Planes; Aircraft Construction Materials; National Aerospace Plane Program; Spacecraft Construction Materials

19940029064 NASA John C. Stennis Space Center, Bay Saint Louis, MS, USA High Heat Flux Facility JAN 1, 1993; In English Report No.(s): NASA-TM-109834; NONP-NASA-VT-94-12962; No Copyright; Avail: CASI; B01, Videotape-Beta; V01, Videotape-VHS

151 This video gives an overview of the High Heat Flux Facility being built at Stennis Space Center in conjunction with Wright-Patterson Air Force Base. This facility will simulate ﬂight heat conditions and will be used to test engine and materials for the National Aerospace Plane. CASI Flight Conditions; Heat Flux; National Aerospace Plane Program; Test Facilities

19940029061 NASA John C. Stennis Space Center, Bay Saint Louis, MS, USA John C. Stennis Space Center overview May 1, 1994; In English Report No.(s): NASA-TM-109816; NONP-NASA-VT-94-12944; No Copyright; Avail: CASI; B01, Videotape-Beta; V01, Videotape-VHS An overview of research being conducted at the John C. Stennis Space Center is given. The Space Center is not only a NASA Space Flight Center, but also houses facilities for 22 other governmental agencies. The programs described are Stennis’ High Heat Flux Facility, the Component Test Facility (used to test propulsion rockets and for the development of the National Aerospace Plane), oceanographic and remote sensing research, and contributions to the development of Space Station Freedom. CASI National Aerospace Plane Program; Research Facilities; Space Station Freedom; Test Facilities

19940028594 Rockwell International Corp., Huntsville, AL, USA Advanced transportation system study: Manned launch vehicle concepts for two way transportation system payloads to LEO Duffy, James B.; JAN 1, 1993; In English Contract(s)/Grant(s): NAS8-39207 Report No.(s): NASA-CR-193951; NAS 1.26:193951; SSD93M0025-1; No Copyright; Avail: CASI; A25, Hardcopy The purpose of the Advanced Transportation System Study (ATSS) task area 1 study effort is to examine manned launch vehicle booster concepts and two-way cargo transfer and return vehicle concepts to determine which of the many proposed concepts best meets NASA’s needs for two-way transportation to low earth orbit. The study identified specific configurations of the normally unmanned, expendable launch vehicles (such as the National Launch System family) necessary to fly manned payloads. These launch vehicle configurations were then analyzed to determine the integrated booster/spacecraft performance, operations, reliability, and cost characteristics for the payload delivery and return mission. Design impacts to the expendable launch vehicles which would be required to perform the manned payload delivery mission were also identified. These impacts included the implications of applying NASA’s man-rating requirements, as well as any mission or payload unique impacts. The booster concepts evaluated included the National Launch System (NLS) family of expendable vehicles and several variations of the NLS reference configurations to deliver larger manned payload concepts (such as the crew logistics vehicle (CLV) proposed by NASA JSC). Advanced, clean sheet concepts such as an F-1A engine derived liquid rocket booster (LRB), the single stage to orbit rocket, and a NASP-derived aerospace plane were also included in the study effort. Existing expendable launch vehicles such as the Titan 4, Ariane 5, Energia, and Proton were also examined. Although several manned payload concepts were considered in the analyses, the reference manned payload was the NASA Langley Research Center’s HL-20 version of the personnel launch system (PLS). A scaled up version of the PLS for combined crew/cargo delivery capability, the HL-42 configuration, was also included in the analyses of cargo transfer and return vehicle (CTRV) booster concepts. In addition to strictly manned payloads, two-way cargo transportation systems (CTRV’s) were also examined. The study provided detailed design and analysis of the performance, reliability, and operations of these concepts. The study analyzed these concepts as unique systems and also analyzed several combined CTRV/booster configurations as integrated launch systems (such as for launch abort analyses). Included in the set of CTRV concepts analyzed were the medium CTRV, the integral CTRV (in both a pressurized and unpressurized configuration), the winged CTRV, and an attached cargo carrier for the PLS system known as the PLS caboose. Author (revised) Launch Vehicle Configurations; Launch Vehicles; Manned Space Flight; Manned Spacecraft; Space Logistics; Space Transportation System; Spacecraft Launching

152 19940027298 NASA Lewis Research Center, Cleveland, OH, USA High temperature NASP engine seal development JAN 1, 1992; In English Report No.(s): LERC-92-174; NASA-TM-109750; NONP-NASA-VT-94-9950; No Copyright; Avail: CASI; B01, Videotape- Beta; V01, Videotape-VHS This video details research being conducted at the Lewis Research Center on high temperature engine seal design for the National Aerospace Plane. To maximize the speed, the jets on the NASP extract oxygen from the air rather than carry large liquid fuel tanks; this creates temperatures within the jet of over 5000 F. To prevent these potentially explosive gases from escaping, researchers are developing new technologies for use in the engine seals. Two examples explained are the ceramic wafer seal and the braided ceramic rope seal. Computer simulations and laboratory footage are used to illustrate the workings of these seals. Beneﬁts for other aerospace and industrial applications, as well as for the space shuttle, are explored. CASI Aerospace Planes; Ceramics; Engine Parts; High Temperature; National Aerospace Plane Program; Refractory Materials; Seals (Stoppers)

19940023286 Maryland Univ., College Park, MD, USA A parametric sensitivity study for single-stage-to-orbit hypersonic vehicles using trajectory optimization Lovell, T. Alan; Schmidt, D. K.; Mar 21, 1994; In English Contract(s)/Grant(s): NAG1-1540 Report No.(s): NASA-CR-195703; NAS 1.26:195703; No Copyright; Avail: CASI; A03, Hardcopy The class of hypersonic vehicle configurations with single stage-to-orbit (SSTO) capability reflect highly integrated airframe and propulsion systems. These designs are also known to exhibit a large degree of interaction between the airframe and engine dynamics. Consequently, even simplified hypersonic models are characterized by tightly coupled nonlinear equations of motion. In addition, hypersonic SSTO vehicles present a major system design challenge; the vehicle’s overall mission performance is a function of its subsystem efficiencies including structural, aerodynamic, propulsive, and operational. Further, all subsystem efficiencies are interrelated, hence, independent optimization of the subsystems is not likely to lead to an optimum design. Thus, it is desired to know the effect of various subsystem efficiencies on overall mission performance. For the purposes of this analysis, mission performance will be measured in terms of the payload weight inserted into orbit. In this report, a trajectory optimization problem is formulated for a generic hypersonic lifting body for a specified orbit-injection mission. A solution method is outlined, and results are detailed for the generic vehicle, referred to as the baseline model. After evaluating the performance of the baseline model, a sensitivity study is presented to determine the effect of various subsystem efficiencies on mission performance. This consists of performing a parametric analysis of the basic design parameters, generating a matrix of configurations, and determining the mission performance of each configuration. Also, the performance loss due to constraining the total head load experienced by the vehicle is evaluated. The key results from this analysis include the formulation of the sizing problem for this vehicle class using trajectory optimization, characteristics of the optimal trajectories, and the subsystem design sensitivities. Author (revised) Airframes; Equations of Motion; Hypersonic Vehicles; Nonlinear Equations; Parameter Identification; Propulsion System Configurations; Propulsion System Performance; Single Stage to Orbit Vehicles; Trajectory Optimization; Trajectory Planning

19940022648 NASA, Washington, DC, USA Access to space study Jan 1, 1994; In English Report No.(s): NASA-TM-109693; NAS 1.15:109693; No Copyright; Avail: CASI; A05, Hardcopy This report summarizes the results of a comprehensive NASA in-house study to identify and assess alternate approaches to access to space through the year 2030, and to select and recommend a preferred cause of action. The goals of the study were to identify the best vehicles and transportation architectures to make major reductions in the cost of space transportation (at least 50%), while at the same time increasing safety for ﬂight crews by at least an order of magnitude. In addition, vehicle reliability was to exceed 0.98 percent, and, as important, the robustness, pad time, turnaround time, and other aspects of operability were to be vastly improved. This study examined three major optional architectures: (1) retain and upgrade the Space Shuttle and expendable launch vehicles; (2) develop new expendable vehicles using conventional technologies and transition from current vehicles beginning in 2005; and (3) develop new reusable vehicles using advanced technology, and transition from current vehicles beginning in 2008. The launch-needs, mission model utilized for for the study was based upon

153 today’s projection of civil, defense, and commercial mission payload requirements. Derived from text Expendable Stages (Spacecraft); Launch Vehicles; Mission Planning; Reliability Analysis; Single Stage to Orbit Vehicles; Space Commercialization; Space Missions; Space Shuttles; Space Transportation; Spacecraft Launching

19940020603 Ohio State Univ., Columbus, OH, USA A conceptual design of an unmanned test vehicle using an airbreathing propulsion system JAN 1, 1992; In English Contract(s)/Grant(s): NASW-4435 Report No.(s): NASA-CR-195550; NAS 1.26:195550; No Copyright; Avail: CASI; A04, Hardcopy According to Aviation Week and Space Technology (Nov. 16, 1992), without a redefined approach to the problem of achieving single stage-to-orbit flight, the X-30 program is virtually assured of cancellation. One of the significant design goals of the X-30 program is to achieve single stage to low-earth orbit using airbreathing propulsion systems. In an attempt to avoid cancellation, the NASP Program has decided to design a test vehicle to achieve these goals. This report recommends a conceptual design of an unmanned test vehicle using an airbreathing propulsion system. Author (revised) Aerospace Planes; Air Breathing Engines; Aircraft Design; National Aerospace Plane Program; Pilotless Aircraft; Propulsion System Performance; Research Aircraft; Single Stage to Orbit Vehicles

19940020118 Ohio State Univ., Columbus, OH, USA Configuration development study of the OSU 1 hypersonic research vehicle Stein, Matthew D.; Frankhauser, Chris; Zee, Warner; Kosanchick, Melvin, III; Nelson, Nick; Hunt, William; May 31, 1993; In English Contract(s)/Grant(s): NASW-4435 Report No.(s): NASA-CR-195522; NAS 1.26:195522; No Copyright; Avail: CASI; A08, Hardcopy In an effort to insure the future development of hypersonic cruise aircraft, the possible vehicle configurations were examined to develop a single-stage-to-orbit hypersonic research vehicle (HRV). Based on the needs of hypersonic research and development, the mission goals and requirements are determined. A body type is chosen. Three modes of propulsion and two liquid rocket fuels are compared, followed by the optimization of the body configuration through aerodynamic, weight, and trajectory studies. A cost analysis is included. Author (revised) Aircraft Configurations; Hydrocarbon Fuels; Hydrogen Fuels; Hypersonic Vehicles; Research Vehicles; Single Stage to Orbit Vehicles; Thermal Protection

19940017090 NASA Lewis Research Center, Cleveland, OH, USA Thrust augmentation options for the Beta 2 two-stage-to-orbit vehicle Snyder, Christopher A.; Dec 1, 1993; In English; AIAA Aircraft Design, Systems and Operations Meeting, 11-13 Aug. 1993, Monterey, CA, USA Contract(s)/Grant(s): RTOP 505-70-00 Report No.(s): NASA-TM-106448; E-8301; NAS 1.15:106448; AIAA PAPER 93-4014; No Copyright; Avail: CASI; A03, Hardcopy NASA LeRC is continuing to study propulsion concepts for a horizontal takeoff and landing, fully reusable, two-stage-to-orbit vehicle. This will be capable of launching and returning a 10,000 pound payload to a 100 nautical mile polar orbit using low-risk technology. The vehicle, Beta 2, is a derivative of the USAF/Boeing Beta vehicle which was designed to deliver a 50,000 pound payload to a similar orbit. Beta 2 stages at Mach 6.5 and about 100,000 ft altitude. The propulsion system for the booster is an over/under turbine bypass engine/ramjet conﬁguration. In this paper, several options for thrust augmentation were studied in order to improve the performance of this engine where there was a critical need. Options studies were turbine engine overspeed in the transonic region, water injection at a various turbine engine locations also during the transonic region, and water injection at the turbine engine face during high speed operation. The methodology, constraints, propulsion performance, and mission study results are presented. Author (revised) Aerospace Planes; Air Breathing Boosters; Gas Turbine Engines; Hypersonic Speed; Propulsion System Conﬁgurations; Propulsive Effıciency; Thrust Augmentation; Water Injection

154 19940012847 NASA Lyndon B. Johnson Space Center, Houston, TX, USA Human Transportation System (HTS) study, volume 2 Lance, N.; Geyer, M. S.; Gaunce, M. T.; Oct 1, 1993; In English Contract(s)/Grant(s): RTOP 906-11-01-01 Report No.(s): NASA-TM-104779-VOL-2; S-739; NAS 1.15:104779-VOL-2; No Copyright; Avail: CASI; A24, Hardcopy This report summarizes work completed under the Human Transportation System Study. This study was conducted by the New Initiatives Office at JSC with the technical support of Boeing, General Dynamics, Lockheed, McDonnell-Douglas, Martin Marietta, and Rockwell. The study was designed to generate information on determining the appropriate path to follow for new system development to meet the Nation’s space transportation needs. The study evaluates 18 transportation architecture options using a parametric set of mission requirements. These options include use of current systems (e.g., Shuttle, Titan, etc. ) as well as proposed systems (e.g., PLS, Single-Stage-to-Orbit, etc.) to assess the impact of various considerations, such as the cost of alternate access, or the benefit of separating people and cargo. The architecture options are compared to each other with six measurable evaluation criteria or attributes. They are: funding profile, human safety, probability of mission success, architecture cost risk, launch schedule confidence, and environmental impact. Values for these attributes are presented for the architecture options, with pertinent conclusions and recommendations. Derived from text Human Factors Engineering; Manned Space Flight; Manned Spacecraft; Mission Planning; Space Transportation; Systems Engineering

19940012582 NASA Langley Research Center, Hampton, VA, USA Aerothermoelastic analysis of a NASP demonstrator model Heeg, Jennifer; Zeiler, Thomas A.; Pototzky, Anthony S.; Spain, Charles V.; Engelund, Walter C.; Oct 1, 1993; In English Contract(s)/Grant(s): RTOP 505-63-50-15 Report No.(s): NASA-TM-109007; NAS 1.15:109007; No Copyright; Avail: CASI; A03, Hardcopy The proposed National Aerospace Plane (NASP) is designed to travel at speeds up to Mach 25. Because aerodynamic heating during high-speed flight through the atmosphere could destiffen a structure, significant couplings between the elastic and rigid body modes could result in lower flutter speeds and more pronounced aeroelastic response characteristics. These speeds will also generate thermal loads on the structure. The purpose of this research is to develop methodologies applicable to the NASP and to apply them to a representative model to determine its aerothermoelastic characteristics when subjected to these thermal loads. This paper describes an aerothermoelastic analysis of the generic hypersonic vehicle configuration. The steps involved in this analysis were: generating vehicle surface temperatures at the appropriate flight conditions; applying these temperatures to the vehicle’s structure to predict changes in the stiffness resulting from material property degradation; predicting the vibration characteristics of the heated structure at the various temperature conditions; performing aerodynamic analyses; and conducting flutter analysis of the heated vehicle. Results of these analyses and conclusions representative of a NASP vehicle are provided in this paper. Author Aerodynamic Heating; Aerodynamic Stability; Aerospace Planes; Aerothermoelasticity; Dynamic Response; Hypersonic Vehicles; Mach Number; Supersonic Flutter; Temperature Distribution

19940010851 NASA, Washington, DC, USA National Aero-Space Plane resource reel Aug 1, 1991; In English Report No.(s): NASA-TM-109451; NONP-NASA-VT-93-190248; No Copyright; Avail: CASI; B02, Videotape-Beta; V02, Videotape-VHS This document presents a series of takes and sequences of model photography of the 1991 NASP design. CASI Aircraft Models; National Aerospace Plane Program; Photography

19940010850 NASA, Washington, DC, USA National Aero-Space Plane Jul 1, 1990; In English Report No.(s): ASR-253; NASA-TM-109450; NONP-NASA-VT-93-190247; No Copyright; Avail: CASI; B01, Videotape- Beta; V01, Videotape-VHS

155 This document presents updated model photography of ‘old’ NASP design. CASI Aircraft Models; National Aerospace Plane Program; Photography

19940009631 Iowa State Univ. of Science and Technology, Ames, IA, USA Trajectory optimization for the national aerospace plane Lu, Ping; Jul 1, 1993; In English Contract(s)/Grant(s): NAG1-1255 Report No.(s): NASA-CR-194146; NAS 1.26:194146; No Copyright; Avail: CASI; A03, Hardcopy During the past six months the research objectives outlined in the last semi-annual report were accomplished. Speciﬁcally, these are: three-dimensional (3-D) fuel-optimal ascent trajectory of the aerospace plane and the effects of thrust vectoring control (TVC) on the fuel consumption and trajectory shaping were investigated; the maximum abort landing area (footprint) was studied; preliminary assessment of simultaneous design of the ascent trajectory and the vehicle conﬁguration for the aerospace plane was also conducted. The work accomplished in the reporting period is summarized. Author (revised) Aerospace Planes; Ascent Trajectories; Fuel Consumption; National Aerospace Plane Program; Trajectory Optimization

19940009145 North Carolina State Univ., Raleigh, NC, USA Multidisciplinary design techniques applied to conceptual aerospace vehicle design Olds, John Robert; Walberg, Gerald D.; JAN 1, 1993; In English Contract(s)/Grant(s): NCC1-168 Report No.(s): NASA-CR-194409; NAS 1.26:194409; No Copyright; Avail: CASI; A09, Hardcopy Multidisciplinary design optimization (MDO) is an emerging discipline within aerospace engineering. Its goal is to bring structure and efficiency to the complex design process associated with advanced aerospace launch vehicles. Aerospace vehicles generally require input from a variety of traditional aerospace disciplines - aerodynamics, structures, performance, etc. As such, traditional optimization methods cannot always be applied. Several multidisciplinary techniques and methods were proposed as potentially applicable to this class of design problem. Among the candidate options are calculus-based (or gradient-based) optimization schemes and parametric schemes based on design of experiments theory. A brief overview of several applicable multidisciplinary design optimization methods is included. Methods from the calculus-based class and the parametric class are reviewed, but the research application reported focuses on methods from the parametric class. A vehicle of current interest was chosen as a test application for this research. The rocket-based combined-cycle (RBCC) single-stage-to-orbit (SSTO) launch vehicle combines elements of rocket and airbreathing propulsion in an attempt to produce an attractive option for launching medium sized payloads into low earth orbit. The RBCC SSTO presents a particularly difficult problem for traditional one-variable-at-a-time optimization methods because of the lack of an adequate experience base and the highly coupled nature of the design variables. MDO, however, with it’s structured approach to design, is well suited to this problem. The result of the application of Taguchi methods, central composite designs, and response surface methods to the design optimization of the RBCC SSTO are presented. Attention is given to the aspect of Taguchi methods that attempts to locate a ‘robust’ design - that is, a design that is least sensitive to uncontrollable inﬂuences on the design. Near-optimum minimum dry weight solutions are determined for the vehicle. A summary and evaluation of the various parametric MDO methods employed in the research are included. Recommendations for additional research are provided. Author (revised) Aerospace Vehicles; Design Analysis; Optimization; Single Stage to Orbit Vehicles

19940008783 NASA Hugh L. Dryden Flight Research Facility, Edwards, CA, USA Perspective on the National Aero-Space Plane Program instrumentation development Bogue, Rodney K.; Erbland, Peter; May 1, 1993; In English; Aero Space Planes Conference, 1-3 Dec. 1992, Orlando, FL, USA Contract(s)/Grant(s): RTOP 763-21-41 Report No.(s): NASA-TM-4505; H-1916; NAS 1.15:4505; Copyright; Avail: CASI; A03, Hardcopy A review of the requirement for, and development of, advanced measurement technology for the National Aerospace Plane program is presented. The objective is to discuss the technical need and the program commitment required to ensure that adequate and timely measurement capabilities are provided for ground and ﬂight testing in the NASP program. The scope of the measurement problem is presented, the measurement process is described, how instrumentation technology

156 development has been affected by NASP program evolution is examined, the national effort to deﬁne measurement requirements and assess the adequacy of current technology to support the NASP program is discussed, and the measurement requirements are summarized. The unique features of the NASP program that complicate the understanding of requirements and the development of viable solutions are illustrated. Author (revised) Aerospace Planes; Measuring Instruments; National Aerospace Plane Program; Spacecraft Instruments

19940008377 NASA Langley Research Center, Hampton, VA, USA A method of predicting quasi-steady aerodynamics for flutter analysis of high speed vehicles using steady CFD calculations Scott, Robert C.; Pototzky, Anthony S.; Jul 1, 1993; In English; AIAA Structures, Structural Dynamics, and Materials Conference,, 19-21 Apr. 1993, LaJolla, CA, USA Contract(s)/Grant(s): RTOP 505-63-50-15 Report No.(s): NASA-TM-109009; NAS 1.15:109009; No Copyright; Avail: CASI; A02, Hardcopy High speed linear aerodynamic theories like piston theory and Newtonian impact theory are relatively inexpensive to use for flutter analysis. These theories have limited areas of applicability depending on the configuration and the flow conditions. In addition, these theories lack the ability to capture viscous, shock, and real gas effects. CFD methods can model all of these effects accurately, but the unsteady calculations required for flutter are expensive and often impractical. This paper describes a method for using steady CFD calculations to approximate the generalized aerodynamic forces for a flutter analysis. Example two-and three-dimensional aerodynamic force calculations are provided. In addition, a flutter analysis of a NASP-type wing will be discussed. Author (revised) Aerodynamic Characteristics; Aerodynamic Forces; Aerospace Planes; Computational Fluid Dynamics; Dynamic Structural Analysis; Flutter Analysis; Prediction Analysis Techniques; Unsteady Aerodynamics; Wings

19940006270 California Univ., Los Angeles, CA, USA Lee waves: Benign and malignant Wurtele, M. G.; Datta, A.; Sharman, R. D.; Jun 1, 1993; In English Contract(s)/Grant(s): NCC2-0374; RTOP 505-68-50 Report No.(s): NASA-CR-186024; H-1890; NAS 1.26:186024; No Copyright; Avail: CASI; A03, Hardcopy The flow of an incompressible fluid over an obstacle will produce an oscillation in which buoyancy is the restoring force, called a gravity wave. For disturbances of this scale, the atmosphere may be treated as dynamically incompressible, even though there exists a mean static upward density gradient. Even in the linear approximation - i.e., for small disturbances - this model explains a great many of the flow phenomena observed in the lee of mountains. However, nonlinearities do arise importantly, in three ways: (1) through amplification due to the decrease of mean density with height; (2) through the large (scaled) size of the obstacle, such as a mountain range; and (3) from dynamically singular levels in the fluid field. These effects produce a complicated array of phenomena - large departure of the streamlines from their equilibrium levels, high winds, generation of small scales, turbulence, etc. - that present hazards to aircraft and to lee surface areas. The nonlinear disturbances also interact with the larger-scale flow in such a manner as to impact global weather forecasts and the climatological momentum balance. If there is no dynamic barrier, these waves can penetrate vertically into the middle atmosphere (30-100 km), where recent observations show them to be of a length scale that must involve the coriolis force in any modeling. At these altitudes, the amplitude of the waves is very large, and the phenomena associated with these wave dynamics are being studied with a view to their potential impact on high performance aircraft, including the projected National Aerospace Plane (NASP). The presentation shows the results of analysis and of state-of-the-art numerical simulations, validated where possible by observational data, and illustrated with photographs from nature. Author (revised) Aircraft Hazards; Atmospheric Turbulence; Flow Distribution; Gravity Waves; Incompressible Fluids; Lee Waves

19940003888 VRA, Inc., Blacksburg, VA, USA PNS predictions of external/internal hypersonic ﬂows for NASP propulsion applications Bhutta, Bilal A.; Lewis, Clark H.; Jun 1, 1990; In English Contract(s)/Grant(s): NAS3-25450; SBIR-01.05-2036 Report No.(s): NASA-CR-194246; NAS 1.26:194246; VRA-TR-90-01; No Copyright; Avail: CASI; A10, Hardcopy

157 Because the development of National Aerospace Plane (NASP) propulsion technology and airframe design relies heavily upon computational fluid dynamics, a keen need exists for developing accurate, reliable, and computationally efficient numerical schemes for predicting large angle-of-attack, supersonic/hypersonic external flows over realistic (3-D) NASP configurations. In this SBIR Phase II study a new three dimensional fully iterative PNS scheme has been developed to study hypersonic external flows over 3-D lifting configurations. It has been demonstrated that this baseline PNS scheme can also be extended to predict 2-D/axisymmetric supersonic/hypersonic internal flows. The three dimensional PNS scheme developed is inherently stable in the subsonic as well as the supersonic flow regions and, thus, does not require any sublayer approximation. It has the capability of treating perfect-gas, equilibrium-air, and nonequilibrium-air gas models in a unified manner. A fourth-order crossflow and a second-order streamwise smoothing approach are used to damp the solution oscillations. In addition, a pseudo-unsteady approach is used to dramatically improve the solution efficiency without compromising the solution accuracy. The scheme is formulated in terms of a general curvilinear coordinate system and is capable of treating cylindrical, parabolic, body-normal, a modified body-normal, and elliptic grid-generation algorithms for highly nonaxisymmetric configurations. A new fully implicit and crossflow-coupled shock-fitting approach has also been developed. With this solution approach the shock is fully coupled in the crossflow direction, while within the shock-layer the solution is obtained using a new predictor-corrector solution scheme which can treat strong crossflow solution-coupling effects. The resulting overall computing times are of the order of the noniterative PNS schemes using approximate factorization, because apart from being numerically efficient the present solution scheme allows much larger marching steps to be taken. In addition to these mathematical and numerical developments, significant effort was spent in developing a vectorization strategy for this 3-D PNS scheme. Six cases were considered to test this new PNS scheme for external and internal hypersonic flows. The first three test cases were for the 5-deg angle of attack flow over a 3-D lifting configuration at Mach 20 and 125 kft altitude. Perfect-gas, equilibrium-air, and nonequilibrium-air gas models were used to study these cases, and it was observed that the type of gas model used had a significant influence on the flowfield predictions. The next three test cases considered 2-D supersonic/hypersonic internal flows through different channel configurations. In each of these internal flow cases, comparisons were made with the predictions obtained using classical 2-D inviscid theory, and the agreement between inviscid and viscous predictions was found to be excellent. In general, the results of these numerical tests were very encouraging and substantiated the accuracy, efficiency, and stability claims of the PNS scheme. Author (revised) Aerospace Planes; Channel Flow; Computational Fluid Dynamics; Grid Generation (Mathematics); Hypersonic Flow; Hypersonic Vehicles; Navier-Stokes Equation; Supersonic Flow; Three Dimensional Flow; Two Dimensional Flow

19930022451 Cornell Univ., Ithaca, NY, USA Real-time trajectory optimization on parallel processors Psiaki, Mark L.; Jul 10, 1993; In English Contract(s)/Grant(s): NAG1-1009 Report No.(s): NASA-CR-193303; NAS 1.26:193303; No Copyright; Avail: CASI; A10, Hardcopy A parallel algorithm has been developed for rapidly solving trajectory optimization problems. The goal of the work has been to develop an algorithm that is suitable to do real-time, on-line optimal guidance through repeated solution of a trajectory optimization problem. The algorithm has been developed on an INTEL iPSC/860 message passing parallel processor. It uses a zero-order-hold discretization of a continuous-time problem and solves the resulting nonlinear programming problem using a custom-designed augmented Lagrangian nonlinear programming algorithm. The algorithm achieves parallelism of function, derivative, and search direction calculations through the principle of domain decomposition applied along the time axis. It has been encoded and tested on 3 example problems, the Goddard problem, the acceleration-limited, planar minimum-time to the origin problem, and a National Aerospace Plane minimum-fuel ascent guidance problem. Execution times as fast as 118 sec of wall clock time have been achieved for a 128-stage Goddard problem solved on 32 processors. A 32-stage minimum-time problem has been solved in 151 sec on 32 processors. A 32-stage National Aerospace Plane problem required 2 hours when solved on 32 processors. A speed-up factor of 7.2 has been achieved by using 32-nodes instead of 1-node to solve a 64-stage Goddard problem. Derived from text Aerospace Planes; Lagrangian Function; Message Processing; Nonlinear Programming; On-Line Systems; Parallel Processing (Computers); Real Time Operation; Trajectory Optimization

158 19930017829 NASA Lewis Research Center, Cleveland, OH, USA The design and evolution of the beta two-stage-to-orbit horizontal takeoff and landing launch system Burkardt, Leo A.; Norris, Rick B.; Dec 1, 1992; In English; 4th International Aerospace Planes Conference, 1-4 Dec. 1992, Orlando, FL, USA Contract(s)/Grant(s): RTOP 505-69-40 Report No.(s): NASA-TM-106118; E-7776; NAS 1.15:106118; AIAA PAPER 92-5080; No Copyright; Avail: CASI; A03, Hardcopy The Beta launch system was originally conceived in 1986 as a horizontal takeoff and landing, fully reusable, two-stage-to-orbit, manned launch vehicle to replace the Shuttle. It was to be capable of delivering a 50,000 lb. payload to low polar orbit. The booster propulsion system consisted of JP fueled turbojets and LH fueled ramjets mounted in pods in an over/under arrangement, and a single LOX/LH fueled SSME rocket. The second stage orbiter, which staged at Mach 8, was powered by an SSME rocket. A major goal was to develop a vehicle design consistent with near term technology. The vehicle design was completed with a GLOW of approximately 2,000,000 lbs. All design goals were met. Since then, interest has shifted to the 10,000 lbs. to low polar orbit payload class. The original Beta was down-sized to meet this payload class. The GLOW of the down-sized vehicle was approximately 1,000,000 lbs. The booster was converted to exclusively air-breathing operation. Because the booster depends on conventional air-breathing propulsion only, the staging Mach number was reduced to 5.5. The orbiter remains an SSME rocket-powered stage. Author Air Breathing Engines; Air Launching; Horizontal Spacecraft Landing; Hydrogen Oxygen Engines; Hypersonic Vehicles; Multistage Rocket Vehicles; Ramjet Engines; Reusable Launch Vehicles; Space Shuttle Main Engine; Spacecraft Launching

19930017827 NASA Langley Research Center, Hampton, VA, USA Analysis of the staging maneuver and booster glideback guidance for a two-stage, winged, fully reusable launch vehicle Naftel, J. Christopher; Powell, Richard W.; Apr 1, 1993; In English Contract(s)/Grant(s): RTOP 906-11-01-01 Report No.(s): NASA-TP-3335; L-17066; NAS 1.60:3335; No Copyright; Avail: CASI; A03, Hardcopy One of the promising launch concepts that could replace the current space shuttle launch system is a two-stage, winged, vertical-takeoff, fully reusable launch vehicle. During the boost phase of ascent, the booster provides propellant for the orbiter engines through a cross-feed system. When the vehicle reaches a Mach number of 3, the booster propellants are depleted and the booster is staged and glides unpowered to a horizontal landing at a launch site runway. Two major design issues for this class of vehicle are the staging maneuver and the booster glideback. For the staging maneuver analysis, a technique was developed that provides for a successful separation of the booster from the orbiter over a wide range of staging angles of attack. A longitudinal ﬂight control system was developed for control of the booster during the staging maneuver. For the booster glide back analysis, a guidance algorithm was developed that successfully guides the booster from the completion of the staging maneuver to a launch site runway while encountering many off-nominal atmospheric, aerodynamic, and staging conditions. Author (revised) Booster Rocket Engines; Control Systems Design; Multistage Rocket Vehicles; Reusable Launch Vehicles; Spacecraft Control; Spacecraft Launching; Spacecraft Maneuvers; Vertical Takeoff

19930017364 NASA Langley Research Center, Hampton, VA, USA Supersonic aeroelastic instability results for a NASP-like wing model Cole, Stanley R.; Florance, James R.; Thomason, Lee B.; Spain, Charles V.; Bullock, Ellen P.; Apr 1, 1993; In English; AIAA 34th SDM Conference, 19-23 Apr. 1993, Hampton, VA, USA Contract(s)/Grant(s): RTOP 763-23-41 Report No.(s): NASA-TM-107739; NAS 1.15:107739; No Copyright; Avail: CASI; A03, Hardcopy An experimental study and an analytical study have been conducted to examine static divergence for hypersonic-vehicle wing models at supersonic conditions. A supersonic test in the Langley Unitary Plan Wind Tunnel facility was conducted for two wind-tunnel models. These models were nearly identical with the exception of airfoil shape. One model had a four-percent maximum thickness airfoil and the other model had an eight-percent maximum thickness airfoil. The wing models had low-aspect ratios and highly swept leading edges. The all-movable wing models were supported by a single-pivot mechanism along the wing root. For both of the wind-tunnel models, conﬁguration changes could be made in the wing-pivot location along the wing root and in the wing-pivot pitch stiffness. Three divergence conditions were measured for the four-percent thick airfoil model in the Mach number range of 2.6 to 3.6 and one divergence condition was measured for the eight-percent thick

159 airfoil model at a Mach number of 2.9. Analytical divergence calculations were made for comparison with experimental results and to evaluate the parametric effects of wing-pivot stiffness, wing-pivot location, and airfoil thickness variations. These analyses showed that decreasing airfoil thickness, moving the wing-pivot location upstream, or increasing the pitch-pivot stiffness have the beneﬁcial effect of increasing the divergence dynamic pressures. The calculations predicted the trend of experimental divergence dynamic pressure with Mach number accurately; however, the calculations were approximately 25 percent conservative with respect to dynamic pressure. Author Aeroelasticity; Aerospace Planes; Dynamic Stability; Hypersonic Vehicles; Low Aspect Ratio Wings; Supersonic Speed; Swept Wings; Wind Tunnel Tests

19930016481 Iowa State Univ. of Science and Technology, Ames, IA, USA Trajectory optimization for the National aerospace plane Lu, Ping; Jan 1, 1993; In English Contract(s)/Grant(s): NAG1-1255 Report No.(s): NASA-CR-192954; NAS 1.26:192954; No Copyright; Avail: CASI; A02, Hardcopy While continuing the application of the inverse dynamics approach in obtaining the optimal numerical solutions, the research during the past six months has been focused on the formulation and derivation of closed-form solutions for constrained hypersonic flight trajectories. Since it was found in the research of the first year that a dominant portion of the optimal ascent trajectory of the aerospace plane is constrained by dynamic pressure and heating constraints, the application of the analytical solutions significantly enhances the efficiency in trajectory optimization, provides a better insight to understanding of the trajectory and conceivably has great potential in guidance of the vehicle. Work of this period has been reported in four technical papers. Two of the papers were presented in the AIAA Guidance, Navigation, and Control Conference (Hilton Head, SC, August, 1992) and Fourth International Aerospace Planes Conference (Orlando, FL, December, 1992). The other two papers have been accepted for publication by Journal of Guidance, Control, and Dynamics, and will appear in 1993. This report briefly summarizes the work done in the past six months and work currently underway. Derived from text Aerospace Planes; Hypersonic Flight; National Aerospace Plane Program; Spacecraft Guidance; Trajectory Optimization

19930014233 NASA Langley Research Center, Hampton, VA, USA An overview of aeroelasticity studies for the National Aerospace Plane Ricketts, Rodney H.; Noll, Thomas E.; Huttsell, Lawrence J.; Hutsell, Lawrence J.; Mar 1, 1993; In English; AIAA 34th SDM Conference,, 19-23 Apr. 1993, Hampton, VA, USA Contract(s)/Grant(s): RTOP 763-23-41 Report No.(s): NASA-TM-107728; NAS 1.15:107728; No Copyright; Avail: CASI; A03, Hardcopy The National Aero-Space Plane (NASP), or X-30, is a single-stage-to-orbit vehicle that is designed to takeoff and land on conventional runways. Research in aeroelasticity was conducted by NASA and the Wright Laboratory to support the design of a ﬂight vehicle by the national contractor team. This research includes the development of new computational codes for predicting unsteady aerodynamic pressures. In addition, studies were conducted to determine the aerodynamic heating effects on vehicle aeroelasticity and to determine the effects of fuselage ﬂexibility on the stability of the control systems. It also includes the testing of scale models to better understand the aeroelastic behavior of the X-30 and to obtain data for code validation and correlation. This paper presents an overview of the aeroelastic research which has been conducted to support the airframe design. Author (revised) Aerodynamic Heating; Aeroelasticity; Airframes; Fuselages; National Aerospace Plane Program; Single Stage to Orbit Vehicles; Unsteady Aerodynamics; X-30 Vehicle

19930013369 NASA Langley Research Center, Hampton, VA, USA Subsonic aerodynamic characteristics of a proposed advanced manned launch system orbiter conﬁguration Ware, George M.; Fox, Charles H., Jr.; Feb 1, 1993; In English Contract(s)/Grant(s): RTOP 506-40-61-01 Report No.(s): NASA-TM-4439; L-17182; NAS 1.15:4439; No Copyright; Avail: CASI; A03, Hardcopy The Advanced Manned Launch System is a proposed near-term technology, two-stage, fully reusable launch system that consists of an unmanned glide-back booster and a manned orbiter. An orbiter model that featured a large fuselage and an aft

160 delta wing with tip fins was tested in the Langley 7- by 10-Foot High-Speed Tunnel. A crew cabin, large payload fairing, and crew access tunnel were mounted on the upper body. The results of the investigation indicated that the configuration was longitudinally stable to an angle of attack of about 6 deg about a center-of-gravity position of 0.7 body length. The model had an untrimmed lift-drag ratio of 6.6, but could not be trimmed at positive lift. The orbiter model was also directionally unstable. The payload fairing was responsible for about half the instability. The tip-fin controllers, which are designed as active controls to produce artificial directional stability, were effective in producing yawing moment, but sizable adverse rolling moment occurred at angles of attack above 6 deg. Differential deflection of the elevon surfaces was effective in producing rolling moment with only small values of adverse yawing moment. Author Aerodynamic Characteristics; Directional Stability; Reusable Launch Vehicles; Space Transportation System; Spacecraft Reentry; Wind Tunnel Tests

19930011110 NASA Langley Research Center, Hampton, VA, USA Computational parametric study of sidewall-compression scramjet inlet performance at Mach 10 Holland, Scott D.; Feb 1, 1993; In English Contract(s)/Grant(s): RTOP 506-40-71-04 Report No.(s): NASA-TM-4411; L-17134; NAS 1.15:4411; No Copyright; Avail: CASI; A03, Hardcopy A computational parametric study of three-dimensional, sidewall-compression scramjet inlets was performed to identify the effects of geometric parameters on inlet performance. The parameters were the leading-edge sweep angle, varied between 30 and 60 deg, and the leading-edge position of the cowl, located at the throat and at two forward positions. A laminar boundary layer with cold-wall (T(sub wall) = 300 K (540 R)) boundary conditions was imposed. The parametric study was performed for a Mach number of 10 and a unit free-stream Reynolds number of 7.06 x 10(exp 6) per meter (2.15 x 10(exp 6) per foot) at a geometric contraction ratio of 5. The performance of each configuration was evaluated in terms of the mass capture, throat Mach number, total pressure recovery, kinetic energy efficiency, and internal compression. One computation of an unswept configuration was included as a baseline to determine the effects of introducing leading-edge sweep on the flow-field parameters. The purpose of the computational parametric study was to perform a trade-off of the effects of various parameters on the global performance of the inlet. Although no single optimal configuration emerged, trade-offs among the stated performance parameters identified a leading-edge sweep angle of 45 deg as possessing the most attractive performance characteristics. Author Computational Fluid Dynamics; Engine Inlets; Hypersonic Speed; Leading Edge Sweep; Supersonic Combustion Ramjet Engines; Sweep Angle

19930011048 Sverdrup Technology, Inc., Brook Park, OH, USA Computation of H2/air reacting flowfields in drag-reduction external combustion Lai, H. T.; Jul 1, 1992; In English; 28th Joint Propulsion Conference and Exhibit, 6-8 Jul. 1992, Nashville, TN, USA Report No.(s): NASA-CR-191071; E-7624; NAS 1.26:191071; AIAA PAPER 92-3672; No Copyright; Avail: CASI; A03, Hardcopy Numerical simulation and analysis of the solution are presented for a laminar reacting flowfield of air and hydrogen in the case of external combustion employed to reduce base drag in hypersonic vehicles operating at transonic speeds. The flowfield consists of a transonic air stream at a Mach number of 1.26 and a sonic transverse hydrogen injection along a row of 26 orifices. Self-sustained combustion is computed over an expansion ramp downstream of the injection and a flameholder, using the recently developed RPLUS code. Measured data is available only for surface pressure distributions and is used for validation of the code in practical 3D reacting flowfields. Pressure comparison shows generally good agreements, and the main effects of combustion are also qualitatively consistent with experiment. Author (revised) Aerodynamic Drag; Base Flow; Combustion Physics; Drag Reduction; Flow Distribution; Hydrogen Fuels; Hypersonic Vehicles; Supersonic Combustion; Transonic Speed

19930011046 MCAT Inst., San Jose, CA, USA Turbulence modeling for hypersonic ﬂight Bardina, Jorge E.; Feb 1, 1993; In English Contract(s)/Grant(s): NCC2-585 Report No.(s): NASA-CR-192288; NAS 1.26:192288; MCAT-93-07; No Copyright; Avail: CASI; A03, Hardcopy

161 The objective of the proposed work is to continue to develop, verify, and incorporate the baseline two-equation turbulence models, which account for the effects of compressibility at high speeds, into a three-dimensional Reynolds averaged Navier-Stokes (RANS) code. Additionally, we plan to provide documented descriptions of the models and their numerical procedures so that they can be implemented into the NASP CFD codes. Derived from text Aerospace Planes; Compressibility; Hypersonic Flight; National Aerospace Plane Program; Navier-Stokes Equation; Turbulence Models

19930010263 North Carolina Agricultural and Technical State Univ., Greensboro, NC, USA, Futron Corp., Bethesda, MD, USA The Center for Aerospace Research: A NASA Center of Excellence at North Carolina Agricultural and Technical State University Lai, Steven H.-Y.; Dec 12, 1992; In English; See also N93-19453 through N93-19466 Contract(s)/Grant(s): NAGW-2924 Report No.(s): NASA-CR-191362; NAS 1.26:191362; No Copyright; Avail: CASI; A11, Hardcopy This report documents the efforts and outcomes of our research and educational programs at NASA-CORE in NCA&TSU. The goal of the center was to establish a quality aerospace research base and to develop an educational program to increase the participation of minority faculty and students in the areas of aerospace engineering. The major accomplishments of this center in the ﬁrst year are summarized in terms of three different areas, namely, the center’s research programs area, the center’s educational programs area, and the center’s management area. In the center’s research programs area, we focus on developing capabilities needed to support the development of the aerospace plane and high speed civil transportation system technologies. In the educational programs area, we developed an aerospace engineering option program ready for university approval. Aerospace Engineering; Aerospace Planes; Education; Hypersonic Vehicles; Research and Development; University Program

19930009413 Iowa State Univ. of Science and Technology, Ames, IA, USA Analytical solutions to constrained hypersonic flight trajectories Lu, Ping; Nov 30, 1992; In English Contract(s)/Grant(s): NAG1-1255 Report No.(s): NASA-CR-191987; NAS 1.26:191987; No Copyright; Avail: CASI; A03, Hardcopy The flight trajectory of aerospace vehicles subject to a class of path constraints is considered. The constrained dynamics is shown to be a natural two-time-scale system. Asymptotic analytical solutions are obtained. Problems of trajectory optimization and guidance can be dramatically simplified with these solutions. Applications in trajectory design for an aerospace plane strongly support the theoretical development. Author Aerospace Planes; Aerospace Vehicles; Ascent Trajectories; Cruising Flight; Hypersonic Flight; Trajectory Analysis; Trajectory Optimization

19930009189 Eloret Corp., Sunnyvale, CA, USA Hypersonic flows as related to the national aerospace plane Kussoy, Marvin; Menter, F.; Huang, P. G.; Jul 30, 1992; In English Contract(s)/Grant(s): NCC2-452 Report No.(s): NASA-CR-191980; NAS 1.26:191980; No Copyright; Avail: CASI; A02, Hardcopy The study in the last 6 months has observed a clear evidence that the current two-equation models tend to under-predict flow separation and over-predict heat transfer rate near flow re-attachment regions. In hypersonic flow calculations, these model deficiencies appear to be even more pronounced. This is particularly true in the incapability of the model to predict the extent of the flow separation. Two major deficiencies of the current two-equation models in predicting complex hypersonic flows have been reported, i.e., under-prediction of flow separation and over-prediction of peak heat transfer rate. Two modifications to the k - epsilon model were reported and tested over a range of flows. Based on our limited study, the modified models have been found to give better agreements in both surface pressure and heat transfer predictions for several complex shock-wave boundary-layer interaction flows. However, in order to confirm our observation, more calculations will be

162 performed in the future study covering a wider range of ﬂows and conditions than reported here. Author Aerospace Planes; Boundary Layer Separation; Hypersonic Flow; National Aerospace Plane Program; Pressure Distribution; Separated Flow

19930007564 NASA Flight Research Center, Edwards, CA, USA The F-18 systems research aircraft facility Sitz, Joel R.; Dec 1, 1992; In English; 1992 Aerotech Conference, 5-8 Oct. 1992, Anaheim, CA, USA Contract(s)/Grant(s): RTOP 505-68-52 Report No.(s): NASA-TM-4433; H-1844; NAS 1.15:4433; No Copyright; Avail: CASI; A03, Hardcopy To help ensure that new aerospace initiatives rapidly transition to competitive U.S. technologies, NASA Dryden Flight Research Facility has dedicated a systems research aircraft facility. The primary goal is to accelerate the transition of new aerospace technologies to commercial, military, and space vehicles. Key technologies include more-electric aircraft concepts, fly-by-light systems, flush airdata systems, and advanced computer architectures. Future aircraft that will benefit are the high-speed civil transport and the National AeroSpace Plane. This paper describes the systems research aircraft flight research vehicle and outlines near-term programs. Author Aerospace Engineering; Aircraft Design; F-18 Aircraft; Research Aircraft; Research Facilities

19930005622 McDonnell Aircraft Co., Saint Louis, MO, USA Advanced aeroservoelastic stabilization techniques for hypersonic flight vehicles Chan, Samuel Y.; Cheng, Peter Y.; Myers, Thomas T.; Klyde, David H.; Magdaleno, Raymond E.; Mcruer, Duane T.; Nov 1, 1992; In English Contract(s)/Grant(s): NAS1-18763; RTOP 505-70-64-01 Report No.(s): NASA-CR-189702; NAS 1.26:189702; No Copyright; Avail: CASI; A06, Hardcopy Advanced high performance vehicles, including Single-Stage-To-Orbit (SSTO) hypersonic flight vehicles, that are statically unstable, require higher bandwidth flight control systems to compensate for the instability resulting in interactions between the flight control system, the engine/propulsion dynamics, and the low frequency structural modes. Military specifications, such as MIL-F-9490D and MIL-F-87242, tend to limit treatment of structural modes to conventional gain stabilization techniques. The conventional gain stabilization techniques, however, introduce low frequency effective time delays which can be troublesome from a flying qualities standpoint. These time delays can be alleviated by appropriate blending of gain and phase stabilization techniques (referred to as Hybrid Phase Stabilization or HPS) for the low frequency structural modes. The potential of using HPS for compensating structural mode interaction was previously explored. It was shown that effective time delay was significantly reduced with the use of HPS; however, the HPS design was seen to have greater residual response than a conventional gain stablized design. Additional work performed to advance and refine the HPS design procedure, to further develop residual response metrics as a basis for alternative structural stability specifications, and to develop strategies for validating HPS design and specification concepts in manned simulation is presented. Stabilization design sensitivity to structural uncertainties and aircraft-centered requirements are also assessed. Author Aeroservoelasticity; Aircraft Control; Design Analysis; Dynamic Models; Dynamic Structural Analysis; Flexible Spacecraft; Flight Control; Hypersonic Vehicles; Stabilization; Structural Stability

19930004117 Iowa State Univ. of Science and Technology, Ames, IA, USA An inverse dynamics approach to trajectory optimization for an aerospace plane Lu, Ping; JAN 1, 1992; In English Contract(s)/Grant(s): NAG1-1255 Report No.(s): NASA-CR-191277; NAS 1.26:191277; No Copyright; Avail: CASI; A03, Hardcopy An inverse dynamics approach for trajectory optimization is proposed. This technique can be useful in many difficult trajectory optimization and control problems. The application of the approach is exempliﬁed by ascent trajectory optimization for an aerospace plane. Both minimum-fuel and minimax types of performance indices are considered. When rocket augmentation is available for ascent, it is shown that accurate orbital insertion can be achieved through the inverse control of the rocket in the presence of disturbances. Author Aerospace Planes; Minimax Technique; Spacecraft Control; Spacecraft Trajectories; Trajectory Control; Trajectory Optimization

163 19930004100 NASA Hugh L. Dryden Flight Research Facility, Edwards, CA, USA Stratospheric turbulence measurements and models for aerospace plane design Ehernberger, L. J.; Dec 1, 1992; In English; 4th AIAA International Aerospace Planes Conference, 1-4 Dec. 1992, Orlando, FL, USA Contract(s)/Grant(s): RTOP 763-21-51 Report No.(s): NASA-TM-104262; H-1865; NAS 1.15:104262; Copyright; Avail: CASI; A03, Hardcopy Progress in computational atmospheric dynamics is exhibiting the ability of numerical simulation to describe instability processes associated with turbulence observed at altitudes between 15 and 25 km in the lower stratosphere. As these numerical simulation tools mature, they can be used to extend estimates of atmospheric perturbations from the present gust database for airplane design at altitudes below 15 km to altitudes between 25 and 50 km where aerospace plane operation would be at hypersonic speeds. The amount of available gust data and number of temperature perturbation observations are limited at altitudes between 15 and 25 km. On the other hand, in-situ gust data at higher altitudes are virtually nonexistent. The uncertain potential for future airbreathing hypersonic ﬂight research vehicles to encounter strong turbulence at higher altitudes could penalize the design of these vehicles by undue cost or limitations on performance. Because the atmospheric structure changes markedly with altitude, direct extrapolation of gust magnitudes and encounter probabilities to the higher ﬂight altitudes is not advisable. This paper presents a brief review of turbulence characteristics observed in the lower stratosphere and highlights the progress of computational atmospheric dynamics that may be used to estimate the severity of atmospheric transients at higher altitudes. Author Aerospace Planes; Atmospheric Models; Atmospheric Turbulence; Flight Conditions; Gusts; Stratosphere; Wind Velocity Measurement

19930003922 NASA, Washington, DC, USA NASA aeronautics: Research and technology program highlights JAN 1, 1990; In English; 44 functional color pages Report No.(s): NASA-NP-159; NAS 1.83:159; No Copyright; Avail: CASI; A05, Hardcopy; 44 functional color pages This report contains numerous color illustrations to describe the NASA programs in aeronautics. The basic ideas involved are explained in brief paragraphs. The seven chapters deal with Subsonic aircraft, High-speed transport, High-performance military aircraft, Hypersonic/Transatmospheric vehicles, Critical disciplines, National facilities and Organizations & installations. Some individual aircraft discussed are : the SR-71 aircraft, aerospace planes, the high-speed civil transport (HSCT), the X-29 forward-swept wing research aircraft, and the X-31 aircraft. Critical disciplines discussed are numerical aerodynamic simulation, computational ﬂuid dynamics, computational structural dynamics and new experimental testing techniques. R.L.B. Aeronautics; Aerospace Planes; Computational Fluid Dynamics; Computerized Simulation; Dynamic Structural Analysis; NASA Programs; Research Aircraft; X-29 Aircraft

19930003880 NASA Langley Research Center, Hampton, VA, USA Hypersonic vehicle control law development using H infinity and mu-synthesis Gregory, Irene M.; Chowdhry, Rajiv S.; Mcminn, John D.; Shaughnessy, John D.; Oct 1, 1992; In English; 4th AIAA International Aerospace Planes Conference, 1-4 Dec. 1992, Orlando, FL, USA Contract(s)/Grant(s): RTOP 505-64-40-01 Report No.(s): NASA-TM-107689; NAS 1.15:107689; No Copyright; Avail: CASI; A03, Hardcopy Applicability and effectiveness of robust control techniques to a single-stage-to-orbit (SSTO) airbreathing hypersonic vehicle on an ascent accelerating path and their effectiveness are explored in this paper. An SSTO control system design problem, requiring high accuracy tracking of velocity and altitude commands while limiting angle of attack oscillations, minimizing control power usage and stabilizing the vehicle all in the presence of atmospheric turbulence and uncertainty in the system, was formulated to compare results of the control designs using H infinity and mu-synthesis procedures. The math model, an integrated flight/propulsion dynamic model of a conical accelerator class vehicle, was linearized as the vehicle accelerated through Mach 8. Controller analysis was conducted using the singular value technique and the mu-analysis approach. Analysis results were obtained in both the frequency and the time domains. The results clearly demonstrate the inherent advantages of the structured singular value framework for this class of problems. Since payload performance margins

164 are so critical for the SSTO mission, it is crucial that adequate stability margins be provided without sacriﬁcing any payload mass. Author Control Systems Design; Control Theory; H-Inﬁnity Control; Hypersonic Flight; Hypersonic Vehicles; Single Stage to Orbit Vehicles; Spacecraft Control

19930003259 NASA Langley Research Center, Hampton, VA, USA Current Technology for Thermal Protection Systems Scotti, Stephen J., compiler; Oct 1, 1992; In English, 11-12 Feb. 1992, Hampton, VA, USA; See also N93-12448 through N93-12460 Contract(s)/Grant(s): RTOP 506-43-31-04 Report No.(s): NASA-CP-3157; L-17118; NAS 1.55:3157; No Copyright; Avail: CASI; A14, Hardcopy Interest in thermal protection systems for high-speed vehicles is increasing because of the stringent requirements of such new projects as the Space Exploration Initiative, the National Aero-Space Plane, and the High-Speed Civil Transport, as well as the needs for improved capabilities in existing thermal protection systems in the Space Shuttle and in turbojet engines. This selection of 13 papers from NASA and industry summarizes the history and operational experience of thermal protection systems utilized in the national space program to date, and also covers recent development efforts in thermal insulation, refractory materials and coatings, actively cooled structures, and two-phase thermal control systems. Aerospace Planes; Composite Materials; Composite Structures; Refractory Coatings; Refractory Materials; Space Shuttles; Supersonic Transports; Temperature Control; Thermal Insulation; Thermal Protection; Two Phase Flow

19930003225 Draper (Charles Stark) Lab., Inc., Cambridge, MA, USA Air-breathing hypersonic vehicle guidance and control studies: An integrated trajectory/control analysis methodology, phase 2 Hattis, Philip D.; Malchow, Harvey L.; Oct 1, 1992; In English Contract(s)/Grant(s): NAS1-18565; RTOP 505-70-64-01 Report No.(s): NASA-CR-189703; NAS 1.26:189703; No Copyright; Avail: CASI; A05, Hardcopy An integrated trajectory/control analysis algorithm has been used to generate trajectories and desired control strategies for two different hypersonic air-breathing vehicle models and orbit targets. Both models used cubic spline curve fit tabulated winged-cone accelerator vehicle representations. Near-fuel-optimal, horizontal takeoff trajectories, imposing a dynamic pressure limit of 1000 psf, were developed. The first model analysis case involved a polar orbit and included the dynamic effects of using elevons to maintain longitudinal trim. Analysis results indicated problems with the adequacy of the propulsion model and highlighted dynamic pressure/altitude instabilities when using vehicle angle of attack as a control variable. Also, the magnitude of computed elevon deflections to maintain trim suggested a need for alternative pitch moment management strategies. The second analysis case was reformulated to use vehicle pitch attitude relative to the local vertical as the control variable. A new, more realistic, air-breathing propulsion model was incorporated. Pitch trim calculations were dropped and an equatorial orbit was specified. Changes in flight characteristics due to the new propulsion model have been identified. Flight regimes demanding rapid attitude changes have been noted. Also, some issues that would affect design of closed-loop controllers were ascertained. Author Air Breathing Engines; Control Systems Design; Feedback Control; Hypersonic Vehicles; Single Stage to Orbit Vehicles; Spacecraft Control; Spacecraft Design; Spacecraft Guidance; Spacecraft Trajectories; Trajectory Analysis; Trajectory Control

19930001912 NASA Langley Research Center, Hampton, VA, USA Winds of change: Expanding the frontiers of flight. Langley Research Center’s 75 years of accomplishment, 1917-1992 Schultz, James; JAN 1, 1992; In English Report No.(s): NASA-NP-130; NAS 1.83:130; No Copyright; Avail: CASI; A07, Hardcopy; Limited Reproducibility: More than 20% of this doc This commemorative volume highlights in pictures and text seventy five years of accomplishments of the Langley Research Center. The introductory matter features wind tunnels and their contribution to the development of aeronautics. A chronological survey details four different periods in Langley’s history. The first period, 1917-1939, is subtitled ‘Perfecting the Plane’ which details Langley’s contribution to early aeronautics with examples from specific aircraft. The second period,

165 1940-1957, focuses on the development of military aircraft during and after World War II. The third period, 1958-1969, tells the story of Langley’s involvement with NASA and the satellite and Apollo era. The fourth period, entitled ‘Charting New Courses: 1970-1992 and Beyond’, treats various new topics from aerospace planes to Mars landing, as well as older topics such as the Space Shuttle and research spinoffs. R.L.B. Aeronautics; Aerospace Sciences; Aircraft; Histories; Military Aircraft; NASA Programs; Wind Tunnels

19930001855 Alabama Univ., Huntsville, AL, USA Rocket Based Combined Cycle (RBCC) Propulsion Workshop, volume 2 Chojnacki, Kent T.; Aug 28, 1992; In English Contract(s)/Grant(s): NAS8-38609 Report No.(s): NASA-CR-184383; NAS 1.26:184383; UAH-TR-91/001-VOL-2; No Copyright; Avail: CASI; A04, Hardcopy The goal of the Rocket Based Combined Cycle (RBCC) Propulsion Technology Workshop, was to impart technology information to the propulsion community with respect to hypersonic combined cycle propulsion capabilities. The major recommendation resulting from this technology workshop was as follows: conduct a systems-level applications study to deﬁne the desired propulsion system and vehicle technology requirements for LEO launch vehicles. All SSTO and TSTO options using the various propulsion systems (airbreathing combined cycle, rocket-based combined cycle, and all rocket) must be considered. Such a study should be accomplished as soon as possible. It must be conducted with a consistent set of ground rules and assumptions. Additionally, the study should be conducted before any major expenditures on a RBCC technology development program occur. Author Air Breathing Engines; Conferences; Hypersonic Flight; Launch Vehicles; Propulsion System Conﬁgurations; Propulsion System Performance; Single Stage to Orbit Vehicles

19930001854 Alabama Univ., Huntsville, AL, USA Rocket Based Combined Cycle (RBCC) Propulsion Technology Workshop. Volume 1: Executive summary Chojnacki, Kent T.; Aug 28, 1992; In English Contract(s)/Grant(s): NAS8-38609 Report No.(s): NASA-CR-184382; NAS 1.26:184382; UAH-TR-92/001-VOL-1; No Copyright; Avail: CASI; A03, Hardcopy The goal of the Rocket-Based Combined Cycle (RBCC) Propulsion Technology Workshop was to assess the RBCC propulsion system’s viability for Earth-to-Orbit (ETO) transportation systems. This was accomplished by creating a forum (workshop) in which past work in the field of RBCC propulsion systems was reviewed, current technology status was evaluated, and future technology programs in the field of RBCC propulsion systems were postulated, discussed, and recommended. Author Aerospace Planes; Air Breathing Engines; Propulsion System Configurations; Propulsion System Performance; Rocket Engine Design; Single Stage to Orbit Vehicles

19920018966 Draper (Charles Stark) Lab., Inc., Cambridge, MA, USA Robust intelligent flight control for hypersonic vehicles Chamitoff, Gregory Errol; Feb 20, 1992; In English Contract(s)/Grant(s): NAS9-18426 Report No.(s): NASA-CR-185680; NAS 1.26:185680; CSDL-T-1106; Copyright; Avail: CASI; A15, Hardcopy Intelligent optimization methods are applied to the problem of real-time flight control for a class of airbreathing hypersonic vehicles (AHSV). The extreme flight conditions that will be encountered by single-stage-to-orbit vehicles, such as the National Aerospace Plane, present a tremendous challenge to the entire spectrum of aerospace technologies. Flight control for these vehicles is particularly difficult due to the combination of nonlinear dynamics, complex constraints, and parametric uncertainty. An approach that utilizes all available a priori and in-flight information to perform robust, real time, short-term trajectory planning is presented. CASI Aerospace Planes; Air Breathing Engines; Control Systems Design; Flight Control; Hypersonic Vehicles; Real Time Operation; Single Stage to Orbit Vehicles; Spacecraft Control; Trajectory Planning

166 19920016739 Systems Technology, Inc., Hawthorne, CA, USA, PRC Kentron, Inc., Edwards, CA, USA Assessment of flying-quality criteria for air-breathing aerospacecraft Mcruer, Duane T.; Myers, Thomas T.; Hoh, Roger H.; Ashkenas, Irving L.; Johnston, Donald E.; Jun 1, 1992; In English Contract(s)/Grant(s): NAS2-12722; ATD-88-STI-0221; RTOP 505-64-40 Report No.(s): NASA-CR-4442; H-1758; NAS 1.26:4442; STI-TR-2361-2; No Copyright; Avail: CASI; A05, Hardcopy A study of flying quality requirements for air breathing aerospacecraft gives special emphasis to the unusual operational requirements and characteristics of these aircraft, including operation at hypersonic speed. The report considers distinguishing characteristics of these vehicles, including dynamic deficiencies and their implications for control. Particular emphasis is given to the interaction of the airframe and propulsion system, and the requirements for dynamic systems integration. Past operational missions are reviewed to define tasks and maneuvers to be considered for this class of aircraft. Areas of special concern with respect to vehicle dynamics and control are identified. Experience with the space shuttle orbiter is reviewed with respect to flight control system mechanization and flight experience in approach and landing flying qualities for the National Aerospace Plane (NASP). CASI Aerospace Planes; Air Breathing Engines; Flight Characteristics; Flight Control; Hypersonic Speed; National Aerospace Plane Program; Systems Integration

19920016714 NASA Lewis Research Center, Cleveland, OH, USA Slush hydrogen transfer studies at the NASA K-Site Test Facility Hardy, Terry L.; Whalen, Margaret V.; JAN 1, 1992; In English; 28th Joint Propulsion Conference and Exhibit, 6-8 Jul. 1992, Nashville, TN, USA Contract(s)/Grant(s): RTOP 763-22-21 Report No.(s): NASA-TM-105596; E-6932; NAS 1.15:105596; AIAA PAPER 92-3384; No Copyright; Avail: CASI; A02, Hardcopy An experimental study was performed as part of the National Aerospace Plane (NASP) effort to determine slush hydrogen production and transfer characteristics. Flow rate and pressure drop characteristics were determined for slush hydrogen ﬂow through a vacuum-jacketed transfer system. These characteristics were compared to similar tests using normal boiling point and triple point hydrogen. In addition, experimental ﬂow characteristic data was compared with predictions from the FLUSH analytical model. Slush hydrogen density loss during the transfer process was also examined. CASI Aerospace Planes; Cryogenic Fluid Storage; Fuel Production; Hydrogen Fuels; Liquid Hydrogen; National Aerospace Plane Program; Slush Hydrogen; Transferring

19920016034 NASA Lewis Research Center, Cleveland, OH, USA Technology issues associated with using densified hydrogen for space vehicles Hardy, Terry L.; Whalen, Margaret V.; JAN 1, 1992; In English; 28th Joint Propulsion Conference and Exhibit, 6-8 Jul. 1992, Nashville, TN, USA Contract(s)/Grant(s): RTOP 763-22-21 Report No.(s): NASA-TM-105642; E-6988; NAS 1.15:105642; AIAA PAPER 92-3079; No Copyright; Avail: CASI; A03, Hardcopy Slush hydrogen and triple-point hydrogen offer the potential for reducing the size and weight of future space vehicles because these fluids have greater densities than normal-boiling-point liquid hydrogen. In addition, these fluids have greater heat capacities, which make them attractive fuels for such applications as the National Aerospace Plane and cryogenic depots. Some of the benefits of using slush hydrogen and triple-point hydrogen for space missions are quantified. Some of the major issues associated with using these densified cryogenic fuels for space applications are examined, and the technology efforts that have been made to address many of these issues are summarized. CASI Aerospace Planes; Cryogenics; Hydrogen; Hydrogen Fuels; Liquid Hydrogen; Slush Hydrogen; Spacecraft Propulsion

19920015921 NASA Lewis Research Center, Cleveland, OH, USA The future challenge for aeropropulsion Rosen, Robert; Bowditch, David N.; JAN 1, 1992; In English; Aeroengine 1992, 6-12 Apr. 1992, Moscow, USSR Contract(s)/Grant(s): RTOP 505-62-00 Report No.(s): NASA-TM-105613; E-6943; NAS 1.15:105613; No Copyright; Avail: CASI; A03, Hardcopy

167 NASA’s research in aeropropulsion is focused on improving the efficiency, capability, and environmental compatibility for all classes of future aircraft. The development of innovative concepts, and theoretical, experimental, and computational tools provide the knowledge base for continued propulsion system advances. Key enabling technologies include advances in internal fluid mechanics, structures, light-weight high-strength composite materials, and advanced sensors and controls. Recent emphasis has been on the development of advanced computational tools in internal fluid mechanics, structural mechanics, reacting flows, and computational chemistry. For subsonic transport applications, very high bypass ratio turbofans with increased engine pressure ratio are being investigated to increase fuel efficiency and reduce airport noise levels. In a joint supersonic cruise propulsion program with industry, the critical environmental concerns of emissions and community noise are being addressed. NASA is also providing key technologies for the National Aerospaceplane, and is studying propulsion systems that provide the capability for aircraft to accelerate to and cruise in the Mach 4-6 speed range. The combination of fundamental, component, and focused technology development underway at NASA will make possible dramatic advances in aeropropulsion efficiency and environmental compatibility for future aeronautical vehicles. CASI Aerospace Planes; Aircraft Engines; Aircraft Noise; Fluid Mechanics; NASA Programs; Propulsion; Propulsion System Configurations; Propulsion System Performance; Structural Analysis; Turbofans

19920012304 NASA Lewis Research Center, Cleveland, OH, USA Mission and sizing analysis for the Beta 2 two-stage-to-orbit vehicle Nadell, Shari-Beth; Baumgarten, William J.; Alexander, Stephen W.; JAN 1, 1992; In English; 1992 Aerospace Design Conference, 3-6 Feb. 1992, Irvine, CA, USA Contract(s)/Grant(s): RTOP 505-69-40 Report No.(s): NASA-TM-105559; E-6883; NAS 1.15:105559; AIAA PAPER 92-1264; No Copyright; Avail: CASI; A03, Hardcopy NASA Lewis Research Center studied a horizontal takeoff and landing, fully reusable, two-stage-to-orbit (TSTO) vehicle capable of launching and returning a 10,000 pound payload to low Earth polar orbit using low-risk technology. The vehicle, called Beta 2, was derived from the USAF/Boeing Beta vehicle, a TSTO study vehicle capable of launching a 50,000 pound payload to low Earth polar orbit. Development of the Beta 2 from the USAF/Boeing Beta vehicle occurred in a series of iterations during which the size of the vehicle was decreased to accommodate the smaller payload, the staging Mach number was decreased from 8.0 to 6.5, and the rocket propulsion system was removed from the booster. The ﬁnal Beta 2 vehicle consisted of a rocket powered orbiter and an all airbreathing booster. The gross takeoff weight of the Beta 2 vehicle was approximately 1.1 million pounds. In addition to its baseline mission, the Beta 2 was capable of delivering approximately 17,500 pounds to the Space Station with the same takeoff gross weight. The mission and sizing analysis performed to arrived at the Beta 2 vehicle is discussed. CASI Aerospace Vehicles; Air Breathing Boosters; Aircraft Design; Launch Vehicle Conﬁgurations; Reusable Launch Vehicles; Reusable Spacecraft; Spacecraft Design; Spacecraft Launching

19920012274 NASA Lewis Research Center, Cleveland, OH, USA Rocket-Based Combined-Cycle (RBCC) Propulsion Technology Workshop. Tutorial session JAN 1, 1992; In English, 23-27 Mar. 1992, Huntsville, AL, USA; See also N92-21518 through N92-21538 Contract(s)/Grant(s): RTOP 590-21-11 Report No.(s): NASA-CP-10090; E-6929; NAS 1.55:10090; No Copyright; Avail: CASI; A12, Hardcopy The goal of this workshop was to illuminate the nation’s space transportation and propulsion engineering community on the potential of hypersonic combined cycle (airbreathing/rocket) propulsion systems for future space transportation applications. Four general topics were examined: (1) selections from the expansive advanced propulsion archival resource; (2) related propulsion systems technical backgrounds; (3) RBCC engine multimode operations related subsystem background; and (4) focused review of propulsion aspects of current related programs. Aerospace Planes; Engine Parts; Hypersonic Flight; Rocket Engine Design; Rocket Engines

19920011425 Washington Univ., Seattle, WA, USA Project Antares: A low cost modular launch vehicle for the future Aarnio, Steve; Anderson, Hobie; Arzaz, El Mehdi; Bailey, Michelle; Beeghly, Jeff; Cartwright, Curt; Chau, William; Dawdy, Andrew; Detert, Bruce; Ervin, Miles, et al.; Jun 14, 1991; In English Contract(s)/Grant(s): NASW-4435 Report No.(s): NASA-CR-190018; NAS 1.26:190018; No Copyright; Avail: CASI; A10, Hardcopy

168 The single stage to orbit launch vehicle Antares is based upon the revolutionary concept of modularity, enabling the Antares to efficiently launch communications satellites, as well as heavy payloads, into Earth’s orbit and beyond. The basic unit of the modular system, a single Antares vehicle, is aimed at launching approximately 10,000 kg into low Earth orbit (LEO). When coupled with a Centaur upper stage it is capable of placing 3500 kg into geostationary orbit. The Antares incorporates a reusable engine, the Dual Mixture Ratio Engine (DMRE), as its propulsive device. This enables Antares to compete and excel in the satellite launch market by dramatically reducing launch costs. Antares’ projected launch costs are $1340 per kg to LEO which offers a tremendous savings over launch vehicles available today. Inherent in the design is the capability to attach several of these vehicles together to provide heavy lift capability. Any number of these vehicles, up to seven, can be attached depending on the payload and mission requirements. With a seven vehicle configuration Antares’s modular concept provides a heavy lift capability of approximately 70,000 kg to LEO. This expandability allows for a wider range of payload options such as large Earth satellites, Space Station Freedom support, and interplanetary spacecraft, and also offers a significant cost savings over a mixed fleet based on different launch vehicles. CASI Antares Rocket Vehicle; Heavy Lift Launch Vehicles; Single Stage to Orbit Vehicles; Spacecraft Design

19920011424 Case Western Reserve Univ., Cleveland, OH, USA Conceptual design of two-stage-to-orbit hybrid launch vehicle Jul 1, 1991; In English Contract(s)/Grant(s): NASW-4435 Report No.(s): NASA-CR-190006; NAS 1.26:190006; No Copyright; Avail: CASI; A12, Hardcopy The object of this design class was to design an earth-to orbit vehicle to replace the present NASA space shuttle. The major motivations for designing a new vehicle were to reduce the cost of putting payloads into orbit and to design a vehicle that could better service the space station with a faster turn-around time. Another factor considered in the design was that near-term technology was to be used. Materials, engines and other important technologies were to be realized in the next 10 to 15 years. The first concept put forth by NASA to meet these objectives was the National Aerospace Plane (NASP). The NASP is a single-stage earth-to-orbit air-breathing vehicle. This concept ran into problems with the air-breathing engine providing enough thrust in the upper atmosphere, among other things. The solution of this design class is a two-stage-to-orbit vehicle. The first stage is air-breathing and the second stage is rocket-powered, similar to the space shuttle. The second stage is mounted on the top of the first stage in a piggy-back style. The vehicle takes off horizontally using only air-breathing engines, flies to Mach six at 100,000 feet, and launches the second stage towards its orbital path. The first stage, or booster, will weigh approximately 800,000 pounds and the second stage, or orbiter will weigh approximately 300,000 pounds. The major advantage of this design is the full recoverability of the first stage compared with the present solid rocket booster that are only partially recoverable and used only a few times. This reduces the cost as well as providing a more reliable and more readily available design for servicing the space station. The booster can fly an orbiter up, turn around, land, refuel, and be ready to launch another orbiter in a matter of hours. CASI Aerospace Planes; Cost Reduction; Design Analysis; Hypersonic Vehicles; Launch Vehicles; Orbital Launching; Recoverability; Reusable Spacecraft; Transatmospheric Vehicles

19920010472 State Univ. of New York, Buffalo, NY, USA Development of an integrated BEM approach for hot fluid structure interaction Dargush, Gary F.; Banerjee, Prasanta K.; Honkala, Keith A.; Nov 1, 1991; In English Contract(s)/Grant(s): NAG3-712; RTOP 553-13-00 Report No.(s): NASA-CR-189052; NAS 1.26:189052; No Copyright; Avail: CASI; A06, Hardcopy The development of a boundary element formulation for the study of hot fluid-structure interaction in earth-to-orbit engine hot section components is described. The initial primary thrust of the program to date was directed quite naturally toward the examination of fluid flow, since boundary element methods for fluids are at a much less developed state. This required the development of integral formulations for both the solid and fluid, and some preliminary infrastructural enhancements to a boundary element code to permit coupling of the fluid-structure problem. Boundary element formulations are implemented in two dimensions for both the solid and the fluid. The solid is modeled as an uncoupled thermoelastic medium under plane strain conditions, while several formulations are investigated for the fluid. For example, both vorticity and primitive variable approaches are implemented for viscous, incompressible flow, and a compressible version is developed. All of the above boundary element implementations are incorporated in a general purpose two-dimensional code. Thus, problems involving

169 intricate geometry, multiple generic modeling regions, and arbitrary boundary conditions are all supported. CASI Boundary Element Method; Compressible Flow; Elastic Media; Fluid-Solid Interactions; Incompressible Flow; Single Stage to Orbit Vehicles; Spacecraft Propulsion; Thermal Stresses; Thermoelasticity; Viscous Flow

19920010243 Rockwell International Corp., Canoga Park, CA, USA Operationally Efficient Propulsion System Study (OEPSS) data book. Volume 2: Ground operations problems Waldrop, Glen S.; Apr 24, 1990; In English Contract(s)/Grant(s): NAS10-11568 Report No.(s): NASA-CR-188750; NAS 1.26:188750; RI/RD90-149-2-VOL-2; No Copyright; Avail: CASI; A05, Hardcopy Operations problems and cost drivers were identified for current propulsion systems and design and technology approaches were identified to increase the operational efficiency and to reduce operations costs for future propulsion systems. To provide readily usable data for the ALS program, the results of the OEPSS study were organized into a series of OEPSS Data Books. This volume presents a detailed description of 25 major problems encountered during launch processing of current expendable and reusable launch vehicles. A concise description of each problem and its operational impact on launch processing is presented, along with potential solutions and technology recommendation. CASI Propulsion System Configurations; Propulsion System Performance; Reusable Launch Vehicles; Spacecraft Launching; Spacecraft Propulsion

19920007788 Lockheed Aeronautical Systems Co., Burbank, CA, USA Heat flux sensor research and development: The cool film calorimeter Abtahi, A.; Dean, P.; Jun 30, 1990; In English Contract(s)/Grant(s): NAS1-18570 Report No.(s): NASA-CR-189789; NAS 1.26:189789; LR-31824; No Copyright; Avail: CASI; A04, Hardcopy The goal was to meet the measurement requirement of the NASP program for a gauge capable of measuring heat flux into a ‘typical’ structure in a ‘typical’ hypersonic flight environment. A device is conceptually described that has fast response times and is small enough to fit in leading edge or cowl lip structures. The device relies heavily on thin film technology. The main conclusion is the description of the limitations of thin film technology both in the art of fabrication and in the assumption that thin films have the same material properties as the original bulk material. Three gauges were designed and fabricated. Thin film deposition processes were evaluated. The effect of different thin film materials on the performance and fabrication of the gauge was studied. The gauges were tested in an arcjet facility. Survivability and accuracy were determined under various hostile environment conditions. CASI Calorimeters; Heat Flux; Leading Edges; Thin Films

19920005858 NASA Langley Research Center, Hampton, VA, USA Trim drag reduction concepts for horizontal takeoff single-stage-to-Orbit vehicles Shaughnessy, John D.; Gregory, Irene M.; Dec 1, 1991; In English Contract(s)/Grant(s): RTOP 505-64-40-01 Report No.(s): NASA-TM-102687; NAS 1.15:102687; No Copyright; Avail: CASI; A03, Hardcopy The results of a study to investigate concepts for minimizing trim drag of horizontal takeoff single-stage-to-orbit (SSTO) vehicles are presented. A generic hypersonic airbreathing conical configuration was used as the subject aircraft. The investigation indicates that extreme forward migration of the aerodynamic center as the vehicle accelerates to orbital velocities causes severe aerodynamic instability and trim moments that must be counteracted. Adequate stability can be provided by active control of elevons and rudder, but use of elevons to produce trim moments results in excessive trim drag and fuel consumption. To alleviate this problem, two solution concepts are examined. Active control of the center of gravity (COG) location to track the aerodynamic center decreases trim moment requirements, reduces elevon deflections, and leads to significant fuel savings. Active control of the direction of the thrust vector produces required trim moments, reduces elevon deflections, and also results in significant fuel savings. It is concluded that the combination of active flight control to provide stabilization, (COG) position control to minimize trim moment requirements, and thrust vectoring to generate required trim

170 moments has the potential to signiﬁcantly reduce fuel consumption during ascent to orbit of horizontal takeoff SSTO vehicles. CASI Aerodynamic Drag; Aerodynamic Stability; Air Breathing Engines; Drag Reduction; Flight Control; Fuel Consumption; Horizontal Flight; Hypersonic Vehicles; Single Stage to Orbit Vehicles

19920005095 Nevada Univ., Reno, NV, USA Numerical investigations in three-dimensional internal flows Rose, William C.; JAN 1, 1991; In English Contract(s)/Grant(s): NCC2-507 Report No.(s): NASA-CR-189467; NAS 1.26:189467; No Copyright; Avail: CASI; A03, Hardcopy The general objectives of NASA’s Generic Hypersonics Research Program (GHP) are to develop a technology background required for aeronautical research in the hypersonic Mach number flow range. These research efforts are to complement the National Aerospace Plane (NASP) Program and are geared toward the development of experimental and computational fluid dynamics (CFD) techniques. One of the prominent goals of inlet design for high speed applications is to produce an inlet that delivers uniform flow at its exit in the shortest possible distance. In previous studies, the technologies for determining contours for both the ramp and cowl were demonstrated that allowed a nearly shock free exiting flow field to be obtained. This technology was developed further during the present reporting period and applied to a preliminary design investigation of a biconic hypersonic research vehicle with a nearly 2-D inlet attached near the aft end of the vehicle. The results of a parametric investigation of this proposed inlet for freestream Mach numbers between 10 and 15 are described. CASI Aircraft Design; Engine Inlets; Hypersonic Flow; Hypersonic Vehicles; Hypersonics; Inlet Flow; Mach Number; Research Vehicles; Three Dimensional Flow

19920004848 Georgia Inst. of Tech., Atlanta, GA, USA Rapid near-optimal aerospace plane trajectory generation and guidance Calise, A. J.; Corban, J. E.; Markopoulos, N.; Nov 1, 1991; In English Contract(s)/Grant(s): NAG1-922 Report No.(s): NASA-CR-189469; NAS 1.26:189469; No Copyright; Avail: CASI; A03, Hardcopy Effort was directed toward the problems of the real time trajectory optimization and guidance law development for the National Aerospace Plane (NASP) applications. In particular, singular perturbation methods were used to develop guidance algorithms suitable for onboard, real time implementation. The progress made in this research effort is reported. CASI Aerospace Planes; Aircraft Guidance; Algorithms; National Aerospace Plane Program; Real Time Operation; Trajectory Control; Trajectory Optimization

19920002784 Draper (Charles Stark) Lab., Inc., Cambridge, MA, USA Air-breathing hypersonic vehicle guidance and control studies; An integrated trajectory/control analysis methodology: Phase 1 Hattis, Philip D.; Malchow, Harvey L.; Sep 1, 1991; In English Contract(s)/Grant(s): NAS1-18565; RTOP 505-64-40-01 Report No.(s): NASA-CR-187623; NAS 1.26:187623; No Copyright; Avail: CASI; A05, Hardcopy A tool which generates optimal trajectory/control histories in an integrated manner is generically adapted to the treatment of single-stage-to-orbit air-breathing hypersonic vehicles. The methodology is implemented as a two point boundary value problem solution technique. Its use permits an assessment of an entire near-minimum-fuel trajectory and desired control strategy from takeoff to orbit while satisfying physically derived inequality constraints and while achieving efficient propulsive mode phasing. A simpler analysis strategy that partitions the trajectory into several boundary condition matched segments is also included to construct preliminary trajectory and control history representations with less computational burden than is required for the overall flight profile assessment. A demonstration was accomplished using a tabulated example (winged-cone accelerator) vehicle model that is combined with a newly developed multidimensional cubic spline data smoothing routine. A constrained near-fuel-optimal trajectory, imposing a dynamic pressure limit of 1000 psf, was developed from horizontal takeoff to 20,000 ft/sec relative air speed while aiming for a polar orbit. Previously unspecified propulsive discontinuities were located. Flight regimes demanding rapid attitude changes were identified, dictating control effector and closed-loop controller authority was ascertained after evaluating effector use for vehicle trim. Also, inadequacies in vehicle model representations

171 and speciﬁc subsystem models with insufficient ﬁdelity were determined based on unusual control characteristics and/or excessive sensitivity to uncertainty. CASI Boundary Value Problems; Controllers; Data Smoothing; Feedback Control; Flight Paths; Hypersonic Vehicles; Single Stage to Orbit Vehicles; Trajectories; Trajectory Control

19920002173 Washington Univ., Seattle, WA, USA Thermoviscoplastic response of Ti-15-3 under various loading conditions Tuttle, M. E.; Rogacki, J.; Oct 1, 1991; In English Contract(s)/Grant(s): NAG1-974; RTOP 505-63-50-04 Report No.(s): NASA-CR-187621; NAS 1.26:187621; No Copyright; Avail: CASI; A08, Hardcopy Metal matrix composites (MMC’s) are candidate materials for use in high temperature, high loading applications. In particular, an MMC consisting of a titanium alloy reinforced with silicon-carbide ﬁbers is being considered for use on the National Aerospace Plane (NASP). Compared to other metals and metallic alloys, titanium alloys retain relatively high stiffness, strength, and corrosion resistance at elevated temperatures. However, above roughly 316 C titanium exhibits a signiﬁcant thermoviscoplastic (creep) response. Since the temperatures encountered in many regions of the NASP are expected to exceed 316 C, the potential thermoviscoplastic behavior of titanium-based MMC’s at elevated temperatures must be thoroughly investigated. CASI Creep Properties; Cyclic Loads; Heat Resistant Alloys; Metal Matrix Composites; Spacecraft Construction Materials; Thermoplasticity; Titanium Alloys; Viscoplasticity

19920001816 NASA Lewis Research Center, Cleveland, OH, USA Overview of the Beta II Two-Stage-To-Orbit vehicle design Plencner, Robert M.; Oct 1, 1991; In English; Aircraft Design Systems and Operations Meeting, 23-25 Sep. 1991, Baltimore, MD, USA Contract(s)/Grant(s): RTOP 505-69-40 Report No.(s): NASA-TM-105298; E-6641; NAS 1.15:105298; AIAA PAPER 91-3175; Copyright; Avail: CASI; A03, Hardcopy A study of a near-term, low risk two-stage-to-orbit (TSTO) vehicle was undertaken. The goal of the study was to assess a fully reusable TSTO vehicle with horizontal takeoff and landing capability that could deliver 10,000 pounds to a 120 nm polar orbit. The configuration analysis was based on the Beta vehicle design. A cooperative study was performed to redesign and refine the Beta concept to meet the mission requirements. The vehicle resulting from this study was named Beta II. It has an all-airbreathing first stage and a staging Mach number of 6.5. The second stage is a conventional wing-body configuration with a single SSME. CASI Aerospace Planes; Air Breathing Engines; Launch Vehicle Configurations; Multiengine Vehicles; Multistage Rocket Vehicles; Polar Orbits; Space Shuttle Main Engine

19920000786 Eloret Corp., Sunnyvale, CA, USA Hypersonic flows as related to the National Aerospace plane Kussoy, Marvin; Levy, Lionel; Menter, F.; Mar 19, 1991; In English Contract(s)/Grant(s): NCC2-452 Report No.(s): NASA-CR-187985; NAS 1.26:187985; No Copyright; Avail: CASI; A04, Hardcopy Experimental data for a series of 2-D and 3-D shock wave/boundary layer interaction flows at Mach 8.2 are presented. The test bodies, composed of simple geometric shapes fastened to a flat plate test bed, were designed to generate flows with varying degrees of pressure gradient, boundary layer separation, and turning angle. The data include surface pressure and heat transfer distributions as well as limited mean flowfield surveys both in the undisturbed and interaction regimes. The data are presented in a convenient form to be used to validate existing or future computational models of these hypersonic flows. CASI Aerospace Planes; Boundary Layer Flow; Boundary Layer Separation; Flow Distribution; Heat Transfer; Hypersonic Flow; National Aerospace Plane Program; Shock Wave Interaction

172 19910021224 NASA Lewis Research Center, Cleveland, OH, USA High temperature NASP engine seals: A technology review Steinetz, Bruce M.; Dellacorte, Christopher; Tong, Mike; Aug 1, 1991; In English; 10th National Aerospace Plane Technology Symposium, 23-26 Apr. 1991, Monterey, CA, USA Contract(s)/Grant(s): RTOP 763-21-41 Report No.(s): NASA-TM-104468; E-6317; NAS 1.15:104468; No Copyright; Avail: CASI; A03, Hardcopy Progress in developing advanced high temperature engine seal concepts and related sealing technologies for advanced hypersonic engines are reviewed. Design attributes and issues requiring further development for both the ceramic wafer seal and the braided ceramic rope seal are examined. Leakage data are presented for these seals for engine simulated pressure and temperature conditions and compared to a target leakage limit. Basic elements of leakage ﬂow models to predict leakage rates for each of these seals over the wide range of pressure and temperature conditions anticipated in the engine are also presented. CASI Aerospace Planes; Engine Design; High Temperature; Leakage; Lubrication; Sealing; Seals (Stoppers)

19910020811 Oklahoma Univ., Norman, OK, USA An integrated aerodynamic/propulsion study for generic aero-space planes based on waverider concepts Emanuel, G.; Rasmussen, M. L.; Aug 1, 1991; In English Contract(s)/Grant(s): NAG1-886 Report No.(s): NASA-CR-188691; NAS 1.26:188691; No Copyright; Avail: CASI; A03, Hardcopy Research efforts related to the development of a unified aerospace plane analysis based on waverider technology are summarized. Viscous effects on the forebodies of cone-derived waverider configurations were studied. A simple means for determining the average skin friction coefficient of laminar boundary layers was established. This was incorporated into a computer program that provides lift and drag coefficients and lift/drag ratio for on-design waveriders when the temperature and Reynolds number based on length are specified. An effort was made to carry out parabolized Navier-Stokes (PNS) calculations for cone-derived waveriders. When the viscous terms were turned off (in the Euler mode) computations for elliptic cone-derived waveriders could be carried out for a wide range of on-design and off-design situations. Work related to waveriders derived from power law shocks is described in some detail. M.G. Aerodynamic Characteristics; Aerospace Planes; Spacecraft Design; Spacecraft Propulsion; Waveriders

19910019135 NASA Lewis Research Center, Cleveland, OH, USA Slush hydrogen propellant production, transfer, and expulsion studies at the NASA K-Site Facility Hardy, Terry L.; Whalen, Margaret V.; JAN 1, 1991; In English; Conference on Advanced Space Exploration Initiative Technologies, 4-6 Sep. 1991, Cleveland, OH, USA Contract(s)/Grant(s): RTOP 763-22-21 Report No.(s): NASA-TM-105191; E-6493; NAS 1.15:105191; AIAA PAPER 91-3550; No Copyright; Avail: CASI; A03, Hardcopy Slush hydrogen is currently being considered as a fuel for the National AeroSpace Plane (NASP) because it offers the potential for decreased vehicle size and weight. However, no large scale data was available on the production, transfer, and tank pressure control characteristics required to use the fuel for the NASP. Therefore, experiments were conducted at NASA-Lewis K-Site Facility to improve the slush hydrogen data base. Slush hydrogen was produced using the evaporative cooling, or freeze-thaw, technique in batches for approx. 800 gallons. This slush hydrogen was pressure transferred toa5ft diameter spherical test tank following production, and ﬂow characteristics were measured during this transfer process. The slush hydrogen in the test tank was pressurized and expelled using a pressurized expulsion technique to obtain information on tank pressure control for the NASP. Results from the production, transfer, pressurization, and pressurized expulsion tests are described. CASI Cryogenic Fluid Storage; Flow Characteristics; Hydrogen Production; Propellant Transfer; Slush Hydrogen

19910017819 NASA Langley Research Center, Hampton, VA, USA Structural dynamic and aeroelastic considerations for hypersonic vehicles Cazier, F. W., Jr.; Doggett, Robert V., Jr.; Ricketts, Rodney H.; Jun 1, 1991; In English; AIAA/ASME/ASCE/AHS/ASC 32d Structures, Structural Dynamics, and Materials Conference, 8-10 Apr. 1991, Baltimore, MD, USA Contract(s)/Grant(s): RTOP 763-23-41 Report No.(s): NASA-TM-104110; NAS 1.15:104110; No Copyright; Avail: CASI; A03, Hardcopy

173 The specific geometrical, structural, and operational environment characteristics of hypersonic vehicles are discussed with particular reference to aerospace plane type configurations. A discussion of the structural dynamic and aeroelastic phenomena that must be addressed for this class of vehicles is presented. These phenomena are in the aeroservothermoelasticity technical area. Some illustrative examples of recent experimental and analytical work are given. Some examples of current research are pointed out. CASI Aerodynamic Configurations; Aerospace Planes; Aerothermoelasticity; Hypersonic Vehicles

19910017813 NASA Langley Research Center, Hampton, VA, USA NASP aeroservothermoelasticity studies Doggett, Robert V., Jr.; Ricketts, Rodney H.; Noll, T. E.; Malone, John B.; Apr 1, 1991; In English Contract(s)/Grant(s): RTOP 763-23-41-71 Report No.(s): NASA-TM-104058; NAS 1.15:104058; No Copyright; Avail: CASI; A03, Hardcopy Some illustrative results obtained from work accomplished under the aerothermoelasticity work breakdown structure (WBS) element of the National Aerospace Plane (NASP) Technology Maturation Program (TMP) are presented and discussed. The objectives of the aerothermoelasticity element were to develop analytical methods applicable to aerospace plane type configurations, to conduct analytical studies to identify potential problems, to evaluate potential solutions to problems, and to provide an experimental data base to verify codes and analytical trends. Work accomplished in the three areas of experimental data base, unsteady aerodynamics, and integrated analysis methodology are described. Some of the specific topics discussed are: (1) transonic wind tunnel aeroelastic model tests of cantilever delta wing models, of an all-moveable delta-wing model, and of aileron buzz models; (2) unsteady aerodynamic theory correlation with experiment and theory improvements; and (3) integrated analysis methodology results for thermal effects on vibration, for thermal effects on flutter, and for improving aeroelastic performance by using active controls. CASI Aeroservoelasticity; Aerospace Planes; Aerothermoelasticity; Data Bases; Delta Wings; National Aerospace Plane Program; Servocontrol; Unsteady Aerodynamics; Wind Tunnel Tests

19910011134 NASA Langley Research Center, Hampton, VA, USA Addition of equilibrium air to an upwind Navier-Stokes code and other first steps toward a more generalized flow solver Rosen, Bruce S.; Mar 1, 1991; In English Contract(s)/Grant(s): RTOP 505-80-11-02 Report No.(s): NASA-TM-102616; NAS 1.15:102616; No Copyright; Avail: CASI; A04, Hardcopy An upwind three-dimensional volume Navier-Stokes code is modified to facilitate modeling of complex geometries and flow fields represented by proposed National Aerospace Plane concepts. Code enhancements include an equilibrium air model, a generalized equilibrium gas model and several schemes to simplify treatment of complex geometric configurations. The code is also restructured for inclusion of an arbitrary number of independent and dependent variables. This latter capability is intended for eventual use to incorporate nonequilibrium/chemistry gas models, more sophisticated turbulence and transition models, or other physical phenomena which will require inclusion of additional variables and/or governing equations. Comparisons of computed results with experimental data and results obtained using other methods are presented for code validation purposes. Good correlation is obtained for all of the test cases considered, indicating the success of the current effort. CASI Aerospace Planes; Computational Fluid Dynamics; Flow Distribution; National Aerospace Plane Program; Navier-Stokes Equation

19910010870 Arizona State Univ., Tempe, AZ, USA Dynamics of aerospace vehicles Schmidt, David K.; Apr 1, 1991; In English Contract(s)/Grant(s): NAG1-758 Report No.(s): NASA-CR-188015; NAS 1.26:188015; No Copyright; Avail: CASI; A06, Hardcopy Papers on the following subjects are presented: (1) multivariable ﬂight control synthesis and literal robustness analysis

174 for an aeroelastic vehicles; (2) numerical and literal aeroelastic-vehicle-model reduction for feedback control synthesis; and (3) dynamics of aerospace vehicles. K.S. Aeroelasticity; Aerospace Planes; Aerospace Vehicles; Aircraft Models; Control Systems Design; Dynamic Structural Analysis; Feedback Control; Flexible Spacecraft

19910008803 Boeing Co., Seattle, WA, USA Personnel launch system autoland development study Bossi, J. A.; Langehough, M. A.; Tollefson, J. C.; Jan 1, 1991; In English Contract(s)/Grant(s): NAS1-18762; RTOP 906-11-01-01 Report No.(s): NASA-CR-187495; NAS 1.26:187495; No Copyright; Avail: CASI; A08, Hardcopy The Personnel Launch System (PLS) Autoland Development Study focused on development of the guidance and control system for the approach and landing (A/L) phase and the terminal area energy management (TAEM) phase. In the A/L phase, a straight-in trajectory profile was developed with an initial high glide slope, a pull-up and flare to lower glide slope, and the final flare touchdown. The TAEM system consisted of using a heading alignment cone spiral profile. The PLS autopilot was developed using integral LQG design techniques. The guidance and control design was verified using a nonlinear 6 DOF simulation. Simulation results demonstrated accurate steering during the TAEM phase and adequate autoland performance in the presence of wind turbulence and wind shear. CASI Aerospace Planes; Approach Control; Automatic Landing Control; Automatic Pilots; Control Systems Design; Ferry Spacecraft; Spacecraft Guidance; Spacecraft Landing

19910007821 Loyola Coll., Baltimore, MD, USA JTEC panel report on space and transatmospheric propulsion technology Shelton, Duane; Aug 1, 1990; In English Contract(s)/Grant(s): NSF ECS-89-02528 Report No.(s): NASA-CR-187670; NAS 1.26:187670; PB90-215732; No Copyright; Avail: CASI; A11, Hardcopy An assessment of Japan’s current capabilities in the areas of space and transatmospheric propulsion is presented. The report focuses primarily upon Japan’s programs in liquid rocket propulsion and in propulsion for spaceplanes and related transatmospheric areas. It also includes brief reference to Japan’s solid rocket programs, as well as to supersonic air-breathing propulsion efforts that are just getting underway. The results are based upon the ﬁndings of a panel of U.S. engineers made up of individuals from academia, government, and industry, and are derived from a review of a broad array of the open literature, combined with visits to the primary propulsion laboratories and development agencies in Japan. NTIS Aerospace Planes; Japan; Jet Propulsion; Liquid Propellant Rocket Engines; Spacecraft Propulsion; Technology Assessment; Transatmospheric Vehicles

19910007799 NASA Ames Research Center, Moffett Field, CA, USA The aerospace plane design challenge: Credible computational fluid dynamics results Mehta, Unmeel B.; Dec 1, 1990; In English; 2nd AIAA International Aerospace Planes Conference, 29-31 Oct. 1990, Orlanda, FL, USA Contract(s)/Grant(s): RTOP 505-60-00 Report No.(s): NASA-TM-102887; A-91016; NAS 1.15:102887; No Copyright; Avail: CASI; A03, Hardcopy Computational fluid dynamics (CFD) is necessary in the design processes of all current aerospace plane programs. Single-stage-to-orbit (STTO) aerospace planes with air-breathing supersonic combustion are going to be largely designed by means of CFD. The challenge of the aerospace plane design is to provide credible CFD results to work from, to assess the risk associated with the use of those results, and to certify CFD codes that produce credible results. To establish the credibility of CFD results used in design, the following topics are discussed: CFD validation vis-a-vis measurable fluid dynamics (MFD) validation; responsibility for credibility; credibility requirement; and a guide for establishing credibility. Quantification of CFD uncertainties helps to assess success risk and safety risks, and the development of CFD as a design tool requires code certification. This challenge is managed by designing the designers to use CFD effectively, by ensuring quality control, and by balancing the design process. For designing the designers, the following topics are discussed: how CFD design technology is developed; the reasons Japanese companies, by and large, produce goods of higher quality than the U.S. counterparts;

175 teamwork as a new way of doing business; and how ideas, quality, and teaming can be brought together. Quality control for reducing the loss imparted to the society begins with the quality of the CFD results used in the design process, and balancing the design process means using a judicious balance of CFD and MFD. CASI Aerospace Planes; Computational Fluid Dynamics; Design Analysis; National Aerospace Plane Program; Supersonic Combustion

19910006700 Rice Univ., Houston, TX, USA Optimal trajectories for an aerospace plane. Part 1: Formulation, results, and analysis Miele, Angelo; Lee, W. Y.; Wu, G. D.; JAN 1, 1990; In English; American Control Conference, 23-25 May 1990, San Diego, CA, USA Contract(s)/Grant(s): TATP-003604020; NAG1-1029 Report No.(s): NASA-CR-187868; NAS 1.26:187868; AAR-247; No Copyright; Avail: CASI; A04, Hardcopy The optimization of the trajectories of an aerospace plane is discussed. This is a hypervelocity vehicle capable of achieving orbital speed, while taking off horizontally. The vehicle is propelled by four types of engines: turbojet engines for flight at subsonic speeds/low supersonic speeds; ramjet engines for flight at moderate supersonic speeds/low hypersonic speeds; scramjet engines for flight at hypersonic speeds; and rocket engines for flight at near-orbital speeds. A single-stage-to-orbit (SSTO) configuration is considered, and the transition from low supersonic speeds to orbital speeds is studied under the following assumptions: the turbojet portion of the trajectory has been completed; the aerospace plane is controlled via the angle of attack and the power setting; the aerodynamic model is the generic hypersonic aerodynamics model example (GHAME). Concerning the engine model, three options are considered: (EM1), a ramjet/scramjet combination in which the scramjet specific impulse tends to a nearly-constant value at large Mach numbers; (EM2), a ramjet/scramjet combination in which the scramjet specific impulse decreases monotonically at large Mach numbers; and (EM3), a ramjet/scramjet/rocket combination in which, owing to stagnation temperature limitations, the scramjet operates only at M approx. less than 15; at higher Mach numbers, the scramjet is shut off and the aerospace plane is driven only by the rocket engines. Under the above assumptions, four optimization problems are solved using the sequential gradient-restoration algorithm for optimal control problems: (P1) minimization of the weight of fuel consumed; (P2) minimization of the peak dynamic pressure; (P3) minimization of the peak heating rate; and (P4) minimization of the peak tangential acceleration. CASI Aerodynamic Characteristics; Aerospace Planes; Hypersonic Speed; Hypersonics; Optimization; Supersonic Speed; Trajectories

19910006698 Rice Univ., Houston, TX, USA Optimal trajectories for an aerospace plane. Part 2: Data, tables, and graphs Miele, Angelo; Lee, W. Y.; Wu, G. D.; JAN 1, 1990; In English; 1990 American Control Conference, 23-25 May 1990, San Diego, CA, USA Contract(s)/Grant(s): NAG1-1029 Report No.(s): NASA-CR-187848; NAS 1.26:187848; AAR-248-PT-2; No Copyright; Avail: CASI; A05, Hardcopy Data, tables, and graphs relative to the optimal trajectories for an aerospace plane are presented. A single-stage-to-orbit (SSTO) conﬁguration is considered, and the transition from low supersonic speeds to orbital speeds is studied for a single aerodynamic model (GHAME) and three engine models. Four optimization problems are solved using the sequential gradient-restoration algorithm for optimal control problems: (1) minimization of the weight of fuel consumed; (2) minimization of the peak dynamic pressure; (3) minimization of the peak heating rate; and (4) minimization of the peak tangential acceleration. The above optimization studies are carried out for different combinations of constraints, speciﬁcally: initial path inclination that is either free or given; dynamic pressure that is either free or bounded; and tangential acceleration that is either free or bounded. CASI Aerospace Planes; Algorithms; Engine Design; Hypersonic Flight; Mathematical Models; Optimal Control; Supersonic Speed; Trajectory Optimization

176 19900016844 NASA Lewis Research Center, Cleveland, OH, USA Prediction of the ullage gas thermal stratification in a NASP vehicle propellant tank experimental simulation using FLOW-3D Hardy, Terry L.; Tomsik, Thomas M.; Jul 1, 1990; In English Contract(s)/Grant(s): RTOP 763-01-21 Report No.(s): NASA-TM-103217; E-5629; NAS 1.15:103217; No Copyright; Avail: CASI; A03, Hardcopy As part of the National Aero-Space Plane (NASP) project, the multi-dimensional effects of gravitational force, initial tank pressure, initial ullage temperature, and heat transfer rate on the 2-D temperature profiles were studied. FLOW-3D, a commercial finite difference fluid flow model, was used for the evaluation. These effects were examined on the basis of previous liquid hydrogen experimental data with gaseous hydrogen pressurant. FLOW-3D results were compared against an existing 1-D model. In addition, the effects of mesh size and convergence criteria on the analytical results were investigated. Suggestions for future modifications and uses of FLOW-3D for modeling of a NASP tank are also presented. CASI Aerospace Planes; Computer Programs; Fluid Flow; Heat Transfer; National Aerospace Plane Program; Propellant Tanks; Ullage

19900016739 NASA Lewis Research Center, Cleveland, OH, USA Background, current status, and prognosis of the ongoing slush hydrogen technology development program for the NASP Dewitt, R. L.; Hardy, T. L.; Whalen, M. V.; Richter, G. P.; Tomsik, T. M.; JAN 1, 1990; In English; 8th World Hydrogen Energy Conference, 22-27 Jul. 1990, Honolulu, HI, USA Contract(s)/Grant(s): RTOP 763-01-21 Report No.(s): NASA-TM-103220; E-5634; NAS 1.15:103220; No Copyright; Avail: CASI; A03, Hardcopy Among the Hydrogen Projects at the NASA Lewis Research Center (NASA LeRC), is the task of implementing and managing the Slush Hydrogen (SLH2) Technology Program for the USA’ National AeroSpace Plane Joint Program Office (NASP JPO). The objectives of this NASA LeRC program are to provide verified numerical models of fluid production, storage, transfer, and feed systems and to provide verified design criteria for other engineered aspects of SLH2 systems germane to a NASP. The pursuit of these objectives is multidimensional, covers a range of problem areas, works these to different levels of depth, and takes advantage of the resources available in private industry, academia, and the U.S. Government. A summary of the NASA LeRC overall SLH2 program plan, is presented along with its implementation, the present level of effort in each of the program areas, some of the results already in hand, and the prognosis for the effort in the immediate future. CASI Aerospace Planes; Design Analysis; Mathematical Models; National Aerospace Plane Program; Project Planning; Slush Hydrogen

19900009959 Georgia Inst. of Tech., Atlanta, GA, USA Rapid near-optimal trajectory generation and guidance law development for single-stage-to-orbit airbreathing vehicles Calise, A. J.; Flandro, G. A.; Corban, J. E.; Jan 1, 1990; In English Contract(s)/Grant(s): NAG1-922 Report No.(s): NASA-CR-186299; NAS 1.26:186299; No Copyright; Avail: CASI; A04, Hardcopy General problems associated with on-board trajectory optimization, propulsion system cycle selection, and with the synthesis of guidance laws were addressed for an ascent to low-earth-orbit of an air-breathing single-stage-to-orbit vehicle. The NASA Generic Hypersonic Aerodynamic Model Example and the Langley Accelerator aerodynamic sets were acquired and implemented. Work related to the development of purely analytic aerodynamic models was also performed at a low level. A generic model of a multi-mode propulsion system was developed that includes turbojet, ramjet, scramjet, and rocket engine cycles. Provisions were made in the dynamic model for a component of thrust normal to the flight path. Computational results, which characterize the nonlinear sensitivity of scramjet performance to changes in vehicle angle of attack, were obtained and incorporated into the engine model. Additional trajectory constraints were introduced: maximum dynamic pressure; maximum aerodynamic heating rate per unit area; angle of attack and lift limits; and limits on acceleration both along and normal to the flight path. The remainder of the effort focused on required modifications to a previously derived algorithm when the model complexity cited above was added. In particular, analytic switching conditions were derived which, under appropriate assumptions, govern optimal transition from one propulsion mode to another for two cases: the case in which engine cycle operations can overlap, and the case in which engine cycle operations are mutually exclusive. The resulting guidance

177 algorithm was implemented in software and exercised extensively. It was found that the approximations associated with the assumed time scale separation employed in this work are reasonable except over the Mach range from roughly 5 to 8. This phenomenon is due to the very large thrust capability of scramjets in this Mach regime when sized to meet the requirement for ascent to orbit. By accounting for flight path angle and flight path angle rate in construction of the flight path over this Mach range, the resulting algorithm provides the means for rapid near-optimal trajectory generation and propulsion cycle selection over the entire Mach range from take-off to orbit. CASI Aerospace Planes; Hypersonic Vehicles; Propulsion System Performance; Single Stage to Orbit Vehicles; Spacecraft Guidance; Spacecraft Propulsion; Spacecraft Trajectories; Trajectory Optimization

19900009436 NASA Langley Research Center, Hampton, VA, USA Structures and materials technology for hypersonic aerospacecraft Mccomb, Harvey G., Jr.; Murrow, Harold N.; Card, Michael F.; Jan 1, 1990; In English Contract(s)/Grant(s): RTOP 506-43-71-04 Report No.(s): NASA-TM-102583; NAS 1.15:102583; No Copyright; Avail: CASI; A03, Hardcopy Major considerations in structural design of a transatmospheric aerospacecraft are discussed. The general direction of progress in structures and materials technology is indicated, and technical areas in structures and materials where further research and development is necessary are indicated. Various structural concepts under study and materials which appear to be most applicable are discussed. Structural design criteria are discussed with particular attention to the factor-of-safety approach and the probabilistic approach. Structural certification requirements for the aerospacecraft are discussed. The kinds of analyses and tests which would be required to certify the structural integrity, safety, and durability of the aerospacecraft are discussed, and the type of test facility needed to perform structural certification tests is identified. CASI Aerospace Planes; Aerospace Vehicles; Certification; Hypersonics; Structural Design; Structural Design Criteria; Structural Engineering

19900007739 McDonnell Aircraft Co., Saint Louis, MO, USA Heat pipes for wing leading edges of hypersonic vehicles Boman, B. L.; Citrin, K. M.; Garner, E. C.; Stone, J. E.; Jan 1, 1990; In English Contract(s)/Grant(s): NAS1-18144; RTOP 506-49-11-05 Report No.(s): NASA-CR-181922; NAS 1.26:181922; No Copyright; Avail: CASI; A06, Hardcopy Wing leading edge heat pipes were conceptually designed for three types of vehicle: an entry research vehicle, aero-space plane, and advanced shuttle. A full scale, internally instrumented sodium/Hastelloy X heat pipe was successfully designed and fabricated for the advanced shuttle application. The 69.4 inch long heat pipe reduces peak leading edge temperatures from 3500 F to 1800 F. It is internally instrumented with thermocouples and pressure transducers to measure sodium vapor qualities. Large thermal gradients and consequently large thermal stresses, which have the potential of limiting heat pipe life, were predicted to occur during startup. A test stand and test plan were developed for subsequent testing of this heat pipe. Heat pipe manufacturing technology was advanced during this program, including the development of an innovative technique for wick installation. CASI Fabrication; Heat Pipes; Hypersonic Vehicles; Leading Edges; Wings

178 Subject Term Index

ABLATIVE MATERIALS The Reusable Launch Vehicle Technol- Structural dynamic and aeroelastic con- Development of Processing Techniques ogy Program and the X-33 Advanced siderations for hypersonic ve- for Advanced Thermal Protection Materi- Technology Demonstrator – 128 hicles – 173 als – 112 AERODYNAMIC CHARACTERISTICS X-33 (Rev-F) Aeroheating Results of Test Materials Testing on the DC-X and DC- A method of predicting quasi-steady 6770 in NASA Langley 20-Inch Mach 6 XA – 121 aerodynamics for flutter analysis of high Air Tunnel – 96 Task 4 supporting technology. Part 1: speed vehicles using steady CFD calcu- AERODYNAMIC DRAG lations – 157 Detailed test plan for leading edge tile A Base Drag Reduction Experiment on development. Leading edge material de- Abort performance for a winged-body the X-33 Linear Aerospike SR-71 Experi- velopment and testing – 139 single-stage to orbit vehicle – 137 ment (LASRE) Flight Program – 97 Thermal Protection Materials for Reentry Aerodynamic Characteristics of a Vertical A simple analytical aerodynamic model Applications – 59 Takeoff Vertical Landing (VTVL) Single of Langley Winged-Cone Aerospace Stage to Orbit Vehicle from M(sub infin- 145 ABORT TRAJECTORIES Plane concept – ity) = 0.1 to 10 – 122 Abort performance for a winged-body Computation of H2/air reacting flowfields single-stage to orbit vehicle – 137 An integrated aerodynamic/propulsion in drag-reduction external combus- study for generic aero-space planes tion – 161 ACCELERATORS based on waverider concepts – 173 NASA’s Advanced Space Transportation Reduction of Base Drag on Launch Ve- Computational Study of a McDonnell 50 Program: RTA Project Summary – 42 hicles – Douglas Single-Stage-to-Orbit Vehicle Trim drag reduction concepts for horizon- Concept for Aerodynamic Analy- ACCIDENT INVESTIGATION tal takeoff single-stage-to-Orbit ve- sis – 126 X-33 LH2 Tank Failure Investigation hicles – 170 Findings – 74 Design and Analysis of Turbines for Wind-Tunnel Investigations of Blunt- Space Applications – 16 ACOUSTIC EMISSION Body Drag Reduction Using Forebody Acoustic Emission Monitoring of the Development of the X-33 Aerodynamic Surface Roughness – 85 DC-XA Composite Liquid Hydrogen Tank Uncertainty Model – 108 AERODYNAMIC FORCES During Structural Testing – 126 Experimental Performance Evaluation of A method of predicting quasi-steady Study Acoustic Emissions from Compos- a Supersonic Turbine for Rocket Engine aerodynamics for flutter analysis of high ites – 111 Applications – 10 speed vehicles using steady CFD calcu- ACOUSTIC FATIGUE Flight Testing the Linear Aerospike lations – 157 SR-71 Experiment (LASRE) – 100 Development of Metallic Thermal Protec- Transonic aerodynamic characteristics of tion Systems for the Reusable Launch Hypersonic vehicle model and control a proposed wing-body reusable launch Vehicle – 124 law development using H(infinity) and vehicle concept – 138 micron synthesis – 144 Panel Flutter and Sonic Fatigue Analysis AERODYNAMIC HEAT TRANSFER for RLV – 77 Optimal trajectories for an aerospace Improvements in Thermal Protection Siz- plane. Part 1: Formulation, results, and ACTIVE CONTROL ing Capabilities for TCAT: Conceptual analysis – 176 Design for Advanced Space Transporta- The 2002 NASA Faculty Fellowship Pro- tion Systems – 61 gram Research Reports – 9 RLV Turbine Performance Optimiza- tion – 83 AERODYNAMIC HEATING ACTUATORS Subsonic aerodynamic characteristics of A Method of Integrating Aeroheating into Direct Fault Tolerant RLV Altitude Con- a proposed advanced manned launch Conceptual Reusable Launch Vehicle trol: A Singular Perturbation Ap- system orbiter configuration – 160 Design: Evaluation of Advanced Thermal proach – 40 Subsonic-to-Hypersonic Aerodynamic Protection Techniques for Future Reus- 77 Using Neural Networks in Decision Mak- Characteristics for a Winged, Circular- able Launch Vehicles – ing for a Reconfigurable Electro Me- Body, Single-Stage-to-Orbit Spacecraft Aerothermoelastic analysis of a NASP chanical Actuator (EMA) – 68 Configuration – 124 demonstrator model – 155 ADHESION TESTS Transonic aerodynamic characteristics of An overview of aeroelasticity studies for Bonding and Sealing Evaluations for a proposed wing-body reusable launch the National Aerospace Plane – 160 Cryogenic Tanks – 116 vehicle concept – 138 Computational/Experimental Aeroheat- ADSORPTION AERODYNAMIC COEFFICIENTS ing Predictions for X-33 – 112 An Investigation of the Effect of Surface A simple analytical aerodynamic model Development of Metallic Thermal Protec- Impurities on the Adsorption Kinetics of of Langley Winged-Cone Aerospace tion Systems for the Reusable Launch Hydrogen Chemisorbed onto Iron – 101 Plane concept – 145 Vehicle – 124 ADVANCED LAUNCH SYSTEM (STS) AERODYNAMIC CONFIGURATIONS Evaluation of Advanced Thermal Protec- Advanced transportation system studies. A Geometry-Centered Approach For tion Techniques for Future Reusable Technical area 2: Heavy lift launch ve- Two-Stage-To-Orbit Vehicle Synthe- Launch Vehicles – 62 hicle development – 129 sis – 45 Evaluation of Thermal Control Coatings Operationally efficient propulsion system Full Envelope Reconfigurable Control for Flexible Ceramic Thermal Protection study (OEPSS) data book – 110 Design for the X-33 Vehicle – 72 Systems – 117

A-1 Infrared Sensing Aeroheating Flight Ex- AEROSPACE ENGINEERING An integrated aerodynamic/propulsion periment: STS-96 Flight Results – 86 2000 NASA Seal/Secondary Air System study for generic aero-space planes 173 X-33 Computational Workshop – 62 based on waverider concepts – Aeroheating/Aerodynamic Predictions A Hazardous Gas Detection System for An inverse dynamics approach to trajec- and Comparisons With Experimental Aerospace and Commercial Applica- tory optimization for an aerospace Data – 25 tions – 98 plane – 163 X-33 (Rev-F) Aeroheating Results of Test Design and Testing of Non-Toxic RCS Analytical solutions to constrained hyper- 6770 in NASA Langley 20-Inch Mach 6 Thrusters for Second Generation Reus- sonic flight trajectories – 162 Air Tunnel – 96 able Launch Vehicle – 14 Assessment of flying-quality criteria for X-33 Rev-F Turbulent Aeroheating Re- Development and Capabilities of Unique air-breathing aerospacecraft – 167 sults From Test 6817 in NASA Langley Structural Seal Test Rigs – 41 20-Inch Mach 6 Air Tunnel and Compari- Background, current status, and progno- General Public Space Travel and Tour- sons With Computations – 26 sis of the ongoing slush hydrogen tech- ism – 96 nology development program for the X-33 Turbulent Aeroheating Measure- NASA’s Next Generation Launch Tech- NASP – 177 ments and Predictions – 32 nology Program - Strategy and Conceptual design of two-stage-to-orbit AERODYNAMIC LOADS 4 Plans – hybrid launch vehicle – 169 A study of facilities and fixtures for testing Space Launch Initiative: New Capabili- of a high speed civil transport wing com- ties ... New Horizons – 36 Constrained minimization of smooth ponent – 127 functions using a genetic algo- Technology and Risk Assessment Using rithm – 141 AERODYNAMIC STABILITY a Robust Design Simulation Methodol- Aerothermoelastic analysis of a NASP ogy – 121 Current Technology for Thermal Protec- demonstrator model – 155 tion Systems – 165 The Center for Aerospace Research: A Flight Stability and Control and Perfor- NASA Center of Excellence at North Dynamics of aerospace vehicles – 174 mance Results from the Linear Aerospike Carolina Agricultural and Technical State SR-71 Experiment (LASRE) – 101 High temperature NASP engine seal de- University – 162 velopment – 153 Transonic aerodynamic characteristics of The F-18 systems research aircraft facil- High temperature NASP engine seals: A a proposed wing-body reusable launch ity – 163 vehicle concept – 138 technology review – 173 AEROSPACE INDUSTRY Trim drag reduction concepts for horizon- Hot dynamic test rig for measuring hyper- Analysis of space concepts enabled by tal takeoff single-stage-to-Orbit ve- sonic engine seal flow and durabil- new transportation (ASCENT) hicles – 170 ity – 151 study – 28 AERODYNAMICS Hypersonic flows as related to the Na- Improved Re-Configurable Sliding Mode AEROSPACE PLANES tional Aerospace Plane – 136 Controller for Reusable Launch Vehicle A conceptual design of an unmanned test vehicle using an airbreathing propulsion Hypersonic flows as related to the na- of Second Generation Addressing Aero- tional aerospace plane – 162 dynamic Surface Failures and Thrust De- system – 154 ficiencies – 34 A method of predicting quasi-steady Hypersonic flows as related to the Na- tional Aerospace plane – 172 Preliminary Sizing of Vertical Take-Off aerodynamics for flutter analysis of high Rocket-Based Combined-Cycle Pow- speed vehicles using steady CFD calcu- JTEC panel report on space and transat- ered Launch Vehicles – 85 lations – 157 mospheric propulsion technology – 175 A simple analytical aerodynamic model AEROELASTICITY Lockheed Martin Response to the OSP of Langley Winged-Cone Aerospace An overview of aeroelasticity studies for Challenge – 20 Plane concept – 145 the National Aerospace Plane – 160 Materials Research for High Speed Civil A summary of the slush hydrogen tech- Dynamics of aerospace vehicles – 174 Transport and Generic Hypersonics- nology program for the National Aero- Metals Durability – 127 Evaluation of Computational Aeroelastic- Space Plane – 140 ity Codes for Loads and Flutter – 82 A Thermal Management Systems Model NASA aeronautics: Research and tech- Supersonic aeroelastic instability results for the NASA GTX RBCC Concept – 53 nology program highlights – 164 for a NASP-like wing model – 159 A Two-Stage-to-Orbit Spaceplane Con- NASA’s New Orbital Space Plane: A AERONAUTICAL ENGINEERING cept With Growth Potential – 80 Bridge to the Future – 19 Aeronautical Engineering: A Continuing Access to space – 143 NASA’s Next Generation Launch Tech- Bibliography with Indexes – 97 nology Program - Strategy and AERONAUTICS Addition of equilibrium air to an upwind Plans – 4 Navier-Stokes code and other first steps NASA aeronautics: Research and tech- toward a more generalized flow NASA’s Orbital Space Plane Risk Re- nology program highlights – 164 solver – 174 duction Strategy – 13 Winds of change: Expanding the frontiers Advanced Range Safety System for High NASP aeroservothermoelasticity stud- of flight. Langley Research Center’s 75 Energy Vehicles – 15 ies – 174 years of accomplishment, 1917- 1992 – 165 Aero-Space Plane: Flexible access to National Aerospace Plane Thermal De- AEROSERVOELASTICITY space – 150 velopment. (Latest Citations from the Aerospace Database) – 113 Advanced aeroservoelastic stabilization Aerothermoelastic analysis of a NASP techniques for hypersonic flight ve- demonstrator model – 155 National Aerospace Plane Thermal De- hicles – 163 Air-breathing aerospace plane develop- velopment. (Latest citations from the 114 NASP aeroservothermoelasticity stud- ment essential: Hypersonic propulsion Aerospace Database) – ies – 174 flight tests – 138 National aerospace plane – 151

A-2 Optimal trajectories for an aerospace The 1995 NASA High-Speed Research Opportunities for Breakthroughs in plane. Part 1: Formulation, results, and Program Sonic Boom Workshop – 126 Large-Scale Computational Simulation 176 54 analysis – The aerospace plane design challenge: and Design – Optimal trajectories for an aerospace Credible computational fluid dynamics Structures and materials technology for plane. Part 2: Data, tables, and results – 176 hypersonic aerospacecraft – 178 graphs – 176 The Center for Aerospace Research: A AEROSPIKE ENGINES Orbital Space Plane Cost Credibility – 4 NASA Center of Excellence at North An Inviscid Computational Study of an Orbital Space Plane (OSP) Program at Carolina Agricultural and Technical State X-33 Configuration at Hypersonic Lockheed Martin – 7 University – 162 Speeds – 94 The future challenge for aeropropul- Orbital Space Plane (OSP) Pro- Analysis of Aerospike Plume Induced sion – 167 gram – 5 Base-Heating Environment – 109 Thrust augmentation options for the Beta Orbital Space Plane Program Flight Base-Bleed Effect on X-33 Aerospike 2 two-stage-to-orbit vehicle – 154 Demonstrators Status – 6 Plume Induced Base-Heating Environ- Trajectory optimization for the national ment During Power-Pack Out – 25 Orbital Space Plane Program Sta- aerospace plane – tus – 5 156 Development of Aluminum Composites for a Rocket Engine’s Lightweight Thrust Overview of the Beta II Two-Stage-To- Trajectory optimization for the National 160 Cell – 80 Orbit vehicle design – 172 aerospace plane – Turbulence modeling for hypersonic Evaluation of the Linear Aerospike SR-71 Personnel launch system autoland de- flight – 161 Experiment (LASRE) Oxygen Sen- velopment study – 175 sor – 91 AEROSPACE SAFETY Perspective on the National Aero-Space Flight Stability and Control and Perfor- A Hazardous Gas Detection System for Plane Program instrumentation develop- mance Results from the Linear Aerospike Aerospace and Commercial Applica- ment – 156 SR-71 Experiment (LASRE) – 101 tions – 98 PNS predictions of external/internal hy- Flight Testing the Linear Aerospike Design and Testing of Non-Toxic RCS personic flows for NASP propulsion ap- SR-71 Experiment (LASRE) – 100 plications – 157 Thrusters for Second Generation Reus- able Launch Vehicle – 18 Linear Aerospike SR-71 Experiment Prediction of the ullage gas thermal (LASRE): Aerospace Propulsion Hazard stratification in a NASP vehicle propellant AEROSPACE SCIENCES Mitigation Systems – 102 tank experimental simulation using Research Reports: 1996 NASA/ASEE FLOW-3D – 177 Summer Faculty Fellowship Pro- Propellant Feed System Leak Detection: gram – 102 Lessons Learned From the Linear Aero- Rapid near-optimal aerospace plane tra- spike SR-71 Experiment (LASRE) – 93 jectory generation and guidance – 171 Winds of change: Expanding the frontiers of flight. Langley Research Center’s 75 Test Report for NASA MSFC Support of Rapid near-optimal trajectory generation the Linear Aerospike SR-71 Experiment and guidance law development for years of accomplishment, 1917- 165 (LASRE) – 89 single-stage-to-orbit airbreathing ve- 1992 – hicles – 177 AEROSPACE SYSTEMS The Control System for the X-33 Linear Aerospike Engine – 105 Real-time trajectory optimization on par- Fuel Cell Activities at the NASA Glenn allel processors – 158 Research Center – 49 X-33 XRS-2200 Linear Aerospike Engine Sea Level Plume Radiation – 66 Robust intelligent flight control for hyper- Integrated Technology Assessment Cen- sonic vehicles – 166 ter (ITAC) Update – 42 AEROTHERMODYNAMICS Rocket Based Combined Cycle (RBCC) Liquid Engine Test Facilities Assess- A Thermal Management Systems Model Propulsion Technology Workshop. Vol- ment – 1 for the NASA GTX RBCC Concept – 53 ume 1: Executive summary – 166 Progress and Status of the 2nd Genera- Statistical Methods for Rapid Aerother- Rocket-Based Combined-Cycle (RBCC) tion Reusable Launch Vehicle Pro- mal Analysis and Design Technol- Propulsion Technology Workshop. Tuto- gram – 76 ogy – 55 rial session – 168 Space Transportation Systems Tech- Thermal Protection System Evaluation Slush hydrogen transfer studies at the nologies – 70 Using Arc-jet Flows: Flight Simulation or 167 Research Tool? – 31 NASA K-Site Test Facility – System Risk Assessment and Allocation Stratospheric turbulence measurements in Conceptual Design – 27 X-33 (Rev-F) Aeroheating Results of Test and models for aerospace plane de- 6770 in NASA Langley 20-Inch Mach 6 sign – 164 AEROSPACE VEHICLES Air Tunnel – 96 Analytical solutions to constrained hyper- AEROTHERMOELASTICITY Structural dynamic and aeroelastic con- sonic flight trajectories – 162 siderations for hypersonic ve- Aerothermoelastic analysis of a NASP hicles – 173 Development of Structural Health Man- demonstrator model – 155 agement Technology for Aerospace Ve- Structures and materials technology for hicles – 1 NASP aeroservothermoelasticity stud- hypersonic aerospacecraft – 178 ies – 174 Dynamics of aerospace vehicles – 174 Supersonic aeroelastic instability results Structural dynamic and aeroelastic con- for a NASP-like wing model – 159 Flight Demonstrations of Orbital Space siderations for hypersonic ve- Plane (OSP) Technologies – 16 Technology and Risk Assessment Using hicles – 173 a Robust Design Simulation Methodol- Mission and sizing analysis for the Beta 2 AFTERBODIES ogy – 121 two-stage-to-orbit vehicle – 168 Aerodynamic Characteristics of a Vertical Technology issues associated with using Multidisciplinary design techniques ap- Takeoff Vertical Landing (VTVL) Single densified hydrogen for space ve- plied to conceptual aerospace vehicle Stage to Orbit Vehicle from M(sub infin- hicles – 167 design – 156 ity) = 0.1 to 10 – 122

A-3 AIR BREATHING BOOSTERS AIR FLOW Rapid near-optimal aerospace plane tra- Mission and sizing analysis for the Beta 2 An implementation of a chemical and jectory generation and guidance – 171 two-stage-to-orbit vehicle – 168 thermal nonequilibrium flow solver on AIRCRAFT HAZARDS unstructured meshes and application to Lee waves: Benign and malig- NASA Glenn Research Center’s Hyper- blunt bodies – 145 sonic Propulsion Program – 95 nant – 157 AIR LAUNCHING AIRCRAFT INSTRUMENTS Structural Sizing of a 25,000-lb Payload, The design and evolution of the beta NASA Stennis Space Center V and V Air-Breathing Launch Vehicle for Single- two-stage-to-orbit horizontal takeoff and Capabilities Overview – 2 Stage-to-Orbit – 85 landing launch system – 159 Thrust augmentation options for the Beta AIRCRAFT CONFIGURATIONS AIRCRAFT LANDING 2 two-stage-to-orbit vehicle – 154 Configuration development study of the Adaptive Critic Neural Network-Based OSU 1 hypersonic research ve- Terminal Area Energy Management and 9 AIR BREATHING ENGINES hicle – 154 Approach and Landing Guidance – A conceptual design of an unmanned test AIRCRAFT MODELS AIRCRAFT CONSTRUCTION MATERIALS vehicle using an airbreathing propulsion Dynamics of aerospace vehicles – 174 system – 154 Development of Processing Techniques for Advanced Thermal Protection Materi- National Aero-Space Plane resource A Two-Stage-to-Orbit Spaceplane Con- als – 112 reel – 155 80 cept With Growth Potential – National aerospace plane – 151 National Aero-Space Plane – 155 Affordable Flight Demonstration of the Thermal Management Design for the Wingless Flight: The Lifting Body GTX Air-Breathing SSTO Vehicle Con- X-33 Lifting Body – 104 Story – 105 cept – 26 AIRCRAFT CONTROL AIRCRAFT NOISE Air-breathing hypersonic vehicle guid- Advanced aeroservoelastic stabilization The future challenge for aeropropul- ance and control studies: An integrated techniques for hypersonic flight ve- sion – 167 trajectory/control analysis methodology, hicles – 163 phase 2 – 165 AIRCRAFT PERFORMANCE AIRCRAFT DESIGN Flight Stability and Control and Perfor- Airbreathing/Rocket Single-Stage-to- A conceptual design of an unmanned test mance Results from the Linear Aerospike Orbit Design Matrix – 124 vehicle using an airbreathing propulsion SR-71 Experiment (LASRE) – 101 Assessment of flying-quality criteria for system – 154 AIRCRAFT RELIABILITY air-breathing aerospacecraft – 167 Affordable Flight Demonstration of the Advanced Durability and Damage Toler- Design and Analysis of a Two-Stage-to- GTX Air-Breathing SSTO Vehicle Con- ance Design and Analysis Methods for Orbit Airbreathing Hypersonic Vehicle cept – 26 Composite Structures: Lessons Learned Concept – 129 Full Envelope Reconfigurable Control from NASA Technology Development Programs – 26 Development and Capabilities of Unique Design for the X-33 Vehicle – 72 Structural Seal Test Rigs – 41 Mission and sizing analysis for the Beta 2 AIRCRAFT two-stage-to-orbit vehicle – 168 Winds of change: Expanding the frontiers GTX Reference Vehicle Structural Verifi- of flight. Langley Research Center’s 75 cation Methods and Weight Sum- NASA Overview – 35 years of accomplishment, 1917- mary – 44 Neural Network and Regression Approxi- 1992 – 165 mations in High Speed Civil Transport NASA Glenn Research Center’s Hyper- AIRFRAMES sonic Propulsion Program – 95 Aircraft Design Optimization – 106 A parametric sensitivity study for single- Overview of the Beta II Two-Stage-To- Numerical investigations in three- stage-to-orbit hypersonic vehicles using Orbit vehicle design – 172 dimensional internal flows – 171 trajectory optimization – 153 System Identification of X-33 Neural Net- Performance Evaluation of the NASA An overview of aeroelasticity studies for work – 175 GTX RBCC Flowpath – 70 the National Aerospace Plane – 160 Technology and Risk Assessment Using Robust intelligent flight control for hyper- Thermal Management Design for the a Robust Design Simulation Methodol- X-33 Lifting Body – 104 sonic vehicles – 166 ogy – 121 AIRLINE OPERATIONS Rocket Based Combined Cycle (RBCC) The F-18 systems research aircraft facil- Propulsion Technology Workshop. Vol- ity – 163 Aircraft Emission Scenarios Projected in ume 1: Executive summary – 166 Year 2015 for the NASA Technology Con- AIRCRAFT ENGINES cept Aircraft (TCA) High Speed Civil Rocket Based Combined Cycle (RBCC) Fuel Cell Activities at the NASA Glenn Transport – 108 Propulsion Workshop, volume 2 – 166 Research Center – 49 ALGORITHMS Systems Challenges for Hypersonic Ve- The future challenge for aeropropul- An Improved RLV Stability Analysis Via a hicles – 117 sion – 167 Continuation Approach – 9 The design and evolution of the beta AIRCRAFT FUELS Evaluation of Advanced Thermal Protec- two-stage-to-orbit horizontal takeoff and Aircraft Emission Scenarios Projected in tion Techniques for Future Reusable landing launch system – 159 Year 2015 for the NASA Technology Con- Launch Vehicles – 62 Trim drag reduction concepts for horizon- cept Aircraft (TCA) High Speed Civil Flush Airdata Sensing (FADS) System tal takeoff single-stage-to-Orbit ve- Transport – 108 Calibration Procedures and Results for hicles – 170 AIRCRAFT GUIDANCE Blunt Forebodies – 93 Update on the Development and Capa- Adaptive Critic Neural Network-Based Integrated G and C Implementation bilities of Unique Structural Seal Test Terminal Area Energy Management and within IDOS: A Simulink Based Reusable Rigs – 3 Approach and Landing Guidance – 9 Launch Vehicle Simulation – 11

A-4 Near-Optimal Operation of Dual-Fuel Computational Study of a McDonnell ASTRIONICS Launch Vehicles – 119 Douglas Single-Stage-to-Orbit Vehicle Research Reports: 1996 NASA/ASEE On-Board Entry Trajectory Planning Ex- Concept for Aerodynamic Analy- Summer Faculty Fellowship Pro- 126 panded to Sub-orbital Flight – 12 sis – gram – 102 On-Board Generation of Three- Effects of Stiffening and Mechanical ASTRONAUTICS Dimensional Constrained Entry Trajecto- Load on Thermal Buckling of Stiffened Selection of Lockheed Martin’s Preferred ries – 35 Cylindrical Shells – 121 TSTO Configurations for the Space Launch Initiative – 12 Optimal trajectories for an aerospace Managing MDO Software Development plane. Part 2: Data, tables, and Projects – 37 ATMOSPHERIC EFFECTS graphs – 176 Optimal technology investment strate- Aircraft Emission Scenarios Projected in Predictor-Corrector Entry Guidance for gies for a reusable launch vehicle – 134 Year 2015 for the NASA Technology Con- cept Aircraft (TCA) High Speed Civil Reusable Launch Vehicles – 74 The NASA Integrated Vehicle Health Transport – 108 Rapid near-optimal aerospace plane tra- Management Technology Experiment for jectory generation and guidance – 171 X-37 – 53 ATMOSPHERIC ENTRY On-Board Generation of Three- APPROACH CONTROL ALLOYS Dimensional Constrained Entry Trajecto- Selection and Evaluation of An Alloy for Personnel launch system autoland de- ries – 35 Nozzle Application – 22 velopment study – 175 Thermal Protection Materials for Reentry ALTITUDE CONTROL ARC HEATING and Planetary Applications – 68 Direct Fault Tolerant RLV Altitude Con- Magnetohydrodynamic Augmented Pro- ATMOSPHERIC MODELS trol: A Singular Perturbation Ap- pulsion Experiment – 52 proach – 40 Natural Atmospheric Environment Model ARC JET ENGINES Development for the National Aeronau- ALUMINUM OXIDES Thermal Protection System Evaluation tics and Space Administration’s Second Task 4 supporting technology. Part 1: Using Arc-jet Flows: Flight Simulation or Generation Reusable Launch Ve- Detailed test plan for leading edge tile Research Tool? – 31 hicle – 57 development. Leading edge material de- Stratospheric turbulence measurements 139 ARCHITECTURE (COMPUTERS) velopment and testing – and models for aerospace plane de- ALUMINUM-LITHIUM ALLOYS Autocommander: A Supervisory Control- sign – 164 ler for Integrated Guidance and Control Fatigue Crack Growth Rate Test Results for the 2nd Generation Reusable Launch ATMOSPHERIC TURBULENCE for Al-Li 2195 Parent Metal, Variable Po- Vehicle – 40 Lee waves: Benign and malig- larity Plasma Arc Welds and Friction Stir nant – 157 Welds – 89 Compute as Fast as the Engineers Can Think! ULTRAFAST COMPUTING TEAM Stratospheric turbulence measurements ALUMINUM FINAL REPORT – 90 and models for aerospace plane de- Development of Aluminum Composites sign – 164 for a Rocket Engine’s Lightweight Thrust ARGUS PROJECT Cell – 80 Highly Reusable Space Transportation ATOMIC RECOMBINATION Surface Catalysis and Characterization ANALOG TO DIGITAL CONVERTERS System Concept Evaluation (The Argus Launch Vehicle) – 103 of Proposed Candidate TPS for Access- Mosad and Stream Vision For A Telero- to-Space Vehicles – 118 botic, Flying Camera System – 59 ARTIFICIAL SATELLITES ATTITUDE CONTROL ANGLE OF ATTACK Estimate of avoidance maneuver rate for HASTOL tether boost facility – 28 Autocommander: A Supervisory Control- A simple analytical aerodynamic model ler for Integrated Guidance and Control of Langley Winged-Cone Aerospace ASCENT PROPULSION SYSTEMS for the 2nd Generation Reusable Launch 145 Plane concept – Technology Roadmap for Dual-Mode Vehicle – 40 Aerodynamic Characteristics of a Vertical Scramjet Propulsion to Support Space- Direct Fault Tolerant RLV Altitude Con- Takeoff Vertical Landing (VTVL) Single Access Vision Vehicle Develop- trol: A Singular Perturbation Ap- Stage to Orbit Vehicle from M(sub infin- ment – 38 proach – 40 ity) = 0.1 to 10 – 122 ASCENT TRAJECTORIES On-Line Computation of a Local Attain- Computational/Experimental Aeroheat- Analytical solutions to constrained hyper- able Moment Set for Reusable Launch ing Predictions for X-33 – 112 sonic flight trajectories – 162 Vehicles – 42 N-Factor Computations for the X-33 Ve- Effect of Aeropropulsive Interactions and Reusable Launch Vehicle Attitude Con- hicle – 76 Design Sensitivities on Optimal Hyper- trol Using a Time-Varying Sliding Mode Subsonic Aerodynamic Characteristics of sonic Ascent Trajectories – 125 Control Technique – 40 a Circular Body Earth-to-Orbit Ve- Near-Optimal Operation of Dual-Fuel AUTOCODERS hicle – 128 Launch Vehicles – 119 Autocommander: A Supervisory Control- Transonic aerodynamic characteristics of Trajectory optimization for the national ler for Integrated Guidance and Control a proposed wing-body reusable launch aerospace plane – 156 for the 2nd Generation Reusable Launch vehicle concept – 138 Vehicle – 40 ASSESSMENTS ANTARES ROCKET VEHICLE AUTOMATIC CONTROL Project Antares: A low cost modular An Overview of PRA with Applications to Aerospace Systems – 51 An Advanced Video Sensor for Auto- launch vehicle for the future – 168 mated Docking – 61 Operations Analysis of Space Shuttle APPLICATIONS PROGRAMS (COMPUT- Autocommander: A Supervisory Control- System – 50 ERS) ler for Integrated Guidance and Control A Thermal Management Systems Model System Level Uncertainty Assessment for the 2nd Generation Reusable Launch for the NASA GTX RBCC Concept – 53 for Collaborative RLV Design – 46 Vehicle – 40

A-5 Automated CFD Parameter Studies on BASE HEATING BOUNDARY CONDITIONS Distributed Parallel Computers – 36 Analysis of Aerospike Plume Induced On-Board Entry Trajectory Planning Ex- Fuzzy Logic Trajectory Design and Guid- Base-Heating Environment – 109 panded to Sub-orbital Flight – 12 ance for Terminal Area Energy Manage- Base-Bleed Effect on X-33 Aerospike BOUNDARY ELEMENT METHOD ment – 8 Plume Induced Base-Heating Environ- Development of an integrated BEM ap- ment During Power-Pack Out – 25 AUTOMATIC LANDING CONTROL proach for hot fluid structure interac- Personnel launch system autoland de- Unic Unstructured CFD Methodology De- tion – 169 velopment study – 175 velopment – 82 BOUNDARY LAYER FLOW AUTOMATIC PILOTS BIBLIOGRAPHIES Computational analysis in support of the SSTO flowpath test – 143 On-Line Computation of a Local Attain- Aeronautical Engineering: A Continuing able Moment Set for Reusable Launch Bibliography with Indexes – 97 Hypersonic flows as related to the Na- Vehicles – 42 National Aerospace Plane Thermal De- tional Aerospace plane – 172 Personnel launch system autoland development. (Latest Citations from the BOUNDARY LAYER SEPARATION velopment study – 175 Aerospace Database) – 113 Analysis of Aerospike Plume Induced Base-Heating Environment – 109 AUTOMATION BISMALEIMIDE Automation of NDE on RSRM Metal Damage tolerance of candidate thermo- Hypersonic flows as related to the na- Components – 57 set composites for use on single stage to tional aerospace plane – 162 orbit vehicles – 148 AUTONOMOUS NAVIGATION Hypersonic flows as related to the Na- Impact Characteristics of Candidate Ma- tional Aerospace plane – 172 Test Results for an Off-the-Shelf terials for Single-Stage-to-Orbit (SSTO) GPS/INS During Approach and Landing Technology – 123 BOUNDARY LAYER TRANSITION Testing of the X-40A – 63 Comparison of Methods for Determining BLANKETS AUTONOMY Boundary Layer Edge Conditions for Lightweight durable TPS: Milestone Transition Correlations – 23 NASA’s Orbital Space Plane Risk Re- 4–151 duction Strategy – 13 Hypersonic Boundary-Layer Transition BLUNT BODIES for X-33 Phase 2 Vehicle – 111 Orbital Space Plane Program Flight Demonstrators Status – 6 An implementation of a chemical and Infrared Sensing Aeroheating Flight Ex- thermal nonequilibrium flow solver on periment: STS-96 Flight Results – 86 Testing the Preliminary X-33 Navigation unstructured meshes and application to System – 84 blunt bodies – 145 X-33 (Rev-F) Aeroheating Results of Test 6770 in NASA Langley 20-Inch Mach 6 AUXILIARY POWER SOURCES BOATTAILS Air Tunnel – 96 High Voltage Design Guidelines: A Computations of Internal and External Timely Update – 61 Axisymmetric Nozzle Aerodynamics at BOUNDARY LAYERS Transonic Speeds – 4 Hypersonic flows as related to the Na- AUXILIARY PROPULSION tional Aerospace Plane – 136 BODY-WING CONFIGURATIONS Non-Toxic Dual Thrust Reaction Control N-Factor Computations for the X-33 Ve- Transonic aerodynamic characteristics of Engine Development for On-Orbit APS hicle – 76 Applications – 11 a proposed wing-body reusable launch vehicle concept – 138 BOUNDARY VALUE PROBLEMS AVIONICS BOLTS Air-breathing hypersonic vehicle guid- An Approach to Establishing System ance and control studies; An integrated X-33 Tank Failure During Autoclave Fab- Benefits for Technology in NASA’s Hy- trajectory/control analysis methodology: rication – 64 personics Investment Area – 67 Phase 1 – 171 NASA’s Spaceliner Investment Area BONDED JOINTS An Automated Method to Compute Or- Technology Activities – 81 Bonding and Sealing Evaluations for bital Re-Entry Trajectories with Heating Cryogenic Tanks – 116 Constraints – 53 AXIAL LOADS Effects of Stiffening and Mechanical BONDING BRANCHING (MATHEMATICS) Load on Thermal Buckling of Stiffened Bonding and Sealing Evaluations for An Improved RLV Stability Analysis Via a Cylindrical Shells – 121 Cryogenic Tanks – 116 Continuation Approach – 9 AXIOMS BOOSTER RECOVERY BRITTLENESS Compute as Fast as the Engineers Can Liquid flyback booster pre-phase: A study Evaluation of Thermal Control Coatings Think! ULTRAFAST COMPUTING TEAM assessment – 146 for Flexible Ceramic Thermal Protection Systems – 117 FINAL REPORT – 90 BOOSTER ROCKET ENGINES BALLISTIC MISSILES Analysis of the staging maneuver and CALIBRATING Reusable launch vehicle development booster glideback guidance for a two- Design and Calibration of the X-33 Flush research – 140 stage, winged, fully reusable launch ve- Airdata Sensing (FADS) System – 113 hicle – 159 Flush Airdata Sensing (FADS) System BASE FLOW The StarBooster System: A Cargo Air- Calibration Procedures and Results for Computation of H2/air reacting flowfields craft for Space – 92 Blunt Forebodies – 92 in drag-reduction external combustion – 161 BORIDES CALORIMETERS Reduction of Base Drag on Launch Ve- Tantalum Addition to Zirconium Diboride Heat flux sensor research and develop- hicles – 50 for Improved Oxidation Resistance – 18 ment: The cool film calorimeter – 170 Wind-Tunnel Investigations of Blunt- BOUNDARIES CAMERAS Body Drag Reduction Using Forebody Panel Flutter and Sonic Fatigue Analysis Mosad and Stream Vision For A Telero- Surface Roughness – 85 for RLV – 77 botic, Flying Camera System – 59

A-6 CARBON FIBER REINFORCED PLAS- CERAMICS CIVIL AVIATION TICS Design and Analysis of UHTC Leading Neural Network and Regression Approxi- Damage tolerance of candidate thermo- Edge Attachment – 44 mations in High Speed Civil Transport set composites for use on single stage to Aircraft Design Optimization – 106 orbit vehicles – 148 Health Monitoring Technology for Ther- mal Protection Systems on Reusable Hy- CLIMATE CHANGE CARBON FIBERS personic Vehicles – 71 Director’s Discretionary Fund Report for Characterization of C/SiC Ceramic Ma- Fiscal Year 1996 – 120 trix Composites (CMCs) with Novel Inter- High temperature NASP engine seal de- face Fiber Coatings – 45 velopment – 153 CLIMBING FLIGHT A Compendium of Wind Statistics and Permeability of Impacted Coated Com- Parametric Weight Comparison of Ad- Models for the NASA Space Shuttle and posite Laminates – 49 vanced Metallic, Ceramic Tile, and Ce- ramic Blanket Thermal Protection Sys- Other Aerospace Vehicle Pro- CARBON-CARBON COMPOSITES tems – 88 grams – 99 Surface Catalytic Efficiency of Advanced Tantalum Addition to Zirconium Diboride CLOSED CYCLES Carbon Carbon Candidate Thermal Pro- for Improved Oxidation Resistance – 18 tection Materials for SSTO Ve- Advanced Low-Cost O2/H2 Engines for hicles – 133 Task 4 supporting technology. Part 1: the SSTO Application – 123 Thermal/Fluid Analysis of a Composite Detailed test plan for leading edge tile COATINGS Heat Exchanger for Use on the RLV development. Leading edge material de- Characterization of C/SiC Ceramic Ma- Rocket Engine – 63 velopment and testing – 139 trix Composites (CMCs) with Novel Inter- face Fiber Coatings – 45 CARBON CERTIFICATION Demonstration of Advanced C/SiC Structures and materials technology for Permeability of Impacted Coated Com- Cooled Ramp – 56 hypersonic aerospacecraft – 178 posite Laminates – 49 CARDIOVASCULAR SYSTEM CHANNEL FLOW Task 4 supporting technology. Part 1: Detailed test plan for leading edge tile Shuttle to Space Station. Heart assist PNS predictions of external/internal hy- implant. Hubble update. X-30 mock- development. Leading edge material de- personic flows for NASP propulsion ap- velopment and testing – 139 up – 138 plications – 157 COAXIAL FLOW CATALYSIS CHANNELS Surface Catalysis and Characterization Experimental Evaluation of a Subscale Comparison of High Aspect Ratio Cool- of Proposed Candidate TPS for Access- Gaseous Hydrogen/Gaseous Oxygen ing Channel Designs for a Rocket Com- to-Space Vehicles – 118 Coaxial Rocket Injector – 34 bustion Chamber – 118 Thrust Enhancement in Hypervelocity COMBINED CYCLE POWER GENERA- Nozzles by Chemical Catalysis – 102 CHARGE COUPLED DEVICES TION Mosad and Stream Vision For A Telero- Second Generation RLV Program – 71 CATALYTIC ACTIVITY botic, Flying Camera System – 59 Surface Catalytic Efficiency of Advanced COMBUSTIBLE FLOW Carbon Carbon Candidate Thermal Pro- CHECKOUT An Experimental Study of Characteristic tection Materials for SSTO Ve- Real-Time Sensor Validation, Signal Re- Combustion-Driven Flow for CFD Valida- hicles – 133 construction, and Feature Detection for tion – 115 98 CERAMIC COATINGS an RLV Propulsion Testbed – COMBUSTION CHAMBERS Evaluation of Thermal Control Coatings CHEMICAL ANALYSIS Comparison of High Aspect Ratio Cool- for Flexible Ceramic Thermal Protection Advanced High Brilliance X-Ray ing Channel Designs for a Rocket Com- 117 Systems – Source – 111 bustion Chamber – 118 CERAMIC FIBERS Evaluation of Thermal Control Coatings Design and Testing of Non-Toxic RCS Test Plans. Lightweight Durable TPS: for Flexible Ceramic Thermal Protection Thrusters for Second Generation Reus- Tasks 1,2,4,5, and 6 – 146 Systems – 117 able Launch Vehicle – 14 CERAMIC MATRIX COMPOSITES Lightweight durable TPS: Milestone Lightweight Chambers for Thrust Assem- Ceramic Matrix Composite Cooled 4–151 blies – 66 Nozzle Material Development Pro- Status of the RBCC Direct-Connect CHEMICAL PROPULSION gram – 21 Mixer Combustor Experiment – 54 Ceramic Matrix Composites (CMC) Life Human lunar mission capabilities using The 2002 NASA Faculty Fellowship Pro- Prediction Method Development – 88 SSTO, ISRU and LOX-augmented NTR technologies: A preliminary assess- gram Research Reports – 9 Characterization of C/SiC Ceramic Ma- ment – 137 trix Composites (CMCs) with Novel Inter- COMBUSTION EFFICIENCY face Fiber Coatings – 45 CHEMICAL REACTIONS Design and Testing of Non-Toxic RCS A Novel Approach for Modeling Chemical Thrusters for Second Generation Reus- Fabrication and Testing of Ceramic Ma- able Launch Vehicle – 14 trix Composite Rocket Propulsion Com- Reaction in Generalized Fluid System 13 ponents – 75 Simulation Program – COMBUSTION PHYSICS Lightweight Chambers for Thrust Assem- CHIPS (ELECTRONICS) Computation of H2/air reacting flowfields blies – 82 Health Monitoring Technology for Ther- in drag-reduction external combus- mal Protection Systems on Reusable Hy- tion – 161 Materials Testing on the DC-X and DC- personic Vehicles – 71 XA – 121 COMBUSTION STABILITY Space Shuttle Main Engine Off-Nominal X-37 C-SiC CMC Control Surface Com- CIRCULAR ORBITS Low Power Level Operation – 117 ponents Development: Status of the Selection process for trade study: Reus- NASA/Boeing/USAF Orbital Vehicle and able Hydrogen Composite Tank System Vision for CFD-Based Combustion Insta- Related Efforts – 20 (RHCTS) – 148 bility Predictions – 23

A-7 COMBUSTION Reusable Launch Vehicle Technology CFD-Based Design Optimization for Analysis of Parallel Burn Without Cross- Program – 127 Single Element Rocket Injector – 25 feed TSTO RLV Architectures and Com- Selection process for trade study: Graph- Comparison of Methods for Determining parison to Parallel Burn With Crossfeed ite Composite Primary Structure Boundary Layer Edge Conditions for and Series Burn Architectures – 19 (GCPS) – 148 Transition Correlations – 23 COMETS Structural Arrangement Trade Study. Vol- Computational Aerodynamics Analysis of Director’s Discretionary Fund Report for ume 1: Reusable Hydrogen Composite Future Launch Vehicle Configura- Fiscal Year 1996 – 120 Tank System (RHCTS) and Graphite tions – 122 COMMAND AND CONTROL Composite Primary Structures (GCPS). Executive summary – 141 Computational parametric study of A guidance and control assessment of sidewall-compression scramjet inlet per- three vertical landing options for Structural arrangement trade study. Vol- formance at Mach 10 – 161 RLV – 136 ume 3: Reusable Hydrogen Composite Tank System (RHCTS) and Graphite Computational/Experimental Aeroheat- COMMAND GUIDANCE ing Predictions for X-33 Phase 2 Ve- A guidance and control assessment of Composite Primary Structures (GCPS). 141 hicle – 110 three vertical landing options for Addendum – RLV – 136 Study Acoustic Emissions from Compos- Computational/Experimental Aeroheat- ing Predictions for X-33 – 112 COMMERCE ites – 111 Analysis of space concepts enabled by Test Plan. GCPS Task 4, subtask 4.2 Design and Analysis of Turbines for new transportation (ASCENT) thrust structure development – 147 Space Applications – 16 study – 28 Trade study plan for Graphite Composite Evaluation and Characterization Study of COMMERCIAL SPACECRAFT Primary Structure (GCPS) – 146 Dual Pulse Laser-Induced Spark (DPLIS) Second Generation Reusable Launch for Rocket Engine Ignition System Appli- Trade study plan for Reusable Hydrogen cation – 14 Vehicle Development and Global Com- Composite Tank System petitiveness of US Space Transportation (RHCTS) – 145 Flight Testing the Linear Aerospike Industry: Critical Success Factors As- SR-71 Experiment (LASRE) – 100 sessment – 51 COMPRESSIBILITY Turbulence modeling for hypersonic Hypersonic flows as related to the Na- COMMERCIALIZATION flight – 161 tional Aerospace Plane – 136 2nd Generation Reusable Launch Ve- hicle Potential Commercial Development COMPRESSIBLE FLOW Launch Vehicle Sizing Benefits Utilizing Scenarios – 79 Development of an integrated BEM ap- Main Propulsion System Crossfeed and proach for hot fluid structure interac- Project Status – 34 COMMUNICATION tion – 169 The X-33 Extended Flight Test NASA aeronautics: Research and tech- Range – 99 COMPRESSION LOADS nology program highlights – 164 Analytical comparison of three stiffened NASA EPSCoR Preparation Grant – 49 COMPOSITE MATERIALS panel concepts – 135 A Damage Resistance Comparison Be- N-Factor Computations for the X-33 Ve- tween Candidate Polymer Matrix Com- COMPRESSION TESTS hicle – 76 posite Feedline Materials – 88 Task 4 supporting technology. Part 2: Detailed test plan for thermal seals. Ther- Opportunities for Breakthroughs in Composite Development and Applica- mal seals evaluation, improvement and Large-Scale Computational Simulation tions for RLV Tankage – 16 test. CAN8-1, Reusable Launch Vehicle and Design – 54 Current Technology for Thermal Protec- (RLV), advanced technology demonstra- Overview of MSFC’s Applied Fluid Dy- tion Systems – 165 tor: X-33. Leading edge and seals ther- namics Analysis Group Activities – 24 mal protection system technology dem- Demonstration of Advanced C/SiC onstration – 139 Performance Evaluation of the NASA Cooled Ramp – 56 GTX RBCC Flowpath – 70 COMPUTATIONAL FLUID DYNAMICS Permeability of Impacted Coated Com- PNS predictions of external/internal hy- posite Laminates – 49 A method of predicting quasi-steady aerodynamics for flutter analysis of high personic flows for NASP propulsion ap- Test Plans. Lightweight Durable TPS: speed vehicles using steady CFD calcu- plications – 157 Tasks 1,2,4,5, and 6 – 146 157 lations – RLV Turbine Performance Optimiza- Thermal-Mechanical Cyclic Test of a Accelerating the Design of Space Ve- tion – 83 Composite Cryogenic Tank for Reusable hicles – 10 Launch Vehicles – 23 Stage Separation CFD Tool Develop- Addition of equilibrium air to an upwind ment and Evaluation – 38 COMPOSITE STRUCTURES Navier-Stokes code and other first steps Current Technology for Thermal Protec- toward a more generalized flow The aerospace plane design challenge: tion Systems – 165 solver – 174 Credible computational fluid dynamics results – 176 Detection of Micro-Leaks Through Com- An Experimental Study of Characteristic plex Geometries Under Mechanical Load Combustion-Driven Flow for CFD Valida- Thermal Protection System Evaluation and at Cryogenic Temperature – 79 tion – 103 Using Arc-jet Flows: Flight Simulation or Research Tool? – 31 Large Composite Structures Processing An Inviscid Computational Study of an Technologies for Reusable Launch Ve- X-33 Configuration at Hypersonic Time-Dependent Simulations of Tur- hicles – 75 Speeds – 94 bopump Flows – 38 Milestone 4: Thrust structure concepts Analysis of X-33 Linear Aerospike Plume Unic Unstructured CFD Methodology De- and IHM screening Graphite Composite Induced Base-Heating Environment Dur- velopment – 82 Primary Structure (GCPS) – 143 ing Power-Pack Out – 29 Using CFD as Rocket Injector Design Reusable launch vehicle: Technology de- Automated CFD Parameter Studies on Tool: Recent Progress at Marshall Space velopment and test program – 135 Distributed Parallel Computers – 36 Flight Center – 3

A-8 Vision for CFD-Based Combustion Insta- Comparison of High Aspect Ratio Cool- Time-Dependent Simulations of Tur- bility Predictions – 23 ing Channel Designs for a Rocket Com- bopump Flows – 38 bustion Chamber – 118 X-33 Computational CONFERENCES Aeroheating/Aerodynamic Predictions Evaluation of Advanced Thermal Protec- 2000 NASA Seal/Secondary Air System and Comparisons With Experimental tion Techniques for Future Reusable Workshop – 62 Data – 25 Launch Vehicles – 62 2nd Generation Reusable Launch Ve- X-33 Rev-F Turbulent Aeroheating Re- Multidisciplinary Optimization Branch Ex- hicle Potential Commercial Development sults From Test 6817 in NASA Langley perience Using iSIGHT Software – 91 Scenarios – 79 20-Inch Mach 6 Air Tunnel and Compari- NASA Overview – 35 sons With Computations – 26 NASA Exhibits – 59 Prediction of the ullage gas thermal X-33 Turbulent Aeroheating Measure- Rocket Based Combined Cycle (RBCC) stratification in a NASP vehicle propellant ments and Predictions – 32 Propulsion Workshop, volume 2 – 166 tank experimental simulation using CONFIGURATION MANAGEMENT COMPUTATIONAL GRIDS FLOW-3D – 177 Reconfigurable Control Design for the Computational Study of a McDonnell Stage Separation CFD Tool Develop- Full X-33 Flight Envelope – 72 Douglas Single-Stage-to-Orbit Vehicle ment and Evaluation – 38 Concept for Aerodynamic Analy- CONTINUITY (MATHEMATICS) X-33 Turbulent Aeroheating Measure- sis – 126 An Improved RLV Stability Analysis Via a ments and Predictions – 32 Progress in Unsteady Turbopump Flow Continuation Approach – 9 COMPUTER STORAGE DEVICES Simulations Using Overset Grid Sys- CONTRACT MANAGEMENT X-33 Development History – 113 tems – 30 Pathfinder Technologies Specialist, Progress in Unsteady Turbopump Flow COMPUTER SYSTEMS DESIGN X-37 – 69 Simulations – 51 Compute as Fast as the Engineers Can Think! ULTRAFAST COMPUTING TEAM CONTROL EQUIPMENT Time-Dependent Simulations of Tur- FINAL REPORT – 90 Direct Fault Tolerant RLV Altitude Con- bopump Flows – 38 trol: A Singular Perturbation Ap- COMPUTER SYSTEMS PERFORMANCE COMPUTATION proach – 40 Multidisciplinary Optimization Branch Ex- Reusable Launch Vehicle Attitude Con- NASA Exhibits – 59 perience Using iSIGHT Software – 91 trol Using a Time-Varying Sliding Mode COMPUTER AIDED DESIGN COMPUTER SYSTEMS PROGRAMS Control Technique – 40 A Geometry-Centered Approach For X-33 Development History – 113 Two-Stage-To-Orbit Vehicle Synthe- CONTROL SURFACES sis – 45 COMPUTERIZED SIMULATION Advanced Control Surface Seal Develop- A Monte Carlo Dispersion Analysis of the ment at NASA GRC for 3rd Generation Managing MDO Software Development X-33 Simulation Software – 72 RLV – 29 Projects – 37 A Novel Approach for Modeling Chemical Advanced High Temperature Structural Multidisciplinary Optimization Branch Ex- Reaction in Generalized Fluid System Seals – 87 perience Using iSIGHT Software – 91 13 Simulation Program – Aerodynamic Characteristics of a Vertical Opportunities for Breakthroughs in Analysis of X-33 Linear Aerospike Plume Takeoff Vertical Landing (VTVL) Single Large-Scale Computational Simulation Induced Base-Heating Environment Dur- Stage to Orbit Vehicle from M(sub infin- and Design – 54 ing Power-Pack Out – 29 ity) = 0.1 to 10 – 122 Optimal technology investment strate- Computational Aerodynamics Analysis of Evaluation of Computational Aeroelastic- gies for a reusable launch vehicle – 134 Future Launch Vehicle Configura- ity Codes for Loads and Flutter – 82 COMPUTER NETWORKS tions – 122 Third Generation RLV Structural Seal Managing MDO Software Development Effects of Stiffening and Mechanical Development Programs at NASA Projects – 37 Load on Thermal Buckling of Stiffened GRC – 3 Cylindrical Shells – 121 COMPUTER PROGRAMS Transonic aerodynamic characteristics of A Geometry-Centered Approach For Improvements in Thermal Protection Siz- a proposed wing-body reusable launch Two-Stage-To-Orbit Vehicle Synthe- ing Capabilities for TCAT: Conceptual vehicle concept – 138 Design for Advanced Space Transporta- sis – 45 Update on the Development and Capa- tion Systems – 61 A Hazardous Gas Detection System for bilities of Unique Structural Seal Test Launch Vehicle Sizing Benefits Utilizing Aerospace and Commercial Applica- Rigs – 3 Main Propulsion System Crossfeed and tions – 98 Project Status – 34 X-37 C-SiC CMC Control Surface Com- A Method of Integrating Aeroheating into ponents Development: Status of the NASA aeronautics: Research and tech- Conceptual Reusable Launch Vehicle NASA/Boeing/USAF Orbital Vehicle and nology program highlights – 164 Design: Evaluation of Advanced Thermal Related Efforts – 20 Protection Techniques for Future Reus- Results of Two-Stage Light-Gas Gun De- CONTROL SYSTEMS DESIGN able Launch Vehicles – 77 velopment Efforts and Hypervelocity Im- A guidance and control assessment of pact Tests of Advanced Thermal Protec- A Monte Carlo Dispersion Analysis of the three vertical landing options for tion Materials – 100 X-33 Simulation Software – 72 RLV – 136 System Level Uncertainty Assessment A Novel Approach for Modeling Chemical Advanced Guidance and Control Meth- for Collaborative RLV Design – 46 Reaction in Generalized Fluid System ods for Reusable Launch Vehicles: Test Simulation Program – 13 Thermal/Fluid Analysis of a Composite Results – 30 Heat Exchanger for Use on the RLV An implementation of a chemical and Air-breathing hypersonic vehicle guid- Rocket Engine – 63 thermal nonequilibrium flow solver on ance and control studies: An integrated unstructured meshes and application to Time Dependent Simulation of Tur- trajectory/control analysis methodology, blunt bodies – 145 bopump Flows – 55 phase 2 – 165

A-9 Analysis of the staging maneuver and Analysis of X-33 Linear Aerospike Plume Estimating the Cost of NASA’s Space booster glideback guidance for a two- Induced Base-Heating Environment Dur- Launch Initiative: How SLI Cost Stack Up stage, winged, fully reusable launch ve- ing Power-Pack Out – 29 Against the Shuttle – 51 hicle – 159 Evaluation of Advanced Thermal Protec- High Altitude Launch for a Practical Dynamics and Stability and Control tion Techniques for Future Reusable SSTO – 32 Characteristics of the X-37 – 62 Launch Vehicles – 62 Orbital Space Plane Cost Credibility – 4 Dynamics of aerospace vehicles – 174 Minimum Heating Re-Entry Trajectories for Advanced Hypersonic Launch Ve- Full Envelope Reconfigurable Control COST REDUCTION hicles – 119 Design for the X-33 Vehicle – 72 Affordable Flight Demonstration of the CONVERGENCE GTX Air-Breathing SSTO Vehicle Con- Hypersonic vehicle control law develop- cept – 46 ment using H infinity and mu- Comparison of Two Multidisciplinary Op- synthesis – 164 timization Strategies for Launch-Vehicle Affordable In-Space Transportation Design – 123 Phase 2: An Advanced Concepts Hypersonic vehicle model and control Project – 113 law development using H(infinity) and CONVERGENT-DIVERGENT NOZZLES micron synthesis – 144 Computations of Internal and External Affordable In-Space Transporta- Axisymmetric Nozzle Aerodynamics at tion – 106 Personnel launch system autoland de- Transonic Speeds – 4 velopment study – 175 An Approach to Establishing System COOLING SYSTEMS Reconfigurable Control Design for the Benefits for Technologies In NASA’s Full X-33 Flight Envelope – 72 Lightweight Chambers for Thrust Assem- Spaceliner Investment Area – 77 blies – 66 Reconfigurable Flight Control Designs Conceptual design of two-stage-to-orbit With Application to the X-33 Ve- COOLING hybrid launch vehicle – 169 hicle – 95 Comparison of High Aspect Ratio Cool- ing Channel Designs for a Rocket Com- Design, Development,and Testing of Um- Robust intelligent flight control for hyper- bustion Chamber – 118 billical System Mechanisms for the X-33 sonic vehicles – 166 Advanced Technology Demonstra- Demonstration of Advanced C/SiC tor – 90 The Control System for the X-33 Linear Cooled Ramp – 56 Aerospike Engine – 105 Highly reusable space transportation: COORDINATION Approaches for reducing ETO launch CONTROL THEORY Affordable In-Space Transporta- costs to $100 - $200 per pound of pay- Hypersonic vehicle control law develop- tion – 106 load – 135 ment using H infinity and mu- COPPER ALLOYS Reusable Launch Vehicle Re- synthesis – 164 Lightweight Chambers for Thrust Assem- search – 119 Reconfigurable Control Design for the blies – 82 Full X-33 Flight Envelope – 72 Second Generation Reusable Launch Oxidation Behavior of Copper Alloy Can- Vehicle Development and Global Com- CONTROLLABILITY didates for Rocket Engine Applications petitiveness of US Space Transportation (Technical Poster) – 48 Industry: Critical Success Factors As- Flight Stability and Control and Perfor- sessment – 51 mance Results from the Linear Aerospike Oxidation-Reduction Resistance of Ad- SR-71 Experiment (LASRE) – 101 vanced Copper Alloys – 11 Structural Seal Development – 11 Subsonic-to-Hypersonic Aerodynamic COST ANALYSIS X-33/RLV System Health Management/ Characteristics for a Winged, Circular- Advanced transportation system studies Vehicle Health Management – 107 Body, Single-Stage-to-Orbit Spacecraft technical area 2 (TA-2): Heavy lift launch Configuration – 124 vehicle development – 130 COSTS Affordable In-Space Transportation CONTROLLERS An Approach to Establishing System Benefits for Technologies In NASA’s Phase 2: An Advanced Concepts A guidance and control assessment of Project – 113 three vertical landing options for Spaceliner Investment Area – 77 RLV – 136 Analysis of space concepts enabled by Characterization of Polyimide Foams for new transportation (ASCENT) Ultra-Lightweight Space Struc- Air-breathing hypersonic vehicle guid- tures – 22 ance and control studies; An integrated study – 28 trajectory/control analysis methodology: O/S analysis of conceptual space ve- Cost Per Pound From Orbit – 47 Phase 1 – 171 hicles – 134 Statistical Methods for Rapid Aerother- Full Envelope Reconfigurable Control COST EFFECTIVENESS mal Analysis and Design Technol- ogy – 55 Design for the X-33 Vehicle – 72 An Approach to Establishing System Improved Re-Configurable Sliding Mode Benefits for Technologies In NASA’s Tether Transportation System Controller for Reusable Launch Vehicle Spaceliner Investment Area – 77 Study – 109 of Second Generation Addressing Aero- Analysis of space concepts enabled by dynamic Surface Failures and Thrust De- new transportation (ASCENT) COWLINGS ficiencies – 34 study – 28 Computational analysis in support of the SSTO flowpath test – 143 Integrated G and C Implementation Single-stage-to-orbit: Meeting the chal- within IDOS: A Simulink Based Reusable lenge – 136 Launch Vehicle Simulation – 11 CRACK PROPAGATION COST ESTIMATES Fatigue Crack Growth Rate Test Results CONVECTIVE HEAT TRANSFER Analysis of space concepts enabled by for Al-Li 2195 Parent Metal, Variable Po- Analysis of Aerospike Plume Induced new transportation (ASCENT) larity Plasma Arc Welds and Friction Stir Base-Heating Environment – 109 study – 28 Welds – 89

A-10 CRACKING (FRACTURING) CRYOGENIC ROCKET PROPELLANTS CYLINDRICAL SHELLS Fatigue Crack Growth Rate Test Results A Cryogenic Propellant Production Depot Effects of Stiffening and Mechanical for Al-Li 2195 Parent Metal, Variable Po- for Low Earth Orbit – 81 Load on Thermal Buckling of Stiffened larity Plasma Arc Welds and Friction Stir Cylindrical Shells – 121 Performance Tests of a Liquid Hydrogen Welds – 89 Propellant Densification Ground System DAMAGE ASSESSMENT CRACKS for the X33/RLV – 118 Wireless Subsurface Microsensors for Fatigue Crack Growth Rate Test Results Preliminary Assessment of Using Gelled Health Monitoring of Thermal Protection for Al-Li 2195 Parent Metal, Variable Po- and Hybrid Propellant Propulsion for Systems on Hypersonic Vehicles – 69 larity Plasma Arc Welds and Friction Stir VTOL/SSTO Launch Systems – 104 Welds – 89 DAMAGE Propellant Densification for Shuttle: The A Damage Resistance Comparison Be- CREEP ANALYSIS SSME Perspective – 32 tween Candidate Polymer Matrix Com- Advanced Metallic Thermal Protection posite Feedline Materials – 88 System Development – 64 Recent Advances and Applications in Cryogenic Propellant Densification Tech- Advanced Durability and Damage Toler- CREEP PROPERTIES nology – 89 ance Design and Analysis Methods for Thermoviscoplastic response of Ti-15-3 Reusable Cryogenic Tank VHM Using Composite Structures: Lessons Learned under various loading conditions – 172 Fiber Optic Distributed Sensing Technol- from NASA Technology Development Programs – 26 CRUISING FLIGHT ogy – 108 Analytical solutions to constrained hyper- Test plan. Task 5, subtask 5.2: Early Test plan. Task 5, subtask 5.2: Early sonic flight trajectories – 162 on-orbit TPSdebris impact tests – 148 on-orbit TPSdebris impact tests – 148 DATA ACQUISITION CRYOGENIC FLUID STORAGE CRYOGENIC TEMPERATURE Infrared Sensing Aeroheating Flight Ex- Advanced Metallic Thermal Protection Cryopumping in Cryogenic Insulations periement: STS-96 Flight Results – 84 System Development – 64 for a Reusable Launch Vehicle – 2 Bonding and Sealing Evaluations for Detection of Micro-Leaks Through Com- DATA BASES Cryogenic Tanks – 116 plex Geometries Under Mechanical Load Highly Reusable Space Transportation Characterization of Thin Film Polymers and at Cryogenic Temperature – 79 (HRST) Baseline Concepts and Analysis: Rocket/RBCC Options – 114 Through Dynamic Mechanical Analysis High Temperature Polyimide Materials in and Permeation – 22 Extreme Temperature Environ- Inlet Development for a Rocket Based Development of a Pressure Box to Evalu- ments – 73 Combined Cycle, Single Stage to Orbit ate Reusable-Launch-Vehicle Vehicle Using Computational Fluid Dy- Study Acoustic Emissions from Compos- namics – 95 Cryogenic-Tank Panels – 132 ites – 111 NASP aeroservothermoelasticity stud- Evaluation of Microcracking in Two ies – Carbon-Fiber/Epoxy-Matrix Composite CRYOGENICS 174 Cryogenic Tanks – 72 Composite Development and Applica- Reusable Rocket Engine Operability tions for RLV Tankage – 16 Manufacturing Process Simulation of Modeling and Analysis – 101 NASA’s Space Launch Initiative Targets Large-Scale Cryotanks – 36 DATA PROCESSING Toxic Propellants – 64 Reusable Cryogenic Tank VHM Using The 2002 NASA Faculty Fellowship Pro- Fiber Optic Distributed Sensing Technol- Slush hydrogen technology program Research Reports – 9 gram – 149 ogy – 108 X-33 Development History – 113 Reusable launch vehicle: Technology de- Technology issues associated with using DATA SMOOTHING velopment and test program – 135 densified hydrogen for space vehicles – 167 Air-breathing hypersonic vehicle guid- Reusable Launch Vehicle Technology ance and control studies; An integrated Program – 127 CRYOPUMPING trajectory/control analysis methodology: Slush hydrogen propellant production, Cryopumping in Cryogenic Insulations Phase 1 – 171 transfer, and expulsion studies at the for a Reusable Launch Vehicle – 2 DATA SYSTEMS NASA K-Site Facility – 173 CURING Design and Calibration of the X-33 Flush Slush hydrogen transfer studies at the Airdata Sensing (FADS) System – 113 A Damage Resistance Comparison Be- NASA K-Site Test Facility – 167 tween Candidate Polymer Matrix Com- DEBONDING (MATERIALS) Support to X-33/Reusable Launch Ve- posite Feedline Materials – 88 X-33 Tank Failure During Autoclave Fab- hicle Technology Program – 86 rication – 64 CYCLIC LOADS Testing of Densified Liquid Hydrogen DECISION MAKING Stratification in a Scale Model Propellant Cyclic Cryogenic Thermal-Mechanical Tank – 67 Testing of an X-33/RLV Liquid Oxygen Using Neural Networks in Decision Mak- Tank Concept – 94 ing for a Reconfigurable Electro Me- Thermal Structures Technology Develop- chanical Actuator (EMA) – 68 ment for Reusable Launch Vehicle Cryo- Thermoviscoplastic response of Ti-15-3 genic Propellant Tanks – 107 under various loading conditions – 172 X-33/RLV System Health Management/ Vehicle Health Management – 107 Thermal-Mechanical Cyclic Test of a CYLINDRICAL BODIES Composite Cryogenic Tank for Reusable An implementation of a chemical and DEFECTS Launch Vehicles – 23 thermal nonequilibrium flow solver on X-33 Tank Failure During Autoclave Fab- rication – 64 CRYOGENIC FLUIDS unstructured meshes and application to blunt bodies – 145 Thermal-Mechanical Cyclic Test of a DEGRADATION Composite Cryogenic Tank for Reusable The 2002 NASA Faculty Fellowship Pro- Oxidation-Reduction Resistance of Ad- Launch Vehicles – 23 gram Research Reports – 9 vanced Copper Alloys – 11

A-11 DELTA CLIPPER Evaluation of Advanced Thermal Protec- System Sensitivity Analysis Applied to Materials Testing on the DC-X and DC- tion Techniques for Future Reusable the Conceptual Design of a Dual-Fuel XA – 121 Launch Vehicles – 62 Rocket SSTO – 125 DELTA WINGS Full Envelope Reconfigurable Control The aerospace plane design challenge: NASP aeroservothermoelasticity stud- Design for the X-33 Vehicle – 72 Credible computational fluid dynamics results – 176 ies – 174 GTX Reference Vehicle Structural Verifi- Subsonic Aerodynamic Characteristics of cation Methods and Weight Sum- Unic Unstructured CFD Methodology De- a Circular Body Earth-to-Orbit Ve- mary – 44 velopment – 82 hicle – 128 High Voltage Design Guidelines: A DESIGN OPTIMIZATION DENSIFICATION Timely Update – 61 An Automated Method to Compute Or- bital Re-Entry Trajectories with Heating Performance Tests of a Liquid Hydrogen Highly Reusable Space Transportation Constraints – 53 Propellant Densification Ground System System Concept Evaluation (The Argus for the X33/RLV – 118 Launch Vehicle) – 103 CFD-Based Design Optimization for Single Element Rocket Injector – 25 Recent Advances and Applications in Inlet Development for a Rocket Based Cryogenic Propellant Densification Tech- Combined Cycle, Single Stage to Orbit Cost Per Pound From Orbit – 47 89 nology – Vehicle Using Computational Fluid Dy- RLV Turbine Performance Optimiza- Task 4 supporting technology. Densifica- namics – 95 tion – 83 tion requirements definition and test ob- Lightweight Chambers for Thrust Assem- Structures for the 3rd Generation Reus- jectives. Propellant densification require- blies – 66 able Concept Vehicle – 69 ments definition – 139 Main Engine Prototype Development for DESIGN TO COST DENSITY 2nd Generation RLV RS-83 – 43 Technology and Risk Assessment Using High Altitude Launch for a Practical Multi-Disciplinary Analysis for Future a Robust Design Simulation Methodol- SSTO – 32 Launch Systems Using NASA’s Ad- ogy – 121 DESCENT TRAJECTORIES vanced Engineering Environment Test Plan. GCPS Task 4, subtask 4.2 Near-Optimal Re-Entry Trajectories for (AEE) – 17 thrust structure development – 147 Reusable Launch Vehicles – 115 Multidisciplinary design techniques ap- The 2002 NASA Faculty Fellowship Pro- DESIGN ANALYSIS plied to conceptual aerospace vehicle gram Research Reports – 9 design – 156 A Two-Stage-to-Orbit Spaceplane Con- Thermal Management Design for the cept With Growth Potential – 80 Multidisciplinary Optimization Branch Ex- X-33 Lifting Body – 104 perience Using iSIGHT Software – 91 Advanced aeroservoelastic stabilization DESIGN techniques for hypersonic flight ve- Operations Analysis of the 2nd Genera- Comparison of High Aspect Ratio Cool- hicles – 163 tion Reusable Launch Vehicle – 43 ing Channel Designs for a Rocket Com- Advanced Control Surface Seal Develop- Optimal technology investment strate- bustion Chamber – 118 ment at NASA GRC for 3rd Generation gies for a reusable launch vehicle – 134 DESTRUCTIVE TESTS RLV – 29 Overview of MSFC’s Applied Fluid Dy- Thermal-Mechanical Cyclic Test of a Advanced Durability and Damage Toler- namics Analysis Group Activities – 24 Composite Cryogenic Tank for Reusable ance Design and Analysis Methods for Launch Vehicles – 23 Composite Structures: Lessons Learned Performance Tests of a Liquid Hydrogen DETECTION from NASA Technology Development Propellant Densification Ground System Programs – 26 for the X33/RLV – 118 Flush Airdata Sensing (FADS) System Calibration Procedures and Results for Advanced High Temperature Structural Propulsion module design review. Mile- Blunt Forebodies – 92 Seals – 43 stone 8: Integrated Propulsion Technol- ogy Demonstrator – 150 DEVELOPMENT Alternate Propulsion Subsystem Con- NASA’s New Orbital Space Plane: A Requirements report for SSTO vertical cepts Tripropellant Comparison Bridge to the Future – 19 Study – 132 take-off and horizontal landing vehicle – 143 DIAGNOSIS Analytical comparison of three stiffened Health Monitoring Technology for Ther- panel concepts – 135 Requirements report for SSTO vertical take-off/horizontal landing ve- mal Protection Systems on Reusable Hy- 71 Background, current status, and progno- hicle – 149 personic Vehicles – sis of the ongoing slush hydrogen tech- NASA IVHM Technology Experiment for Results of a Rocket-Based Combined- nology development program for the X-vehicles (NITEX) – 67 NASP – 177 Cycle SSTO Design Using Parametric MDO Methods – 121 Wireless Subsurface Sensors for Health Comparison of Two Multidisciplinary Op- Monitoring of Thermal Protection Sys- Reusable Rocket Engine Operability timization Strategies for Launch-Vehicle tems on Hypersonic Vehicles – 68 Design – 123 Modeling and Analysis – 101 DIGITAL CAMERAS Selection process for trade study: Reus- Conceptual design of two-stage-to-orbit NASA Stennis Space Center V and V able Hydrogen Composite Tank System hybrid launch vehicle – 169 Capabilities Overview – 2 (RHCTS) – 148 Design and Analysis of Turbines for DIGITAL SIMULATION Space Applications – 16 System Identification of X-33 Neural Net- work – 8 Computational Aerodynamics Analysis of Design and Analysis of UHTC Leading Future Launch Vehicle Configura- Edge Attachment – 44 System Level Uncertainty Assessment tions – 122 for Collaborative RLV Design – 46 Dual-Fuel Propulsion in Single-Stage Ad- Effects of Stiffening and Mechanical vanced Manned Launch System Ve- System Risk Assessment and Allocation Load on Thermal Buckling of Stiffened hicle – 125 in Conceptual Design – 27 Cylindrical Shells – 121

A-12 DIRECT NUMERICAL SIMULATION Evaluation of Thermal Control Coatings EDUCATION Computational/Experimental Aeroheat- for Flexible Ceramic Thermal Protection Research Reports: 1996 NASA/ASEE ing Predictions for X-33 Phase 2 Ve- Systems – 117 Summer Faculty Fellowship Pro- hicle – 110 Lightweight durable TPS: Milestone gram – 102 DIRECTIONAL CONTROL 4–151 The Center for Aerospace Research: A Subsonic Aerodynamic Characteristics of Materials Research for High Speed Civil NASA Center of Excellence at North a Circular Body Earth-to-Orbit Ve- Transport and Generic Hypersonics- Carolina Agricultural and Technical State 162 hicle – 128 Metals Durability – 127 University – EFFICIENCY DIRECTIONAL STABILITY DYNAMIC CHARACTERISTICS Subsonic aerodynamic characteristics of Advanced High Brilliance X-Ray Dynamics and Stability and Control a proposed advanced manned launch Source – 111 Characteristics of the X-37 – 62 160 system orbiter configuration – EJECTORS DISCOVERY (ORBITER) DYNAMIC MODELS Operationally efficient propulsion system STS-102 Photo-op/Suit-up/Depart Advanced aeroservoelastic stabilization study (OEPSS) data book – 110 techniques for hypersonic flight ve- O&C/Launch Discovery On Rocket-Induced Magnetohydrodynamic hicles – 163 Orbit/Landing/Crew Egress – 27 Ejector: A Single-Stage-to-Orbit Ad- DISPERSION An Experimental Study of Characteristic vanced Propulsion Concept – 131 Combustion-Driven Flow for CFD Valida- A Monte Carlo Dispersion Analysis of the Supercharged Ejector Ramjet with Ma- tion – 103 X-33 Simulation Software – 72 glifter Launch Assist (for the Argus Con- DISPLACEMENT Hypersonic vehicle model and control cept) – 120 law development using H(infinity) and Panel Flutter and Sonic Fatigue Analysis micron synthesis – 144 ELASTIC MEDIA for RLV – 77 Development of an integrated BEM ap- DISTRIBUTED PROCESSING DYNAMIC RESPONSE proach for hot fluid structure interac- Automated CFD Parameter Studies on Aerothermoelastic analysis of a NASP tion – 169 demonstrator model – 155 Distributed Parallel Computers – 36 ELECTRIC PROPULSION Managing MDO Software Development Design and Analysis of Turbines for Advanced Electric Propulsion for RLV Projects – 37 Space Applications – 16 Launched Geosynchronous Space- craft – 90 NASA Exhibits – 59 DYNAMIC STABILITY NASA IVHM Technology Experiment for Supersonic aeroelastic instability results Advanced spacecraft designs in support X-vehicles (NITEX) – 67 for a NASP-like wing model – 159 of human missions to earth’s neighborhood – 28 DIVERTERS DYNAMIC STRUCTURAL ANALYSIS ELECTRODES Computational analysis in support of the A method of predicting quasi-steady Durability of Membrane Electrode As- SSTO flowpath test – 143 aerodynamics for flutter analysis of high semblies (MEAs) in PEM Fuel Cells Op- speed vehicles using steady CFD calcu- DOWNLINKING erated on Pure Hydrogen and Oxy- lations – 157 X-33 Integrated Test Facility Extended gen – 6 Range Simulation – 100 Advanced aeroservoelastic stabilization techniques for hypersonic flight ve- ELECTROMAGNETIC ACCELERATION DRAG MEASUREMENT hicles – 163 Magnetohydrodynamic Augmented Pro- A Base Drag Reduction Experiment on pulsion Experiment – 52 the X-33 Linear Aerospike SR-71 Experi- Dynamics of aerospace vehicles – 174 ment (LASRE) Flight Program – 97 NASA aeronautics: Research and tech- ELECTROMECHANICAL DEVICES nology program highlights – 164 Using Neural Networks in Decision Mak- DRAG REDUCTION ing for a Reconfigurable Electro Me- A Base Drag Reduction Experiment on Structural arrangement trade study. Vol- chanical Actuator (EMA) – 68 the X-33 Linear Aerospike SR-71 Experi- ume 3: Reusable Hydrogen Composite ment (LASRE) Flight Program – 97 Tank System (RHCTS) and Graphite ELEVONS Composite Primary Structures (GCPS). Advanced High Temperature Structural Computation of H2/air reacting flowfields Addendum – 141 Seals – 43 in drag-reduction external combustion – 161 DYNAMIC TESTS EMISSION Reduction of Base Drag on Launch Ve- Hot dynamic test rig for measuring hyper- Aircraft Emission Scenarios Projected in hicles – 50 sonic engine seal flow and durabil- Year 2015 for the NASA Technology Con- ity – 151 cept Aircraft (TCA) High Speed Civil Trim drag reduction concepts for horizon- Transport – 108 tal takeoff single-stage-to-Orbit ve- EARTH ORBITS hicles – 170 EMITTANCE Advanced spacecraft designs in support Wind-Tunnel Investigations of Blunt- Surface Catalysis and Characterization of human missions to earth’s neighbor- of Proposed Candidate TPS for Access- Body Drag Reduction Using Forebody hood – 28 Surface Roughness – 85 to-Space Vehicles – 118 Earth-to-Orbit Rocket Propulsion – 17 DURABILITY ENERGY CONVERSION EFFICIENCY Estimate of avoidance maneuver rate for Fuel Cell Activities at the NASA Glenn Ceramic Matrix Composites (CMC) Life HASTOL tether boost facility – 28 Prediction Method Development – 88 Research Center – 49 Durability of Membrane Electrode As- ECONOMICS ENGINE AIRFRAME INTEGRATION semblies (MEAs) in PEM Fuel Cells Op- 2nd Generation Reusable Launch Ve- Affordable Flight Demonstration of the erated on Pure Hydrogen and Oxy- hicle Potential Commercial Development GTX Air-Breathing SSTO Vehicle Con- gen – 6 Scenarios – 79 cept – 46

A-13 ENGINE CONTROL Design and Testing of Non-Toxic RCS EXPERIMENT DESIGN The Control System for the X-33 Linear Thrusters for Second Generation Reus- The 1994 NASA/USRA/ADP Design Aerospike Engine – 105 able Launch Vehicle – 14 Projects – 140 ENGINE DESIGN Engineering the Future of Full-Scale Pro- EXPERT SYSTEMS Advanced Low-Cost O2/H2 Engines for pulsion Testing – 7 Advanced Propulsion System Test and the SSTO Application – 123 Hot dynamic test rig for measuring hyper- Tell Capability – 133 Advances in hypersonic vehicle synthe- sonic engine seal flow and durabil- Real-Time Sensor Validation, Signal Re- sis with application to studies of ad- ity – 151 construction, and Feature Detection for vanced thermal protection sys- Improved Testing Capability and Adapt- an RLV Propulsion Testbed – 98 tem – 138 ability Through the Use of Wireless Sen- Development and Capabilities of Unique sors – 2 F-18 AIRCRAFT Structural Seal Test Rigs – 41 Liquid Engine Test Facilities Assess- The F-18 systems research aircraft facil- Dual-Fuel Propulsion in Single-Stage Ad- ment – 1 ity – 163 vanced Manned Launch System Ve- Options for flight testing rocket-based FABRICATION hicle – 125 combined-cycle (RBCC) engines – 130 Advanced Metallic Thermal Protection High temperature NASP engine seals: A Status of the RBCC Direct-Connect System Development – 64 technology review – 173 Mixer Combustor Experiment – 54 Characterization of Polyimide Foams for Hot dynamic test rig for measuring hyper- Test Report for NASA MSFC Support of Ultra-Lightweight Space Struc- sonic engine seal flow and durabil- the Linear Aerospike SR-71 Experiment tures – 22 ity – 151 (LASRE) – 89 Composite,Cryogenic, Conformal, Com- Lightweight Chambers for Thrust Assem- ENTRY GUIDANCE (STS) mon Bulkhead, Aerogel-Insulated Tank blies – 66 Advanced Guidance and Control Meth- (CBAT) – 74 Main Engine Prototype Development for ods for Reusable Launch Vehicles: Test 2nd Generation RLV RS-83 – 43 Results – 30 Evaluation and Characterization Study of Dual Pulse Laser-Induced Spark (DPLIS) Multidisciplinary Approach to Aerospike ENVIRONMENT EFFECTS for Rocket Engine Ignition System Appli- Nozzle Design – 120 Lightweight durable TPS: Milestone cation – 14 4–151 NASA Overview – 35 Fabrication and Testing of Ceramic Ma- ENVIRONMENT POLLUTION Optimal trajectories for an aerospace trix Composite Rocket Propulsion Com- plane. Part 2: Data, tables, and An Overview of In-Stu Treatability Stud- ponents – 75 graphs – 176 ies at Marshall Space Flight Center, Huntsville, Alabama – 78 Heat pipes for wing leading edges of Unic Unstructured CFD Methodology De- hypersonic vehicles – 178 velopment – 82 EPOXY MATRIX COMPOSITES Permeability of Impacted Coated Com- Hypervelocity Impact Test Results for a ENGINE FAILURE posite Laminates – 49 Metallic Thermal Protection Sys- Abort performance for a winged-body tem – 10 single-stage to orbit vehicle – 137 EPOXY RESINS Damage tolerance of candidate thermo- Large Composite Structures Processing ENGINE INLETS set composites for use on single stage to Technologies for Reusable Launch Ve- Computational analysis in support of the orbit vehicles – 148 hicles – 75 SSTO flowpath test – 143 Impact Characteristics of Candidate Ma- Lightweight Chambers for Thrust Assem- Computational parametric study of terials for Single-Stage-to-Orbit (SSTO) blies – 82 sidewall-compression scramjet inlet per- Technology – 123 formance at Mach 10 – 161 Mosad and Stream Vision For A Telero- Study Acoustic Emissions from Compos- botic, Flying Camera System – 59 Numerical investigations in three- ites – 111 dimensional internal flows – 171 Selection process for trade study: Graph- EQUATIONS OF MOTION ite Composite Primary Structure Space Shuttle Main Engine Start with A parametric sensitivity study for single- (GCPS) – 148 Off-Nominal Propellant Inlet Pres- stage-to-orbit hypersonic vehicles using Test Plans. Lightweight Durable TPS: sures – 116 trajectory optimization – 153 Tasks 1,2,4,5, and 6 – 146 ENGINE MONITORING INSTRUMENTS Near-Optimal Re-Entry Trajectories for Improved Testing Capability and Adapt- Reusable Launch Vehicles – 115 Water Flow Performance of a Superscale ability Through the Use of Wireless Sen- Model of the Fastrac Liquid Oxygen sors – 2 ETHYL ALCOHOL Pump – 55 Design and Testing of Non-Toxic RCS ENGINE PARTS Thrusters for Second Generation Reus- FAILURE ANALYSIS Fabrication and Testing of Ceramic Ma- able Launch Vehicle – 18 Advanced Propulsion System Test and trix Composite Rocket Propulsion Com- Tell Capability – 133 ponents – 75 EVALUATION Multidisciplinary Optimization Branch Ex- X-33 LH2 Tank Failure Investigation High temperature NASP engine seal de- perience Using iSIGHT Software – 91 Findings – 58 velopment – 153 EVOLUTION (DEVELOPMENT) FAILURE MODES Rocket-Based Combined-Cycle (RBCC) International Space Station Evolution Propulsion Technology Workshop. Tuto- Data Book – 87 A Plan for Advanced Guidance and Con- rial session – 168 trol Technology for 2nd Generation Re- EXHAUST FLOW SIMULATION ENGINE TESTS usable Launch Vehicles – 52 NASA EPSCoR Preparation Grant – 49 Air-breathing aerospace plane develop- Rocket Engine Health Management: ment essential: Hypersonic propulsion EXPENDABLE STAGES (SPACECRAFT) Early Definition of Critical Flight Mea- flight tests – 138 Access to space study – 153 surements – 19

A-14 FAILURE FEEDBACK CONTROL FLIGHT CHARACTERISTICS Direct Fault Tolerant RLV Altitude Con- Air-breathing hypersonic vehicle guid- Assessment of flying-quality criteria for trol: A Singular Perturbation Ap- ance and control studies; An integrated air-breathing aerospacecraft – 167 proach – 40 trajectory/control analysis methodology: Phase 1 – 171 Infrared Sensing Aeroheating Flight Ex- Findings from the X-33 Hydrogen Tank periement: STS-96 Flight Results – 84 Failure Investigation – 60 Air-breathing hypersonic vehicle guidance and control studies: An integrated System Identification of X-33 Neural Net- Full Envelope Reconfigurable Control work – 8 Design for the X-33 Vehicle – 72 trajectory/control analysis methodology, phase 2 – 165 X-33 Tank Failure During Autoclave Fab- FLIGHT CONDITIONS rication – 60 Dynamics of aerospace vehicles – 174 High Heat Flux Facility – 151 FALSE ALARMS FERRY SPACECRAFT Stratospheric turbulence measurements Rocket Engine Health Management: Personnel launch system autoland de- and models for aerospace plane de- Early Definition of Critical Flight Mea- velopment study – 175 sign – 164 surements – 19 FIBER COMPOSITES FLIGHT CONTROL FASTENERS Ceramic Matrix Composites (CMC) Life Lightweight durable TPS: Milestone Prediction Method Development – 88 A Plan for Advanced Guidance and Con- 4–151 trol Technology for 2nd Generation Re- Permeability of Impacted Coated Com- usable Launch Vehicles – 52 Parametric Weight Comparison of Ad- posite Laminates – 49 vanced Metallic, Ceramic Tile, and Ce- Advanced aeroservoelastic stabilization ramic Blanket Thermal Protection Sys- FIBER OPTICS techniques for hypersonic flight ve- tems – 88 Reusable Cryogenic Tank VHM Using hicles – 163 Fiber Optic Distributed Sensing Technol- Advanced Guidance and Control Meth- FATIGUE (MATERIALS) ogy – 108 Fatigue Crack Growth Rate Test Results ods for Reusable Launch Vehicles: Test for Al-Li 2195 Parent Metal, Variable Po- FIBER-MATRIX INTERFACES Results – 30 larity Plasma Arc Welds and Friction Stir Characterization of C/SiC Ceramic Ma- An Automated Method to Compute Or- Welds – 89 trix Composites (CMCs) with Novel Inter- bital Re-Entry Trajectories with Heating Materials Research for High Speed Civil face Fiber Coatings – 48 Constraints – 53 Transport and Generic Hypersonics- FINITE ELEMENT METHOD Assessment of flying-quality criteria for Metals Durability – 127 A Strategy for Integrating a Large Finite air-breathing aerospacecraft – 167 FATIGUE TESTS Element Model Using MSC Development of the X-33 Aerodynamic Milestone 4: Test plan for Reusable Hy- NASTRAN/PATRAN: X-33 Lessons Uncertainty Model – 108 drogen Composite Tank System Learned – 97 Full Envelope Reconfigurable Control (RHCTS). Task 3: Composite tank mate- A study of facilities and fixtures for testing 142 Design for the X-33 Vehicle – 72 rials – of a high speed civil transport wing com- FAULT DETECTION ponent – 127 Improved Re-Configurable Sliding Mode Controller for Reusable Launch Vehicle Detection of Micro-Leaks Through Com- Computational Study of a McDonnell of Second Generation Addressing Aero- plex Geometries Under Mechanical Load Douglas Single-Stage-to-Orbit Vehicle dynamic Surface Failures and Thrust De- and at Cryogenic Temperature – 79 Concept for Aerodynamic Analy- ficiencies – 34 NASA IVHM Technology Experiment for sis – 126 67 Reconfigurable Control Design for the X-vehicles (NITEX) – Multidisciplinary Approach to Aerospike Full X-33 Flight Envelope – 72 Rocket Engine Health Management: Nozzle Design – 120 Early Definition of Critical Flight Mea- Structural arrangement trade study. Vol- Reconfigurable Flight Control Designs surements – 19 ume 3: Reusable Hydrogen Composite With Application to the X-33 Ve- FAULT TOLERANCE Tank System (RHCTS) and Graphite hicle – 95 Direct Fault Tolerant RLV Altitude Con- Composite Primary Structures (GCPS). Robust intelligent flight control for hyper- trol: A Singular Perturbation Ap- Addendum – 141 sonic vehicles – 166 proach – 40 FINITE VOLUME METHOD Support to X-33/Resusable Launch Ve- Reusable Launch Vehicle Attitude Con- Computational Aerodynamics Analysis of hicle Technology Program – 86 trol Using a Time-Varying Sliding Mode Future Launch Vehicle Configura- System Identification of X-33 Neural Net- Control Technique – 40 122 tions – work – 8 Test plan. GCPS task 7, subtask 7.1: IHM Hypersonic Boundary-Layer Transition Trim drag reduction concepts for horizon- development – 149 for X-33 Phase 2 Vehicle – 111 tal takeoff single-stage-to-Orbit ve- FEASIBILITY FINS hicles – 170 Manufacturing Process Simulation of Subsonic Aerodynamic Characteristics of Large-Scale Cryotanks – 36 Two Reconfigurable Flight-Control De- a Circular Body Earth-to-Orbit Ve- sign Methods: Robust Servomechanism FEED SYSTEMS hicle – 128 and Control Allocation – 76 Launch Vehicle Sizing Benefits Utilizing FLAME SPECTROSCOPY Main Propulsion System Crossfeed and FLIGHT ENVELOPES Project Status – 34 NASA EPSCoR Preparation Grant – 49 Reconfigurable Control Design for the Propellant Feed System Leak Detection: FLEXIBLE SPACECRAFT Full X-33 Flight Envelope – 72 Lessons Learned From the Linear Aero- Advanced aeroservoelastic stabilization spike SR-71 Experiment (LASRE) – 93 techniques for hypersonic flight ve- FLIGHT OPERATIONS Various advanced design projects pro- hicles – 163 Single-stage-to-orbit: Meeting the chal- moting engineering education – 144 Dynamics of aerospace vehicles – 174 lenge – 136

A-15 Space Ops 2002: Bringing Space Opera- Propellant Feed System Leak Detection: FLUID MANAGEMENT tions into the 21st Century. Track 3: Op- Lessons Learned From the Linear Aero- Reusable Cryogenic Tank VHM Using erations, Mission Planning and Control. spike SR-71 Experiment (LASRE) – 93 Fiber Optic Distributed Sensing Technol- 2nd Generation Reusable Launch Test Results for an Off-the-Shelf ogy – 108 Vehicle-Concepts for Flight Opera- GPS/INS During Approach and Landing FLUID MECHANICS tions – 35 Testing of the X-40A – 63 The future challenge for aeropropul- Systems Engineering Approach to Tech- The X-33 Extended Flight Test sion – 167 nology Integration for NASA’s 2nd Gen- Range – 99 eration Reusable Launch Vehicle – 23 FLUID-SOLID INTERACTIONS Wingless Flight: The Lifting Body Development of an integrated BEM ap- FLIGHT PATHS Story – 105 proach for hot fluid structure interac- Air-breathing hypersonic vehicle guid- tion – 169 FLOOD PREDICTIONS ance and control studies; An integrated FLUTTER ANALYSIS NASA EPSCoR Preparation Grant – 49 trajectory/control analysis methodology: A method of predicting quasi-steady Phase 1 – 171 FLOW CHARACTERISTICS aerodynamics for flutter analysis of high Near-Optimal Operation of Dual-Fuel Overview of GRC’s Advanced Sensor speed vehicles using steady CFD calcu- Launch Vehicles – 119 and Instrumentation Development – 41 lations – 157 Panel Flutter and Sonic Fatigue Analysis FLIGHT SIMULATION Slush hydrogen propellant production, for RLV – 77 Cyclic Cryogenic Thermal-Mechanical transfer, and expulsion studies at the Testing of an X-33/RLV Liquid Oxygen NASA K-Site Facility – 173 FLUTTER Evaluation of Computational Aeroelastic- Tank Concept – 94 FLOW DISTRIBUTION ity Codes for Loads and Flutter – 82 Thermal Protection System Evaluation Addition of equilibrium air to an upwind Using Arc-jet Flows: Flight Simulation or Navier-Stokes code and other first steps FOAMS Research Tool? – 31 toward a more generalized flow Characterization of Polyimide Foams for solver – 174 Ultra-Lightweight Space Struc- FLIGHT SIMULATORS tures – 22 An Experimental Study of Characteristic Orbital Space Plane Program Flight FOREBODIES Demonstrators Status – 6 Combustion-Driven Flow for CFD Valida- tion – 103 Flush Airdata Sensing (FADS) System Calibration Procedures and Results for FLIGHT TESTS Computation of H2/air reacting flowfields Blunt Forebodies – 92 A Base Drag Reduction Experiment on in drag-reduction external combus- the X-33 Linear Aerospike SR-71 Experi- tion – 161 Reduction of Base Drag on Launch Ve- ment (LASRE) Flight Program – 97 hicles – 50 Hypersonic Boundary-Layer Transition Advanced Guidance and Control Meth- for X-33 Phase 2 Vehicle – 111 FRACTALS ods for Reusable Launch Vehicles: Test High Altitude Launch for a Practical Hypersonic flows as related to the Na- Results – 30 SSTO – 32 tional Aerospace Plane – 136 Affordable Flight Demonstration of the FRACTURE STRENGTH Hypersonic flows as related to the Na- GTX Air-Breathing SSTO Vehicle Con- Fractographic Analysis of HfB2-SiC and tional Aerospace plane – 172 cept – 26 ZrB2-SiC Composites – 57 Air-breathing aerospace plane develop- Lee waves: Benign and malig- Materials Research for High Speed Civil 157 ment essential: Hypersonic propulsion nant – Transport and Generic Hypersonics- flight tests – 138 Thrust Enhancement in Hypervelocity Metals Durability – 127 Design, Development,and Testing of Um- Nozzles by Chemical Catalysis – 102 FUEL CELL POWER PLANTS billical System Mechanisms for the X-33 X-33 Computational Performance Evaluation of Electro- Advanced Technology Demonstra- Aeroheating/Aerodynamic Predictions chem’s PEM Fuel Cell Power Plant for tor – 90 and Comparisons With Experimental NASA’s 2nd Generation Reusable Data – 25 Launch Vehicle – 6 Flight Demonstration of X-33 Vehicle Health Management System Compo- FLOW MEASUREMENT FUEL CELLS nents on the F/A-18 Systems Research Hypersonic flows as related to the Na- Fuel Cell Activities at the NASA Glenn Aircraft – 15 tional Aerospace Plane – 136 Research Center – 49 Flight Stability and Control and Perfor- NASA EPSCoR Preparation Grant – 49 Performance Evaluation of Electro- mance Results from the Linear Aerospike chem’s PEM Fuel Cell Power Plant for SR-71 Experiment (LASRE) – 101 FLOW VISUALIZATION NASA’s 2nd Generation Reusable Launch Vehicle – 6 Flight Testing the Linear Aerospike A Novel Approach for Modeling Chemical SR-71 Experiment (LASRE) – 100 Reaction in Generalized Fluid System FUEL CONSUMPTION Simulation Program – 13 Hyper-X Wind Tunnel Program – 109 Effect of Aeropropulsive Interactions and Hypersonic Boundary-Layer Transition Design Sensitivities on Optimal Hyper- Linear Aerospike SR-71 Experiment for X-33 Phase 2 Vehicle – 111 sonic Ascent Trajectories – 125 (LASRE): Aerospace Propulsion Hazard Trajectory optimization for the national Mitigation Systems – 102 Subsonic Aerodynamic Characteristics of a Circular Body Earth-to-Orbit Ve- aerospace plane – 156 NASA’s Hypersonic Investment hicle – 128 Trim drag reduction concepts for horizon- Area – 47 tal takeoff single-stage-to-Orbit ve- FLUID FLOW Options for flight testing rocket-based hicles – 170 combined-cycle (RBCC) engines – 130 Prediction of the ullage gas thermal stratification in a NASP vehicle propellant FUEL PRODUCTION Orbital Space Plane Program Flight tank experimental simulation using Slush hydrogen transfer studies at the Demonstrators Status – 6 FLOW-3D – 177 NASA K-Site Test Facility – 167

A-16 FUEL PUMPS Propellant Feed System Leak Detection: Structural arrangement trade study. Vol- Unshrouded Impeller Technology Task Lessons Learned From the Linear Aero- ume 3: Reusable Hydrogen Composite Status – 81 spike SR-71 Experiment (LASRE) – 93 Tank System (RHCTS) and Graphite Composite Primary Structures (GCPS). FUEL TANKS GAS TURBINE ENGINES Addendum – 141 Acoustic Emission Monitoring of the RLV Turbine Performance Optimiza- DC-XA Composite Liquid Hydrogen Tank tion – 83 Trade study plan for Graphite Composite 126 Primary Structure (GCPS) – 146 During Structural Testing – Thrust augmentation options for the Beta Analytical comparison of three stiffened 2 two-stage-to-orbit vehicle – 154 GRAVITY WAVES panel concepts – 135 Lee waves: Benign and malig- GASEOUS ROCKET PROPELLANTS nant – 157 Development of a Pressure Box to Evalu- Experimental Evaluation of a Subscale ate Reusable-Launch-Vehicle Gaseous Hydrogen/Gaseous Oxygen GRAY GAS Cryogenic-Tank Panels – 132 Coaxial Rocket Injector – 34 Analysis of Aerospike Plume Induced Manufacturing Process Simulation of Base-Heating Environment – 109 GAS-GAS INTERACTIONS Large-Scale Cryotanks – 36 GRID GENERATION (MATHEMATICS) An Experimental Study of Characteristic Reusable Cryogenic Tank VHM Using Combustion-Driven Flow for CFD Valida- An implementation of a chemical and Fiber Optic Distributed Sensing Technol- tion – 103 thermal nonequilibrium flow solver on ogy – 108 unstructured meshes and application to GELLED ROCKET PROPELLANTS blunt bodies – 145 Support to X-33/Reusable Launch Ve- Preliminary Assessment of Using Gelled hicle Technology Program – 86 PNS predictions of external/internal hy- and Hybrid Propellant Propulsion for personic flows for NASP propulsion ap- Test plan. Task 5, subtask 5.2: Early VTOL/SSTO Launch Systems – 104 plications – 157 on-orbit TPSdebris impact tests – 148 GENERAL OVERVIEWS GROUND BASED CONTROL Trade study plan for Graphite Composite Overview of GRC’s Advanced Sensor NASA IVHM Technology Experiment for Primary Structure (GCPS) – 146 and Instrumentation Development – 41 X-vehicles (NITEX) – 67 Various advanced design projects pro- GENETIC ALGORITHMS moting engineering education – 144 GROUND OPERATIONAL SUPPORT Constrained minimization of smooth SYSTEM FUEL TESTS functions using a genetic algo- Design of launch systems using continu- A summary of the slush hydrogen tech- rithm – 141 ous improvement process – 134 nology program for the National Aero- Genetic Algorithm Tuned Fuzzy Logic for Systems Engineering Approach to Tech- Space Plane – 140 Gliding Return Trajectories – 12 nology Integration for NASA’s 2nd Gen- FULL SCALE TESTS eration Reusable Launch Vehicle – 23 GEOSYNCHRONOUS ORBITS Air-breathing aerospace plane develop- Advanced Electric Propulsion for RLV GROUND SUPPORT SYSTEMS ment essential: Hypersonic propulsion High Altitude Launch for a Practical flight tests – 138 Launched Geosynchronous Space- craft – 90 SSTO – 32 Engineering the Future of Full-Scale Pro- Lockheed Martin Response to the OSP pulsion Testing – 7 Affordable In-Space Transportation Phase 2: An Advanced Concepts Challenge – 20 Selection process for trade study: Graph- Project – 113 X-33 Integrated Test Facility Extended ite Composite Primary Structure Range Simulation – 100 (GCPS) – 148 Tether Transportation System Study – 109 GUIDANCE (MOTION) Test Plan. GCPS Task 4, subtask 4.2 thrust structure development – 147 GOVERNMENT/INDUSTRY RELATIONS Near-Optimal Operation of Dual-Fuel Launch Vehicles – 119 FUNCTIONAL DESIGN SPECIFICATIONS Test Plans. Lightweight Durable TPS: Tasks 1,2,4,5, and 6 – 146 Support to X-33/Resusable Launch Ve- The 1994 NASA/USRA/ADP Design hicle Technology Program – 86 Projects – 140 The 1994 NASA/USRA/ADP Design Projects – 140 GUIDANCE SENSORS FUSELAGES An Advanced Video Sensor for Auto- GRAPHITE-EPOXY COMPOSITES An overview of aeroelasticity studies for mated Docking – 61 the National Aerospace Plane – 160 Design and Analysis of a Two-Stage-to- Orbit Airbreathing Hypersonic Vehicle GUST LOADS FUZZY SYSTEMS Concept – 129 A Compendium of Wind Statistics and Fuzzy Logic Trajectory Design and Guid- Models for the NASA Space Shuttle and Evaluation of Microcracking in Two ance for Terminal Area Energy Manage- Other Aerospace Vehicle Pro- Carbon-Fiber/Epoxy-Matrix Composite ment – 8 grams – 99 Cryogenic Tanks – 72 Genetic Algorithm Tuned Fuzzy Logic for GUSTS Gliding Return Trajectories – 12 Milestone 4: Thrust structure concepts and IHM screening Graphite Composite Stratospheric turbulence measurements GAS CHROMATOGRAPHY Primary Structure (GCPS) – 143 and models for aerospace plane de- Director’s Discretionary Fund Report for sign – 164 GRAPHITE Fiscal Year 1996 – 120 HARDWARE Selection process for trade study: Graph- GAS DETECTORS ite Composite Primary Structure A Hazardous Gas Detection System for A Hazardous Gas Detection System for (GCPS) – 148 Aerospace and Commercial Applica- Aerospace and Commercial Applica- tions – 98 Structural Arrangement Trade Study. Vol- tions – 98 ume 1: Reusable Hydrogen Composite HARNESSES Evaluation of the Linear Aerospike SR-71 Tank System (RHCTS) and Graphite A Damage Resistance Comparison Be- Experiment (LASRE) Oxygen Sen- Composite Primary Structures (GCPS). tween Candidate Polymer Matrix Com- sor – 91 Executive summary – 141 posite Feedline Materials – 88

A-17 HAZARDS Advanced Transportation System Stud- HONEYCOMB CORES A Hazardous Gas Detection System for ies Technical Area 2 (TA-2) Heavy Lift Bonding and Sealing Evaluations for Aerospace and Commercial Applica- Launch Vehicle Development Con- Cryogenic Tanks – 116 tions – 98 tract – 129 HONEYCOMB STRUCTURES HEART DISEASES Advanced transportation system studies Development of Metallic Thermal Protec- Shuttle to Space Station. Heart assist technical area 2(TA-2): Heavy lift launch tion Systems for the Reusable Launch implant. Hubble update. X-30 mock- vehicle development – 130 Vehicle – 124 up – 138 Project Antares: A low cost modular HORIZONTAL FLIGHT launch vehicle for the future – 168 HEAT EXCHANGERS Trim drag reduction concepts for horizon- Task 4 supporting technology. Densifica- HETEROGENEITY tal takeoff single-stage-to-Orbit ve- tion requirements definition and test ob- Compute as Fast as the Engineers Can hicles – 170 jectives. Propellant densification require- Think! ULTRAFAST COMPUTING TEAM HORIZONTAL SPACECRAFT LANDING ments definition – 139 FINAL REPORT – 90 Requirements report for SSTO vertical Thermal/Fluid Analysis of a Composite take-off/horizontal landing ve- Heat Exchanger for Use on the RLV HIGH ALTITUDE hicle – 149 Rocket Engine – 37 High Altitude Launch for a Practical SSTO – 29 Selection process for trade study: Reus- HEAT FLUX Towers for Earth Launch – 29 able Hydrogen Composite Tank System Heat flux sensor research and develop- (RHCTS) – 148 ment: The cool film calorimeter – 170 HIGH ASPECT RATIO STS-102 Photo-op/Suit-up/Depart High Heat Flux Facility – 151 Comparison of High Aspect Ratio Cool- O&C/Launch Discovery On Thermal Protection Materials for Reentry ing Channel Designs for a Rocket Com- Orbit/Landing/Crew Egress – 27 bustion Chamber – 118 and Planetary Applications – 68 The design and evolution of the beta X-33 XRS-2200 Linear Aerospike Engine HIGH SPEED two-stage-to-orbit horizontal takeoff and Sea Level Plume Radiation – 66 The 1995 NASA High-Speed Research landing launch system – 159 Program Sonic Boom Workshop – 126 HEAT PIPES HUBBLE SPACE TELESCOPE Heat pipes for wing leading edges of HIGH STRENGTH Shuttle to Space Station. Heart assist hypersonic vehicles – 178 Characterization of C/SiC Ceramic Ma- implant. Hubble update. X-30 mock- HEAT RESISTANT ALLOYS trix Composites (CMCs) with Novel Inter- up – 138 Development of Metallic Thermal Protec- face Fiber Coatings – 48 HUMAN FACTORS ENGINEERING tion Systems for the Reusable Launch HIGH TEMPERATURE TESTS Human Transportation System (HTS) Vehicle – 124 Task 4 supporting technology. Part 2: study, volume 2 – 155 Thermoviscoplastic response of Ti-15-3 Detailed test plan for thermal seals. Ther- HYBRID PROPULSION under various loading conditions – 172 mal seals evaluation, improvement and Preliminary Assessment of Using Gelled HEAT SHIELDING test. CAN8-1, Reusable Launch Vehicle and Hybrid Propellant Propulsion for (RLV), advanced technology demonstra- Materials Testing on the DC-X and DC- VTOL/SSTO Launch Systems – 104 tor: X-33. Leading edge and seals ther- XA – 121 mal protection system technology dem- HYBRID ROCKET ENGINES HEAT TRANSFER onstration – 139 Various advanced design projects pro- Analysis of Aerospike Plume Induced moting engineering education – 144 Base-Heating Environment – 109 HIGH TEMPERATURE Design and Analysis of UHTC Leading HYDROCARBON FUELS Design of Experiments for the Thermal Edge Attachment – 44 Configuration development study of the Characterization of Metallic Foam – 1 OSU 1 hypersonic research ve- High temperature NASP engine seal de- hicle – 154 Hypersonic flows as related to the Na- velopment – 153 tional Aerospace plane – 172 Systems Challenges for Hypersonic Ve- High temperature NASP engine seals: A Prediction of the ullage gas thermal hicles – 117 technology review – 173 stratification in a NASP vehicle propellant HYDROGEN FUELS tank experimental simulation using High Temperature Polyimide Materials in FLOW-3D – 177 Extreme Temperature Environ- A summary of the slush hydrogen tech- ments – 73 nology program for the National Aero- Unic Unstructured CFD Methodology De- Space Plane – 140 velopment – 82 Hot dynamic test rig for measuring hyper- Characterization of Thin Film Polymers X-33 (Rev-F) Aeroheating Results of Test sonic engine seal flow and durability – 151 Through Dynamic Mechanical Analysis 6770 in NASA Langley 20-Inch Mach 6 and Permeation – 22 Air Tunnel – 96 HIGH VOLTAGES Computation of H2/air reacting flowfields HEATING High Voltage Design Guidelines: A in drag-reduction external combus- Infrared Sensing Aeroheating Flight Ex- Timely Update – 61 tion – 161 periement: STS-96 Flight Results – 84 H-INFINITY CONTROL Configuration development study of the Near-Optimal Re-Entry Trajectories for Hypersonic vehicle control law develop- OSU 1 hypersonic research ve- Reusable Launch Vehicles – 115 ment using H infinity and mu- hicle – 154 synthesis – 164 Thermal Protection Materials for Reentry Design and Testing of Non-Toxic RCS and Planetary Applications – 68 HISTORIES Thrusters for Second Generation Reus- able Launch Vehicle – 18 HEAVY LIFT LAUNCH VEHICLES Winds of change: Expanding the frontiers Advanced transportation system studies. of flight. Langley Research Center’s 75 Selection process for trade study: Reus- Technical area 2: Heavy lift launch ve- years of accomplishment, 1917- able Hydrogen Composite Tank System hicle development – 129 1992 – 165 (RHCTS) – 148

A-18 Slush hydrogen transfer studies at the Materials Research for High Speed Civil HYPERSONIC SPEED NASA K-Site Test Facility – 167 Transport and Generic Hypersonics- Air-breathing aerospace plane develop- 127 Structural Arrangement Trade Study. Vol- Metals Durability – ment essential: Hypersonic propulsion flight tests – 138 ume 1: Reusable Hydrogen Composite HYPERSONIC BOUNDARY LAYER Tank System (RHCTS) and Graphite Computational/Experimental Aeroheat- Assessment of flying-quality criteria for Composite Primary Structures (GCPS). ing Predictions for X-33 Phase 2 Ve- air-breathing aerospacecraft – 167 Executive summary – 141 hicle – 110 Computational parametric study of Structural arrangement trade study. Vol- Infrared Sensing Aeroheating Flight Ex- sidewall-compression scramjet inlet per- ume 3: Reusable Hydrogen Composite periment: STS-96 Flight Results – 86 formance at Mach 10 – 161 Tank System (RHCTS) and Graphite Composite Primary Structures (GCPS). HYPERSONIC COMBUSTION Computational Study of a McDonnell Douglas Single-Stage-to-Orbit Vehicle Addendum – 141 SRM-Assisted Trajectory for the GTX Concept for Aerodynamic Analy- Reference Vehicle – 48 Systems Challenges for Hypersonic Ve- sis – 126 hicles – 117 HYPERSONIC FLIGHT Discrete Roughness Effects on Shuttle Technology issues associated with using Air-breathing aerospace plane develop- Orbiter at Mach 6 – 37 densified hydrogen for space ve- ment essential: Hypersonic propulsion hicles – 167 flight tests – 142 Hypersonic flows as related to the Na- tional Aerospace Plane – 136 Test plan. Task 5, subtask 5.2: Early Analytical solutions to constrained hyper- on-orbit TPSdebris impact tests – 148 sonic flight trajectories – 162 Optimal trajectories for an aerospace HYDROGEN OXYGEN ENGINES plane. Part 1: Formulation, results, and Hypersonic vehicle control law develop- analysis – 176 The design and evolution of the beta ment using H infinity and mu- two-stage-to-orbit horizontal takeoff and synthesis – 164 Thrust augmentation options for the Beta landing launch system – 159 2 two-stage-to-orbit vehicle – 154 Optimal trajectories for an aerospace HYDROGEN OXYGEN FUEL CELLS plane. Part 2: Data, tables, and X-33 (Rev-F) Aeroheating Results of Test Durability of Membrane Electrode As- graphs – 176 6770 in NASA Langley 20-Inch Mach 6 Air Tunnel – 96 semblies (MEAs) in PEM Fuel Cells Op- Rocket Based Combined Cycle (RBCC) erated on Pure Hydrogen and Oxy- Propulsion Workshop, volume 2 – 175 X-33 Rev-F Turbulent Aeroheating Re- gen – 6 sults From Test 6817 in NASA Langley Rocket-Based Combined-Cycle (RBCC) 20-Inch Mach 6 Air Tunnel and Compari- HYDROGEN PRODUCTION Propulsion Technology Workshop. Tuto- sons With Computations – 26 Slush hydrogen propellant production, rial session – 168 transfer, and expulsion studies at the HYPERSONIC VEHICLES NASA K-Site Facility – 173 Thrust Enhancement in Hypervelocity Nozzles by Chemical Catalysis – 102 A parametric sensitivity study for single- HYDROGEN Trajectory optimization for the National stage-to-orbit hypersonic vehicles using An Investigation of the Effect of Surface 153 aerospace plane – 160 trajectory optimization – Impurities on the Adsorption Kinetics of Hydrogen Chemisorbed onto Iron – 101 Turbulence modeling for hypersonic Advanced aeroservoelastic stabilization flight – 161 techniques for hypersonic flight ve- Experimental Evaluation of a Subscale hicles – 163 Gaseous Hydrogen/Gaseous Oxygen HYPERSONIC FLOW Coaxial Rocket Injector – 34 Advanced Range Safety System for High An implementation of a chemical and Energy Vehicles – 15 Selection and Evaluation of An Alloy for thermal nonequilibrium flow solver on Nozzle Application – 22 unstructured meshes and application to Advances in hypersonic vehicle synthe- blunt bodies – 145 sis with application to studies of ad- Technology issues associated with using vanced thermal protection sys- densified hydrogen for space ve- Computational Study of a McDonnell tem – 138 hicles – 167 Douglas Single-Stage-to-Orbit Vehicle Concept for Aerodynamic Analy- Aerothermoelastic analysis of a NASP HYDROLOGY MODELS sis – 126 demonstrator model – 155 NASA EPSCoR Preparation Grant – 49 Hot dynamic test rig for measuring hyper- Air-breathing hypersonic vehicle guid- HYDROSTATIC PRESSURE sonic engine seal flow and durabil- ance and control studies; An integrated Space Shuttle Main Engine Off-Nominal ity – 151 trajectory/control analysis methodology: 117 Phase 1 – 171 Low Power Level Operation – Hypersonic flows as related to the Na- HYPERGOLIC ROCKET PROPELLANTS tional Aerospace Plane – 136 Air-breathing hypersonic vehicle guidance and control studies: An integrated Preliminary Assessment of Using Gelled Hypersonic flows as related to the na- trajectory/control analysis methodology, and Hybrid Propellant Propulsion for tional aerospace plane – 162 VTOL/SSTO Launch Systems – 104 phase 2 – 165 Hypersonic flows as related to the Na- Calculation of Supersonic Combustion HYPERSONIC AIRCRAFT tional Aerospace plane – 172 Using Implicit Schemes – 30 A simple analytical aerodynamic model Numerical investigations in three- Computation of H2/air reacting flowfields of Langley Winged-Cone Aerospace dimensional internal flows – 171 Plane concept – 145 in drag-reduction external combus- PNS predictions of external/internal hy- tion – 161 Affordable Flight Demonstration of the personic flows for NASP propulsion ap- Computational/Experimental Aeroheat- GTX Air-Breathing SSTO Vehicle Con- plications – 157 cept – 26 ing Predictions for X-33 Phase 2 Ve- hicle – 110 GTX Reference Vehicle Structural Verifi- HYPERSONIC INLETS cation Methods and Weight Sum- Computational analysis in support of the Conceptual design of two-stage-to-orbit mary – 44 SSTO flowpath test – 143 hybrid launch vehicle – 169

A-19 Configuration development study of the Hot dynamic test rig for measuring hyper- Impact Characteristics of Candidate Ma- OSU 1 hypersonic research ve- sonic engine seal flow and durabil- terials for Single-Stage-to-Orbit (SSTO) hicle – 154 ity – 151 Technology – 123 Design and Analysis of a Two-Stage-to- Hypersonic Boundary-Layer Transition Milestone 4: Test plan for Reusable Hy- Orbit Airbreathing Hypersonic Vehicle for X-33 Phase 2 Vehicle – 111 drogen Composite Tank System 129 Concept – Integrated Technology Assessment Cen- (RHCTS). Task 3: Composite tank mate- 142 Estimate of avoidance maneuver rate for ter (ITAC) Update – 42 rials – HASTOL tether boost facility – 28 Permeability After Impact Testing of NASA’s Hypersonic Investment Composite Laminates – 21 Heat pipes for wing leading edges of Area – 47 hypersonic vehicles – 178 Permeability Testing of Impacted Com- Numerical investigations in three- posite Laminates for Use on Reusable Hypersonic vehicle control law develop- dimensional internal flows – 171 ment using H infinity and mu- Launch Vehicles – 80 synthesis – 164 Optimal trajectories for an aerospace Results of Two-Stage Light-Gas Gun De- plane. Part 1: Formulation, results, and velopment Efforts and Hypervelocity Im- Hypersonic vehicle model and control analysis – 176 law development using H(infinity) and pact Tests of Advanced Thermal Protec- 100 micron synthesis – 144 Structures and materials technology for tion Materials – hypersonic aerospacecraft – 178 Test plan. Task 5, subtask 5.2: Early Hyper-X Wind Tunnel Program – 109 Subsonic-to-Hypersonic Aerodynamic on-orbit TPSdebris impact tests – 148 Linear Aerospike SR-71 Experiment Characteristics for a Winged, Circular- IMPACT VELOCITY (LASRE): Aerospace Propulsion Hazard Body, Single-Stage-to-Orbit Spacecraft Mitigation Systems – 102 Results of Two-Stage Light-Gas Gun De- Configuration – 124 velopment Efforts and Hypervelocity Im- Minimum Heating Re-Entry Trajectories Systems Challenges for Hypersonic Ve- pact Tests of Advanced Thermal Protec- for Advanced Hypersonic Launch Ve- hicles – 117 tion Materials – 100 hicles – 119 The X-33 Extended Flight Test IMPELLERS Numerical investigations in three- Range – 99 Progress in Unsteady Turbopump Flow dimensional internal flows – 171 Simulations Using Overset Grid Sys- PNS predictions of external/internal hy- HYPERVELOCITY FLOW tems – 30 Thrust Enhancement in Hypervelocity personic flows for NASP propulsion ap- Unshrouded Impeller Technology Task Nozzles by Chemical Catalysis – 102 plications – 157 Status – 81 Rapid near-optimal trajectory generation HYPERVELOCITY IMPACT IMPURITIES and guidance law development for Development of Metallic Thermal Protec- An Investigation of the Effect of Surface single-stage-to-orbit airbreathing ve- tion Systems for the Reusable Launch Impurities on the Adsorption Kinetics of 177 hicles – Vehicle – 124 Hydrogen Chemisorbed onto Iron – 101 Robust intelligent flight control for hyper- Hypervelocity Impact Test Results for a INCOMPRESSIBLE FLOW sonic vehicles – 166 Metallic Thermal Protection Sys- Development of an integrated BEM ap- Structural dynamic and aeroelastic con- tem – 10 proach for hot fluid structure interac- siderations for hypersonic ve- Results of Two-Stage Light-Gas Gun De- tion – 169 hicles – 173 velopment Efforts and Hypervelocity Im- INCOMPRESSIBLE FLUIDS Supersonic aeroelastic instability results pact Tests of Advanced Thermal Protec- Lee waves: Benign and malig- for a NASP-like wing model – 159 tion Materials – 100 nant – 157 Systems Challenges for Hypersonic Ve- IGNITION SYSTEMS INFORMATION MANAGEMENT hicles – 117 Evaluation and Characterization Study of Automated CFD Parameter Studies on Dual Pulse Laser-Induced Spark (DPLIS) The Center for Aerospace Research: A Distributed Parallel Computers – 36 for Rocket Engine Ignition System Appli- NASA Center of Excellence at North INFRARED DETECTORS cation – 14 Carolina Agricultural and Technical State Infrared Sensing Aeroheating Flight Ex- University – 162 IMPACT DAMAGE periment: STS-96 Flight Results – 86 The design and evolution of the beta Damage tolerance of candidate thermo- INFRARED INSPECTION two-stage-to-orbit horizontal takeoff and set composites for use on single stage to Thermographic Analysis of Composite 148 landing launch system – 159 orbit vehicles – Cobonds on the X-33 – 75 Trim drag reduction concepts for horizon- IMPACT RESISTANCE INFRARED RADIATION tal takeoff single-stage-to-Orbit ve- Damage tolerance of candidate thermo- Infrared Sensing Aeroheating Flight Ex- hicles – 170 set composites for use on single stage to periement: STS-96 Flight Results – 84 orbit vehicles – 148 Wireless Subsurface Microsensors for INFRARED SPECTRA Health Monitoring of Thermal Protection Impact Characteristics of Candidate Ma- Development of Processing Techniques Systems on Hypersonic Vehicles – 69 terials for Single-Stage-to-Orbit (SSTO) for Advanced Thermal Protection Materi- Wireless Subsurface Sensors for Health Technology – 123 als – 112 Monitoring of Thermal Protection Sys- IMPACT STRENGTH INFRARED SPECTROSCOPY tems on Hypersonic Vehicles – 68 Impact Characteristics of Candidate Ma- Development of Processing Techniques HYPERSONICS terials for Single-Stage-to-Orbit (SSTO) for Advanced Thermal Protection Materi- An Approach to Establishing System Technology – 123 als – 112 Benefits for Technology in NASA’s Hy- IMPACT TESTS INJECTORS personics Investment Area – 67 Hypervelocity Impact Test Results for a An Experimental Study of Characteristic Development of the X-33 Aerodynamic Metallic Thermal Protection Sys- Combustion-Driven Flow for CFD Valida- Uncertainty Model – 108 tem – 10 tion – 103

A-20 CFD-Based Design Optimization for Computational Study of a McDonnell LASER DOPPLER VELOCIMETERS Single Element Rocket Injector – 25 Douglas Single-Stage-to-Orbit Vehicle An Experimental Study of Characteristic Concept for Aerodynamic Analy- Experimental Evaluation of a Subscale Combustion-Driven Flow for CFD Valida- sis – 126 Gaseous Hydrogen/Gaseous Oxygen tion – 103 Coaxial Rocket Injector – 34 N-Factor Computations for the X-33 Ve- LASER INDUCED FLUORESCENCE hicle – 76 Using CFD as Rocket Injector Design An Experimental Study of Characteristic Tool: Recent Progress at Marshall Space ION ENGINES Combustion-Driven Flow for CFD Valida- Flight Center – 3 Ion engine propelled Earth-Mars cycler tion – 103 with nuclear thermal propelled transfer LASER OUTPUTS INLET FLOW vehicle, volume 2 – 144 Evaluation and Characterization Study of Computational analysis in support of the Dual Pulse Laser-Induced Spark (DPLIS) SSTO flowpath test – 143 ION PROPULSION Ion engine propelled Earth-Mars cycler for Rocket Engine Ignition System Appli- Inlet Development for a Rocket Based with nuclear thermal propelled transfer cation – 14 Combined Cycle, Single Stage to Orbit vehicle, volume 2 – 144 LATERAL STABILITY Vehicle Using Computational Fluid Dy- Transonic aerodynamic characteristics of namics – 95 IRON An Investigation of the Effect of Surface a proposed wing-body reusable launch Numerical investigations in three- Impurities on the Adsorption Kinetics of vehicle concept – 138 dimensional internal flows – 171 Hydrogen Chemisorbed onto Iron – 101 LAUNCH VEHICLE CONFIGURATIONS INSTRUMENT PACKAGES ISOLATION Access to space – 143 Improved Testing Capability and Adapt- Parametric Weight Comparison of Ad- Advanced transportation system studies. ability Through the Use of Wireless Sen- vanced Metallic, Ceramic Tile, and Ce- Technical area 2: Heavy lift launch ve- sors – 2 ramic Blanket Thermal Protection Sys- hicle development – 129 tems – 88 INSULATION Advanced transportation system study: Cryopumping in Cryogenic Insulations JAPAN Manned launch vehicle concepts for two for a Reusable Launch Vehicle – 2 JTEC panel report on space and transat- way transportation system payloads to mospheric propulsion technology – 175 LEO – 152 INTAKE SYSTEMS JET FLOW Computational Aerodynamics Analysis of Thrust Enhancement in Hypervelocity Thermal Protection System Evaluation Future Launch Vehicle Configura- 102 Nozzles by Chemical Catalysis – Using Arc-jet Flows: Flight Simulation or tions – 122 INTERMETALLICS Research Tool? – 31 Mission and sizing analysis for the Beta 2 Hypervelocity Impact Test Results for a JET PROPULSION two-stage-to-orbit vehicle – 168 Metallic Thermal Protection Sys- JTEC panel report on space and transat- Overview of the Beta II Two-Stage-To- tem – 10 mospheric propulsion technology – 175 Orbit vehicle design – 172 INTERNATIONAL SPACE STATION JOINTS (JUNCTIONS) Reconfigurable Flight Control Designs A Collection of Technical Papers – 122 Development of Metal Matrix Compos- With Application to the X-33 Ve- ites for NASA’s Advanced Propulsion 95 International Space Station Evolution hicle – Systems – 74 Data Book – 87 Requirements report for SSTO vertical Thermographic Analysis of Composite take-off and horizontal landing ve- NASA Alternate Access to Station Ser- Cobonds on the X-33 – 75 hicle – 143 vice Concept – 65 X-33 Tank Failure During Autoclave Fab- The 1994 NASA/USRA/ADP Design Orbital Space Plane (OSP) Pro- rication – 64 Projects – 140 gram – 5 LAGRANGE MULTIPLIERS Trade study plan for Reusable Hydrogen Space Ops 2002: Bringing Space Opera- Constrained minimization of smooth Composite Tank System tions into the 21st Century. Track 3: Op- functions using a genetic algo- (RHCTS) – 145 erations, Mission Planning and Control. rithm – 141 2nd Generation Reusable Launch LAUNCH VEHICLES Vehicle-Concepts for Flight Opera- LAGRANGIAN FUNCTION A Procedure for Structural Weight Esti- tions – 35 Real-time trajectory optimization on par- mation of Single Stage to Orbit Launch 158 Vehicles (Interim User’s Manual) – 44 STS-102 Photo-op/Suit-up/Depart allel processors – O&C/Launch Discovery On LAMINAR FLOW Access to space study – 153 Orbit/Landing/Crew Egress – 27 Computational/Experimental Aeroheat- Advanced Range Safety System for High ing Predictions for X-33 – 112 INTERSTELLAR CHEMISTRY Energy Vehicles – 15 Director’s Discretionary Fund Report for LAMINATES Advanced transportation system studies Fiscal Year 1996 – 120 Impedance-Based Structural Health technical area 2 (TA-2): Heavy lift launch Monitoring for Composite Laminates at vehicle development – 130 INVARIANT IMBEDDINGS Cryogenic Environments – 7 Advanced transportation system studies An Improved RLV Stability Analysis Via a Milestone 4: Test plan for Reusable Hy- technical area 2(TA-2): Heavy lift launch Continuation Approach – 9 drogen Composite Tank System vehicle development – 130 (RHCTS). Task 3: Composite tank mate- INVISCID FLOW Advanced transportation system study: rials – 142 An Inviscid Computational Study of an Manned launch vehicle concepts for two X-33 Configuration at Hypersonic Permeability After Impact Testing of way transportation system payloads to Speeds – 94 Composite Laminates – 21 LEO – 152 Comparison of Methods for Determining Permeability Testing of Impacted Com- Comparison of Two Multidisciplinary Op- Boundary Layer Edge Conditions for posite Laminates for Use on Reusable timization Strategies for Launch-Vehicle Transition Correlations – 23 Launch Vehicles – 80 Design – 123

A-21 Computational Aerodynamics Analysis of X-33 LH2 Tank Failure Investigation LIFT Future Launch Vehicle Configura- Findings – 14 A simple analytical aerodynamic model tions – 122 LEE WAVES of Langley Winged-Cone Aerospace Plane concept – 145 Conceptual design of two-stage-to-orbit Lee waves: Benign and malig- hybrid launch vehicle – 169 nant – 157 LIGHT (VISIBLE RADIATION) Lockheed Martin Response to the OSP LIFE CYCLE COSTS Mosad and Stream Vision For A Telero- botic, Flying Camera System – 59 Challenge – 20 A Quantitative Reliability, Maintainability NASA Glenn Research Center’s Hyper- and Supportability Approach for NASA’s LIQUID FLOW sonic Propulsion Program – 95 Second Generation Reusable Launch Space Shuttle Main Engine Start with Vehicle – 41 Off-Nominal Propellant Inlet Pres- Orbital Space Plane Program Sta- sures – 116 tus – 5 Access to space – 143 LIQUID HYDROGEN Results of a Rocket-Based Combined- Advanced transportation system studies Acoustic Emission Monitoring of the Cycle SSTO Design Using Parametric technical area 2 (TA-2): Heavy lift launch DC-XA Composite Liquid Hydrogen Tank MDO Methods – 121 vehicle development – 130 During Structural Testing – 126 Rocket Based Combined Cycle (RBCC) Advanced transportation system studies Design and Testing of Non-Toxic RCS Propulsion Workshop, volume 2 – 166 technical area 2(TA-2): Heavy lift launch vehicle development – 130 Thrusters for Second Generation Reus- SRM-Assisted Trajectory for the GTX able Launch Vehicle – 18 Reference Vehicle – 48 Estimating the Cost of NASA’s Space Launch Initiative: How SLI Cost Stack Up Detection of Micro-Leaks Through Com- Various advanced design projects pro- Against the Shuttle – 51 plex Geometries Under Mechanical Load moting engineering education – 144 and at Cryogenic Temperature – 79 Life Cycle Systems Engineering Ap- Vision for CFD-Based Combustion Insta- proach to NASA’s 2nd Generation Reus- Evaluation of Microcracking in Two 23 bility Predictions – able Launch Vehicle – 39 Carbon-Fiber/Epoxy-Matrix Composite X-33 Phase 2 – 115 Cryogenic Tanks – 72 O/S analysis of conceptual space ve- LAUNCHING BASES hicles – 134 Findings from the X-33 Hydrogen Tank Failure Investigation – 60 Business Plan for the Southwest Re- STARS: The Space Transportation Archi- gional Spaceport: Executive Sum- tecture Risk System – 116 Milestone 4: Test plan for Reusable Hy- mary – 114 drogen Composite Tank System LIFE (DURABILITY) (RHCTS). Task 3: Composite tank mate- LAUNCHING A Quantitative Reliability, Maintainability rials – 142 Design of launch systems using continu- and Supportability Approach for NASA’s ous improvement process – 134 Performance Tests of a Liquid Hydrogen Second Generation Reusable Launch Propellant Densification Ground System 41 LEADING EDGE SWEEP Vehicle – for the X33/RLV – 118 Advanced Durability and Damage Toler- Computational parametric study of Preliminary Assessment of Using Gelled ance Design and Analysis Methods for sidewall-compression scramjet inlet per- and Hybrid Propellant Propulsion for 161 Composite Structures: Lessons Learned formance at Mach 10 – VTOL/SSTO Launch Systems – 104 from NASA Technology Development LEADING EDGES Programs – 26 Recent Advances and Applications in Fractographic Analysis of HfB2-SiC and Cryogenic Propellant Densification Tech- Advanced High Temperature Structural ZrB2-SiC Composites – 57 nology – 89 Seals – 43 Heat flux sensor research and develop- Slush hydrogen technology pro- Ceramic Matrix Composites (CMC) Life ment: The cool film calorimeter – 170 gram – 149 Prediction Method Development – 88 Heat pipes for wing leading edges of Slush hydrogen transfer studies at the hypersonic vehicles – 178 LIFT DRAG RATIO NASA K-Site Test Facility – 167 Subsonic Aerodynamic Characteristics of Task 4 supporting technology. Part 1: Task 4 supporting technology. Densifica- Detailed test plan for leading edge tile a Circular Body Earth-to-Orbit Ve- hicle – 128 tion requirements definition and test ob- development. Leading edge material de- jectives. Propellant densification require- velopment and testing – 139 LIFTING BODIES ments definition – 139 LEAKAGE Computational Aerodynamics Analysis of Technology issues associated with using Future Launch Vehicle Configura- Detection of Micro-Leaks Through Com- densified hydrogen for space ve- tions – 122 plex Geometries Under Mechanical Load hicles – 167 Development of the X-33 Aerodynamic and at Cryogenic Temperature – 79 Testing of Densified Liquid Hydrogen Uncertainty Model – 108 High temperature NASP engine seals: A Stratification in a Scale Model Propellant technology review – 173 Linear Aerospike SR-71 Experiment Tank – 67 (LASRE): Aerospace Propulsion Hazard Trade study plan for Reusable Hydrogen Hot dynamic test rig for measuring hyper- Mitigation Systems – 102 sonic engine seal flow and durabil- Composite Tank System ity – 151 Wingless Flight: The Lifting Body (RHCTS) – 145 Story – 105 Permeability After Impact Testing of X-33 LH2 Tank Failure Investigation Composite Laminates – 21 LIFTING REENTRY VEHICLES Findings – 14 Permeability Testing of Impacted Com- Thermal Management Design for the X-33 Tank Failure During Autoclave Fab- posite Laminates for Use on Reusable X-33 Lifting Body – 104 rication – 60 Launch Vehicles – 83 LIFTOFF (LAUNCHING) LIQUID NITROGEN Propellant Feed System Leak Detection: STS-102 Photo-op/Suit-up/Depart Evaluation of the Linear Aerospike SR-71 Lessons Learned From the Linear Aero- O&C/Launch Discovery On Experiment (LASRE) Oxygen Sen- spike SR-71 Experiment (LASRE) – 93 Orbit/Landing/Crew Egress – 27 sor – 91

A-22 LIQUID OXYGEN LOAD TESTING MACHINES Results of a Rocket-Based Combined- An Experimental Study of Characteristic Cyclic Cryogenic Thermal-Mechanical Cycle SSTO Design Using Parametric Combustion-Driven Flow for CFD Valida- Testing of an X-33/RLV Liquid Oxygen MDO Methods – 121 tion – 103 Tank Concept – 94 SRM-Assisted Trajectory for the GTX Design and Testing of Non-Toxic RCS LOAD TESTS Reference Vehicle – 47 Thrusters for Second Generation Reus- Cyclic Cryogenic Thermal-Mechanical Tether Transportation System able Launch Vehicle – 18 Testing of an X-33/RLV Liquid Oxygen Study – 109 Tank Concept – 94 Linear Aerospike SR-71 Experiment LUBRICATION (LASRE): Aerospace Propulsion Hazard Development of a Pressure Box to Evalu- High temperature NASP engine seals: A Mitigation Systems – 102 ate Reusable-Launch-Vehicle technology review – 173 Cryogenic-Tank Panels – 132 Preliminary Assessment of Using Gelled LUNAR BASES LOADS (FORCES) and Hybrid Propellant Propulsion for The rationale/benefits of nuclear thermal VTOL/SSTO Launch Systems – 104 A study of facilities and fixtures for testing rocket propulsion for NASA’s lunar space of a high speed civil transport wing com- transportation system – 141 Recent Advances and Applications in ponent – 127 Cryogenic Propellant Densification Tech- LUNAR LANDING nology – 89 Development of Metallic Thermal Protec- tion Systems for the Reusable Launch Human lunar mission capabilities using Space Shuttle Main Engine Start with Vehicle – 124 SSTO, ISRU and LOX-augmented NTR Off-Nominal Propellant Inlet Pres- technologies: A preliminary assess- sures – 116 Evaluation of Computational Aeroelastic- ment – 137 ity Codes for Loads and Flutter – 82 Task 4 supporting technology. Densifica- LUNAR ORBITAL RENDEZVOUS Parametric Weight Comparison of Ad- tion requirements definition and test ob- The rationale/benefits of nuclear thermal vanced Metallic, Ceramic Tile, and Ce- jectives. Propellant densification require- rocket propulsion for NASA’s lunar space ramic Blanket Thermal Protection Sys- ments definition – 139 transportation system – 141 tems – 88 Water Flow Performance of a Superscale LUNAR ORBITS LOGIC DESIGN Model of the Fastrac Liquid Oxygen Affordable In-Space Transportation Pump – 55 Fuzzy Logic Trajectory Design and Guid- Phase 2: An Advanced Concepts ance for Terminal Area Energy Manage- Project – 113 8 LIQUID PROPELLANT ROCKET EN- ment – LUNAR SURFACE GINES LOGISTICS MANAGEMENT Advanced spacecraft designs in support A Novel Approach for Modeling Chemical A Collection of Technical Papers – 122 of human missions to earth’s neighbor- Reaction in Generalized Fluid System LONGITUDINAL STABILITY hood – 28 Simulation Program – 13 Subsonic-to-Hypersonic Aerodynamic MACH NUMBER Advanced Low-Cost O2/H2 Engines for Characteristics for a Winged, Circular- A simple analytical aerodynamic model the SSTO Application – 123 Body, Single-Stage-to-Orbit Spacecraft of Langley Winged-Cone Aerospace Configuration – 124 Advanced Propulsion System Test and Plane concept – 145 Tell Capability – 133 Transonic aerodynamic characteristics of Aerodynamic Characteristics of a Vertical a proposed wing-body reusable launch Takeoff Vertical Landing (VTVL) Single Alternate Propulsion Subsystem Con- vehicle concept – 138 cepts Tripropellant Comparison Stage to Orbit Vehicle from M(sub infin- Study – 132 LOW ASPECT RATIO WINGS ity) = 0.1 to 10 – 122 Supersonic aeroelastic instability results Aerothermoelastic analysis of a NASP JTEC panel report on space and transat- for a NASP-like wing model – 159 mospheric propulsion technology – 175 demonstrator model – 155 LOW COST An implementation of a chemical and Liquid Engine Test Facilities Assess- Progress and Status of the 2nd Genera- thermal nonequilibrium flow solver on ment – 1 tion Reusable Launch Vehicle Pro- unstructured meshes and application to National Rocket Propulsion Materials gram – 76 blunt bodies – 145 Plan: A NASA, Department of Defense, LOW EARTH ORBITS Estimate of avoidance maneuver rate for and Industry Partnership – 73 A Cryogenic Propellant Production Depot HASTOL tether boost facility – 28 81 Propulsion module design review. Mile- for Low Earth Orbit – Near-Optimal Operation of Dual-Fuel stone 8: Integrated Propulsion Technol- Advanced transportation system studies Launch Vehicles – 119 ogy Demonstrator – 150 technical area 2(TA-2): Heavy lift launch Numerical investigations in three- 130 Test Report for NASA MSFC Support of vehicle development – dimensional internal flows – 171 the Linear Aerospike SR-71 Experiment Affordable In-Space Transportation Operationally efficient propulsion system (LASRE) – 89 Phase 2: An Advanced Concepts study (OEPSS) data book – 110 Project – 113 LIQUID ROCKET PROPELLANTS Transonic aerodynamic characteristics of General Public Space Travel and Tour- a proposed wing-body reusable launch 2nd Generation Reusable Launch Ve- ism – 96 hicle (2G RLV) – 82 vehicle concept – 138 Highly reusable space transportation: Design and Testing of Non-Toxic RCS Approaches for reducing ETO launch MAGNETOHYDRODYNAMIC GENERA- TORS Thrusters for Second Generation Reus- costs to $100 - $200 per pound of pay- able Launch Vehicle – 18 load – 135 Magnetohydrodynamic Augmented Pro- pulsion Experiment – 52 Task 4 supporting technology. Densifica- Human lunar mission capabilities using tion requirements definition and test ob- SSTO, ISRU and LOX-augmented NTR MAGNETOHYDRODYNAMICS jectives. Propellant densification require- technologies: A preliminary assess- Magnetohydrodynamic Augmented Pro- ments definition – 139 ment – 137 pulsion Experiment – 52

A-23 Rocket-Induced Magnetohydrodynamic Durability of Membrane Electrode As- Investigation of Fundamental Modeling Ejector: A Single-Stage-to-Orbit Ad- semblies (MEAs) in PEM Fuel Cells Op- and Thermal Performance Issues for a vanced Propulsion Concept – 131 erated on Pure Hydrogen and Oxy- Metallic Thermal Protection System De- gen – 6 sign – 33 MAINTAINABILITY A Quantitative Reliability, Maintainability Manufacturing Process Simulation of Launch Vehicle Sizing Beneﬁts Utilizing and Supportability Approach for NASA’s Large-Scale Cryotanks – 36 Main Propulsion System Crossfeed and 34 Second Generation Reusable Launch MARKET RESEARCH Project Status – Vehicle – 41 2nd Generation RLV: Program Goals and Optimal trajectories for an aerospace O/S analysis of conceptual space ve- Acquisition Strategy – 79 plane. Part 2: Data, tables, and graphs – 176 hicles – 134 Analysis of space concepts enabled by Reusable launch vehicle reliability, main- new transportation (ASCENT) Vision for CFD-Based Combustion Insta- tainability and operability assess- study – 28 bility Predictions – 23 ment – 150 Transportation Systems Analyses (TSA): MATRICES (MATHEMATICS) MAINTENANCE Highly Reusable Space Transportation Investigation of Fundamental Modeling Operations Analysis of the 2nd Genera- (HRST). A preliminary programmatic as- and Thermal Performance Issues for a 131 tion Reusable Launch Vehicle – 43 sessment – Metallic Thermal Protection System De- MASSIVELY PARALLEL PROCESSORS sign – 33 MANAGEMENT METHODS Compute as Fast as the Engineers Can MATRIX MATERIALS Second Generation RLV Space Vehicle Think! ULTRAFAST COMPUTING TEAM Concept – 65 FINAL REPORT – 90 Damage tolerance of candidate thermoset composites for use on single stage to MANAGEMENT PLANNING MATERIALS SCIENCE orbit vehicles – 148 Advanced spacecraft designs in support Research Reports: 1996 NASA/ASEE of human missions to earth’s neighbor- Summer Faculty Fellowship Pro- MEASURING INSTRUMENTS hood – 28 gram – 102 Development of Metallic Thermal Protec- tion Systems for the Reusable Launch MANAGEMENT SYSTEMS MATERIALS SELECTION Vehicle – 124 Flight Demonstration of X-33 Vehicle High Altitude Launch for a Practical Health Management System Compo- SSTO – 29 Overview of GRC’s Advanced Sensor and Instrumentation Development – 41 nents on the F/A-18 Systems Research MATERIALS TESTS Aircraft – 15 Lightweight durable TPS: Milestone Perspective on the National Aero-Space The NASA Integrated Vehicle Health 4–151 Plane Program instrumentation develop- Management Technology Experiment for ment – 156 MATHEMATICAL MODELS X-37 – 53 A Compendium of Wind Statistics and MECHANICAL ENGINEERING X-33/RLV System Health Management/ Models for the NASA Space Shuttle and The 1994 NASA/USRA/ADP Design Vehicle Health Management – 107 Other Aerospace Vehicle Pro- Projects – 140 grams – 99 MANNED MARS MISSIONS Various advanced design projects pro- Ion engine propelled Earth-Mars cycler A Procedure for Structural Weight Esti- moting engineering education – 144 mation of Single Stage to Orbit Launch with nuclear thermal propelled transfer MECHANICAL PROPERTIES vehicle, volume 2 – 144 Vehicles (Interim User’s Manual) – 44 Characterization of C/SiC Ceramic Ma- A simple analytical aerodynamic model MANNED SPACE FLIGHT trix Composites (CMCs) with Novel Inter- of Langley Winged-Cone Aerospace face Fiber Coatings – 45 Advanced transportation system study: Plane concept – 145 Manned launch vehicle concepts for two Effects of Stiffening and Mechanical A Thermal Management Systems Model way transportation system payloads to Load on Thermal Buckling of Stiffened for the NASA GTX RBCC Concept – 53 LEO – 152 Cylindrical Shells – 121 Advanced Propulsion System Test and Human Transportation System (HTS) Materials Research for High Speed Civil Tell Capability – 133 study, volume 2 – 155 Transport and Generic Hypersonics- Autocommander: A Supervisory Control- MANNED SPACECRAFT Metals Durability – 127 ler for Integrated Guidance and Control Tantalum Addition to Zirconium Diboride Advanced transportation system study: for the 2nd Generation Reusable Launch for Improved Oxidation Resistance – 18 Manned launch vehicle concepts for two Vehicle – 40 way transportation system payloads to Thermal-Mechanical Cyclic Test of a LEO – 152 Background, current status, and prognosis of the ongoing slush hydrogen tech- Composite Cryogenic Tank for Reusable 23 Human Transportation System (HTS) nology development program for the Launch Vehicles – study, volume 2 – 155 NASP – 177 MEMBRANES Ion engine propelled Earth-Mars cycler Constrained minimization of smooth Durability of Membrane Electrode As- with nuclear thermal propelled transfer functions using a genetic algo- semblies (MEAs) in PEM Fuel Cells Op- vehicle, volume 2 – 144 rithm – 141 erated on Pure Hydrogen and Oxy- gen – 6 MANUFACTURING Design and Analysis of UHTC Leading A Damage Resistance Comparison Be- Edge Attachment – 44 Performance Evaluation of Electro- tween Candidate Polymer Matrix Com- chem’s PEM Fuel Cell Power Plant for Flush Airdata Sensing (FADS) System posite Feedline Materials – 88 NASA’s 2nd Generation Reusable Calibration Procedures and Results for Launch Vehicle – 6 Cryopumping in Cryogenic Insulations Blunt Forebodies – 92 for a Reusable Launch Vehicle – 2 Hypersonic vehicle model and control MESSAGE PROCESSING Demonstration of Advanced C/SiC law development using H(inﬁnity) and Real-time trajectory optimization on par- Cooled Ramp – 56 micron synthesis – 144 allel processors – 158

A-24 METAL BONDING MILITARY AIRCRAFT STARS: The Space Transportation Archi- Materials Research for High Speed Civil Winds of change: Expanding the frontiers tecture Risk System – 116 Transport and Generic Hypersonics- of flight. Langley Research Center’s 75 MOON Metals Durability – 127 years of accomplishment, 1917- Human lunar mission capabilities using 1992 – 165 METAL FOAMS SSTO, ISRU and LOX-augmented NTR MIMO (CONTROL SYSTEMS) Design of Experiments for the Thermal technologies: A preliminary assess- 137 Characterization of Metallic Foam – 1 Reconfigurable Flight Control Designs ment – With Application to the X-33 Ve- MULTIDISCIPLINARY DESIGN OPTIMIZA- METAL MATRIX COMPOSITES hicle – 95 TION Development of Aluminum Composites Two Reconfigurable Flight-Control De- Managing MDO Software Development for a Rocket Engine’s Lightweight Thrust sign Methods: Robust Servomechanism Projects – 37 Cell – 80 and Control Allocation – 76 Multidisciplinary Approach to Aerospike Development of Metal Matrix Compos- MINIMA Nozzle Design – 120 ites for NASA’s Advanced Propulsion Constrained minimization of smooth Multidisciplinary Optimization Branch Ex- Systems – 74 functions using a genetic algo- perience Using iSIGHT Software – 91 Lightweight Chambers for Thrust Assem- rithm – 141 Optimal technology investment strate- blies – 66 MINIMAX TECHNIQUE gies for a reusable launch vehicle – 134 Thermoviscoplastic response of Ti-15-3 An inverse dynamics approach to trajec- System Risk Assessment and Allocation under various loading conditions – 172 tory optimization for an aerospace plane – 163 in Conceptual Design – 27 METALS System Sensitivity Analysis Applied to MISSILE DEFENSE Automation of NDE on RSRM Metal the Conceptual Design of a Dual-Fuel Reusable launch vehicle development Components – 57 Rocket SSTO – 125 research – 140 MULTIENGINE VEHICLES METEOROID CONCENTRATION MISSION PLANNING Abort performance for a winged-body Results of Two-Stage Light-Gas Gun De- A Plan for Advanced Guidance and Con- single-stage to orbit vehicle – 137 velopment Efforts and Hypervelocity Im- trol Technology for 2nd Generation Re- pact Tests of Advanced Thermal Protec- usable Launch Vehicles – 52 Overview of the Beta II Two-Stage-To- tion Materials – 100 Orbit vehicle design – 172 Access to space study – 153 METHODOLOGY Advanced Range Safety System for High MULTILAYER INSULATION Advanced Guidance and Control Meth- Energy Vehicles – 15 Effective Thermal Conductivity of High ods for Reusable Launch Vehicles: Test Temperature Insulations for Reusable Results – 30 Human Transportation System (HTS) Launch Vehicles – 97 study, volume 2 – 155 Parametric Weight Comparison of Ad- MICROCRACKS Preliminary Sizing of Vertical Take-Off vanced Metallic, Ceramic Tile, and Ce- Evaluation of Microcracking in Two Rocket-Based Combined-Cycle Pow- ramic Blanket Thermal Protection Sys- Carbon-Fiber/Epoxy-Matrix Composite ered Launch Vehicles – 85 tems – 88 Cryogenic Tanks – 72 Systems Engineering Approach to Tech- MULTIPROCESSING (COMPUTERS) Permeability Testing of Impacted Com- nology Integration for NASA’s 2nd Gen- posite Laminates for Use on Reusable eration Reusable Launch Vehicle – 23 NASA Exhibits – 59 Launch Vehicles – 83 The rationale/benefits of nuclear thermal MULTISTAGE ROCKET VEHICLES X-33 LH2 Tank Failure Investigation rocket propulsion for NASA’s lunar space Analysis of the staging maneuver and Findings – 14 transportation system – 141 booster glideback guidance for a two- stage, winged, fully reusable launch ve- MIXING RATIOS MICROGRAVITY hicle – 159 Second Generation RLV Program – 71 Space Shuttle Main Engine Start with Off-Nominal Propellant Inlet Pres- Overview of the Beta II Two-Stage-To- MICROINSTRUMENTATION sures – 116 Orbit vehicle design – 172 Health Monitoring Technology for Ther- MODELS The design and evolution of the beta mal Protection Systems on Reusable Hy- A Strategy for Integrating a Large Finite two-stage-to-orbit horizontal takeoff and personic Vehicles – 71 Element Model Using MSC landing launch system – 159 Wireless Subsurface Microsensors for NASTRAN/PATRAN: X-33 Lessons NASA PROGRAMS Learned – 97 Health Monitoring of Thermal Protection An Overview of In-Stu Treatability Stud- Systems on Hypersonic Vehicles – 69 Reusable Rocket Engine Operability ies at Marshall Space Flight Center, Modeling and Analysis – 101 Huntsville, Alabama – 78 MICROSTRUCTURE System Risk Assessment and Allocation Advanced Durability and Damage Toler- General Public Space Travel and Tour- in Conceptual Design – 27 ance Design and Analysis Methods for ism – 96 Composite Structures: Lessons Learned MODULES Life Cycle Systems Engineering Ap- from NASA Technology Development On-Line Computation of a Local Attain- proach to NASA’s 2nd Generation Reus- Programs – 26 able Moment Set for Reusable Launch able Launch Vehicle – 39 Characterization of C/SiC Ceramic Ma- Vehicles – 42 NASA aeronautics: Research and tech- trix Composites (CMCs) with Novel Inter- MONTE CARLO METHOD nology program highlights – 164 face Fiber Coatings – 45 A Monte Carlo Dispersion Analysis of the NASA’s Advanced Space Transportation MILITARY AIR FACILITIES X-33 Simulation Software – 72 Program: RTA Project Summary – 42 X-33 Integrated Test Facility Extended Development of the X-33 Aerodynamic NASA’s Spaceliner Investment Area Range Simulation – 100 Uncertainty Model – 108 Technology Activities – 81

A-25 Overview of MSFC’s Applied Fluid Dy- NASTRAN The aerospace plane design challenge: namics Analysis Group Activities – 24 A Strategy for Integrating a Large Finite Credible computational fluid dynamics results – 176 Pathfinder Technologies Specialist, Element Model Using MSC X-37 – 69 NASTRAN/PATRAN: X-33 Lessons Third Generation RLV Structural Seal Learned – 97 Development Programs at NASA Research Reports: 1996 NASA/ASEE GRC – 41 Summer Faculty Fellowship Pro- NATIONAL AEROSPACE PLANE PRO- gram – 102 GRAM Trajectory optimization for the national A conceptual design of an unmanned test aerospace plane – 156 Reusable launch vehicle: Technology de- vehicle using an airbreathing propulsion velopment and test program – 135 Trajectory optimization for the National system – 154 aerospace plane – 160 Second Generation Reusable Launch A summary of the slush hydrogen tech- Turbulence modeling for hypersonic Vehicle Development and Global Com- nology program for the National Aero- flight – 161 petitiveness of US Space Transportation Space Plane – 140 Industry: Critical Success Factors As- sessment – 51 Addition of equilibrium air to an upwind NAVIER-STOKES EQUATION Navier-Stokes code and other first steps Addition of equilibrium air to an upwind Selection of Lockheed Martin’s Preferred toward a more generalized flow Navier-Stokes code and other first steps TSTO Configurations for the Space solver – 174 toward a more generalized flow Launch Initiative – 12 Aero-Space Plane: Flexible access to solver – 174 Space Launch Initiative: New Capabili- space – 150 Computational/Experimental Aeroheat- ties - New Horizons – 38 An overview of aeroelasticity studies for ing Predictions for X-33 Phase 2 Ve- Space Ops 2002: Bringing Space Opera- the National Aerospace Plane – 160 hicle – 110 tions into the 21st Century. Track 3: Op- Assessment of flying-quality criteria for Computational/Experimental Aeroheat- erations, Mission Planning and Control. ing Predictions for X-33 – 112 2nd Generation Reusable Launch air-breathing aerospacecraft – 167 Vehicle-Concepts for Flight Opera- Background, current status, and progno- PNS predictions of external/internal hy- tions – 35 sis of the ongoing slush hydrogen tech- personic flows for NASP propulsion applications – 157 Survey of NASA V and V nology development program for the Processes/Methods – 58 NASP – 177 Turbulence modeling for hypersonic flight – 161 The future challenge for aeropropul- High Heat Flux Facility – 151 sion – 167 High temperature NASP engine seal de- X-33 Computational velopment – 153 Aeroheating/Aerodynamic Predictions Third Generation RLV Structural Seal and Comparisons With Experimental Development Programs at NASA Hypersonic flows as related to the na- Data – 25 GRC – 41 tional aerospace plane – 162 X-33 Turbulent Aeroheating Measure- V and V of Advanced Systems at Hypersonic flows as related to the Na- ments and Predictions – 32 NASA – 58 tional Aerospace plane – 172 NAVIGATION AIDS Winds of change: Expanding the frontiers John C. Stennis Space Center over- of flight. Langley Research Center’s 75 view – 152 Testing the Preliminary X-33 Navigation years of accomplishment, 1917- System – 84 1992 – 165 NASP aeroservothermoelasticity studies – 174 NAVIGATION NASA SPACE PROGRAMS National Aero-Space Plane resource Lockheed Martin Response to the OSP An Approach to Establishing System reel – 155 Challenge – 20 Benefits for Technology in NASA’s Hy- Support to X-33/Resusable Launch Ve- personics Investment Area – 67 National Aerospace Plane Thermal De- velopment. (Latest citations from the hicle Technology Program – 86 NASA 2nd Generation RLV Program In- Aerospace Database) – 114 troduction, Status and Future NEURAL NETS National aerospace plane – 151 Plans – 31 Adaptive Critic Neural Network-Based National Aero-Space Plane – 155 Terminal Area Energy Management and NASA’s Orbital Space Plane Risk Re- Approach and Landing Guidance – 9 duction Strategy – 13 Neural Network and Regression Approxi- mations in High Speed Civil Transport Neural Network and Regression Approxi- Optimal technology investment strate- Aircraft Design Optimization – 106 mations in High Speed Civil Transport gies for a reusable launch vehicle – 134 Aircraft Design Optimization – 106 Perspective on the National Aero-Space Orbital Space Plane (OSP) Program at Plane Program instrumentation develop- Research Reports: 1996 NASA/ASEE Lockheed Martin – 7 ment – 156 Summer Faculty Fellowship Pro- gram – 102 Orbital Space Plane Program Sta- Prediction of the ullage gas thermal tus – 5 stratification in a NASP vehicle propellant System Identification of X-33 Neural Net- Performance Evaluation of Electro- tank experimental simulation using work – 8 FLOW-3D – 177 chem’s PEM Fuel Cell Power Plant for Using Neural Networks in Decision Mak- NASA’s 2nd Generation Reusable Rapid near-optimal aerospace plane tra- ing for a Reconfigurable Electro Me- Launch Vehicle – 6 jectory generation and guidance – 171 chanical Actuator (EMA) – 68

Reusable launch vehicle development Slush hydrogen technology pro- NEWTON THEORY research – 140 gram – 149 Flush Airdata Sensing (FADS) System Space Launch Initiative: New Capabili- Slush hydrogen transfer studies at the Calibration Procedures and Results for ties ... New Horizons – 36 NASA K-Site Test Facility – 167 Blunt Forebodies – 93

A-26 NITROGEN ATOMS NUCLEAR ENGINE FOR ROCKET VE- Optimal trajectories for an aerospace Surface Catalytic Efficiency of Advanced HICLES plane. Part 1: Formulation, results, and Carbon Carbon Candidate Thermal Pro- The rationale/benefits of nuclear thermal analysis – 176 tection Materials for SSTO Ve- rocket propulsion for NASA’s lunar space ORBITAL LAUNCHING transportation system – 141 hicles – 133 Conceptual design of two-stage-to-orbit NONDESTRUCTIVE TESTS NUCLEAR PROPULSION hybrid launch vehicle – 169 Acoustic Emission Monitoring of the Ion engine propelled Earth-Mars cycler ORBITAL MANEUVERING VEHICLES DC-XA Composite Liquid Hydrogen Tank with nuclear thermal propelled transfer NASA’s Orbital Space Plane Risk Re- During Structural Testing – 126 vehicle, volume 2 – 144 duction Strategy – 13 The rationale/benefits of nuclear thermal Automation of NDE on RSRM Metal ORBITAL MANEUVERS Components – 57 rocket propulsion for NASA’s lunar space transportation system – 141 Inlet Development for a Rocket Based Impact Characteristics of Candidate Ma- Combined Cycle, Single Stage to Orbit terials for Single-Stage-to-Orbit (SSTO) NUCLEAR ROCKET ENGINES Vehicle Using Computational Fluid Dy- Technology – 123 Human lunar mission capabilities using namics – 95 Milestone 4: Thrust structure concepts SSTO, ISRU and LOX-augmented NTR OXIDATION RESISTANCE technologies: A preliminary assess- and IHM screening Graphite Composite Oxidation Behavior of Copper Alloy Can- ment – 137 Primary Structure (GCPS) – 143 didates for Rocket Engine Applications Test plan. GCPS task 7, subtask 7.1: IHM Ion engine propelled Earth-Mars cycler (Technical Poster) – 48 with nuclear thermal propelled transfer development – 149 Oxidation-Reduction Resistance of Ad- vehicle, volume 2 – 144 Thermographic Analysis of Composite vanced Copper Alloys – 11 NUMERICAL ANALYSIS Cobonds on the X-33 – 75 Tantalum Addition to Zirconium Diboride Predictor-Corrector Entry Guidance for for Improved Oxidation Resistance – 18 NONEQUILIBRIUM FLOW Reusable Launch Vehicles – 74 An implementation of a chemical and OXIDATION-REDUCTION REACTIONS thermal nonequilibrium flow solver on NUMERICAL FLOW VISUALIZATION Oxidation-Reduction Resistance of Ad- unstructured meshes and application to Computational/Experimental Aeroheat- vanced Copper Alloys – 11 blunt bodies – 145 ing Predictions for X-33 Phase 2 Ve- hicle – 110 OXIDATION NONINTRUSIVE MEASUREMENT Oxidation Behavior of Copper Alloy Can- An Experimental Study of Characteristic ON-LINE SYSTEMS didates for Rocket Engine Applications Combustion-Driven Flow for CFD Valida- On-Line Computation of a Local Attain- (Technical Poster) – 48 tion – 103 able Moment Set for Reusable Launch Vehicles – 42 OXIDIZERS NONLINEAR EQUATIONS Space Shuttle Main Engine Off-Nominal A parametric sensitivity study for single- Real-time trajectory optimization on par- Low Power Level Operation – 117 stage-to-orbit hypersonic vehicles using allel processors – 158 Space Shuttle Main Engine Start with trajectory optimization – 153 OPERATING COSTS Off-Nominal Propellant Inlet Pres- NONLINEAR PROGRAMMING An Approach to Establishing System sures – 116 Benefits for Technologies In NASA’s Neural Network and Regression Approxi- OXYGEN ANALYZERS mations in High Speed Civil Transport Spaceliner Investment Area – 77 Evaluation of the Linear Aerospike SR-71 Aircraft Design Optimization – 106 OPERATIONS RESEARCH Experiment (LASRE) Oxygen Sen- Real-time trajectory optimization on par- Operations Analysis of Space Shuttle sor – 91 allel processors – 158 System – 50 OXYGEN ATOMS NOSE CONES O/S analysis of conceptual space ve- Surface Catalytic Efficiency of Advanced Wingless Flight: The Lifting Body hicles – 134 Carbon Carbon Candidate Thermal Pro- 105 Story – OPTICAL MEASUREMENT tection Materials for SSTO Ve- hicles – 133 NOSES (FOREBODIES) An Experimental Study of Characteristic Aerodynamic Characteristics of a Vertical Combustion-Driven Flow for CFD Valida- OXYGEN-HYDROCARBON ROCKET EN- Takeoff Vertical Landing (VTVL) Single tion – 115 GINES Stage to Orbit Vehicle from M(sub infin- Non-Toxic Dual Thrust Reaction Control OPTICAL RADAR ity) = 0.1 to 10 – 122 Engine Development for On-Orbit APS NASA Stennis Space Center V and V Applications – 11 NOZZLE DESIGN Capabilities Overview – 2 Multidisciplinary Approach to Aerospike OXYGEN Nozzle Design – 120 OPTIMAL CONTROL Evaluation of the Linear Aerospike SR-71 Optimal trajectories for an aerospace Experiment (LASRE) Oxygen Sen- NOZZLE FLOW plane. Part 2: Data, tables, and sor – 92 Computations of Internal and External graphs – 176 Experimental Evaluation of a Subscale Axisymmetric Nozzle Aerodynamics at OPTIMIZATION Gaseous Hydrogen/Gaseous Oxygen Transonic Speeds – 4 34 Comparison of Two Multidisciplinary Op- Coaxial Rocket Injector – Space Shuttle Main Engine Off-Nominal timization Strategies for Launch-Vehicle OZONE DEPLETION Low Power Level Operation – 117 Design – 123 Director’s Discretionary Fund Report for NOZZLE GEOMETRY Constrained minimization of smooth Fiscal Year 1996 – 120 Computations of Internal and External functions using a genetic algo- OZONOSPHERE Axisymmetric Nozzle Aerodynamics at rithm – 141 Aircraft Emission Scenarios Projected in Transonic Speeds – 4 Multidisciplinary design techniques ap- Year 2015 for the NASA Technology Con- Thrust Enhancement in Hypervelocity plied to conceptual aerospace vehicle cept Aircraft (TCA) High Speed Civil Nozzles by Chemical Catalysis – 102 design – 156 Transport – 108

A-27 PANEL FLUTTER Test Results for an Off-the-Shelf POLAR ORBITS Panel Flutter and Sonic Fatigue Analysis GPS/INS During Approach and Landing Overview of the Beta II Two-Stage-To- for RLV – 77 Testing of the X-40A – 63 Orbit vehicle design – 172 PANELS Testing of Densified Liquid Hydrogen POLYIMIDE RESINS Stratification in a Scale Model Propellant Advanced Metallic Thermal Protection Damage tolerance of candidate thermo- Tank – 67 System Development – 64 set composites for use on single stage to Testing the Preliminary X-33 Navigation orbit vehicles – 148 Analytical comparison of three stiffened System – 84 panel concepts – 135 Impact Characteristics of Candidate Ma- Thermal/Fluid Analysis of a Composite terials for Single-Stage-to-Orbit (SSTO) Development of a Pressure Box to Evalu- Heat Exchanger for Use on the RLV Technology – 123 ate Reusable-Launch-Vehicle Rocket Engine – 63 Cryogenic-Tank Panels – 132 POLYIMIDES PERMEABILITY Characterization of Polyimide Foams for PARALLEL COMPUTERS Characterization of Thin Film Polymers Ultra-Lightweight Space Struc- Automated CFD Parameter Studies on Through Dynamic Mechanical Analysis tures – 22 Distributed Parallel Computers – 36 and Permeation – 22 High Temperature Polyimide Materials in Managing MDO Software Development Permeability After Impact Testing of Extreme Temperature Environ- Projects – 37 Composite Laminates – 21 ments – 73 PARALLEL PROCESSING (COMPUTERS) Permeability Testing of Impacted Com- POLYMER MATRIX COMPOSITES NASA Exhibits – 59 posite Laminates for Use on Reusable A Damage Resistance Comparison Be- Launch Vehicles – 80 tween Candidate Polymer Matrix Com- Real-time trajectory optimization on par- posite Feedline Materials – 88 allel processors – 158 PHOTOCONDUCTORS Director’s Discretionary Fund Report for Detection of Micro-Leaks Through Com- PARAMETER IDENTIFICATION Fiscal Year 1996 – 120 plex Geometries Under Mechanical Load A parametric sensitivity study for single- PHOTOGRAPHY and at Cryogenic Temperature – 79 stage-to-orbit hypersonic vehicles using National Aero-Space Plane resource Impact Characteristics of Candidate Ma- trajectory optimization – 153 reel – 155 terials for Single-Stage-to-Orbit (SSTO) PAYLOADS National Aero-Space Plane – 155 Technology – 123 Development of Metallic Thermal Protec- PIEZOELECTRIC ACTUATORS Lightweight Chambers for Thrust Assem- tion Systems for the Reusable Launch blies – 82 Vehicle – 124 Impedance-Based Structural Health Monitoring for Composite Laminates at Permeability After Impact Testing of Highly reusable space transportation: Cryogenic Environments – 7 Composite Laminates – 21 Approaches for reducing ETO launch Permeability Testing of Impacted Com- costs to $100 - $200 per pound of pay- PILOTLESS AIRCRAFT posite Laminates for Use on Reusable load – 135 A conceptual design of an unmanned test vehicle using an airbreathing propulsion Launch Vehicles – 80 Structural Seal Development – 11 system – 154 POLYMERIC FILMS Tether Transportation System PITCHING MOMENTS Characterization of Thin Film Polymers Study – 109 A simple analytical aerodynamic model Through Dynamic Mechanical Analysis 22 PERFORMANCE PREDICTION of Langley Winged-Cone Aerospace and Permeation – Plane concept – 145 2nd Generation Reusable Launch Ve- POTENTIAL FLOW hicle (2G RLV) – 82 PLANETARY ATMOSPHERES Flush Airdata Sensing (FADS) System Design and Analysis of Turbines for Thermal Protection Materials for Reentry Calibration Procedures and Results for Space Applications – 16 and Planetary Applications – 68 Blunt Forebodies – 93 Investigation of Fundamental Modeling Thermal Protection Materials for Reentry POWDER METALLURGY Applications – 59 and Thermal Performance Issues for a Oxidation-Reduction Resistance of Ad- Metallic Thermal Protection System De- PLANNING vanced Copper Alloys – 11 sign – 33 National Rocket Propulsion Materials PREDICTION ANALYSIS TECHNIQUES Performance Evaluation of Electro- Plan: A NASA, Department of Defense, and Industry Partnership – 73 A method of predicting quasi-steady chem’s PEM Fuel Cell Power Plant for aerodynamics for flutter analysis of high NASA’s 2nd Generation Reusable PLASMA FREQUENCIES speed vehicles using steady CFD calcu- Launch Vehicle – 6 The X-33 Extended Flight Test lations – 157 Range – 99 PERFORMANCE TESTS Ceramic Matrix Composites (CMC) Life Fabrication and Testing of Ceramic Ma- PLASMA RADIATION Prediction Method Development – 88 Development of Processing Techniques trix Composite Rocket Propulsion Com- Vision for CFD-Based Combustion Insta- 75 for Advanced Thermal Protection Materi- ponents – bility Predictions – 23 als – 112 Infrared Sensing Aeroheating Flight Ex- periement: STS-96 Flight Results – 84 PLUMES PREDICTOR-CORRECTOR METHODS Predictor-Corrector Entry Guidance for Milestone 4: Test plan for Reusable Hy- Analysis of Aerospike Plume Induced 109 Reusable Launch Vehicles – 74 drogen Composite Tank System Base-Heating Environment – (RHCTS). Task 3: Composite tank mate- Base-Bleed Effect on X-33 Aerospike PRESSURE DISTRIBUTION rials – 142 Plume Induced Base-Heating Environ- Hypersonic flows as related to the na- ment During Power-Pack Out – 25 tional aerospace plane – 162 Performance Tests of a Liquid Hydrogen Propellant Densification Ground System X-33 XRS-2200 Linear Aerospike Engine Status of the RBCC Direct-Connect for the X33/RLV – 118 Sea Level Plume Radiation – 66 Mixer Combustor Experiment – 54

A-28 X-33 Computational Integrated propulsion technology demon- Two Reconfigurable Flight-Control De- Aeroheating/Aerodynamic Predictions strator. Program plan – 147 sign Methods: Robust Servomechanism 76 and Comparisons With Experimental Operations Analysis of Space Shuttle and Control Allocation – 25 Data – System – 50 PROPOSALS PRESSURE MEASUREMENT Structural Arrangement Trade Study. Vol- Business Plan for the Southwest Re- Flush Airdata Sensing (FADS) System ume 1: Reusable Hydrogen Composite gional Spaceport: Executive Sum- Calibration Procedures and Results for Tank System (RHCTS) and Graphite mary – 114 92 Blunt Forebodies – Composite Primary Structures (GCPS). PROPULSION SYSTEM CONFIGURA- PRESSURE SENSORS Executive summary – 141 TIONS Propellant Feed System Leak Detection: PROPELLANT STORAGE A parametric sensitivity study for single- stage-to-orbit hypersonic vehicles using Lessons Learned From the Linear Aero- Task 4 supporting technology. Densifica- 93 trajectory optimization – 153 spike SR-71 Experiment (LASRE) – tion requirements definition and test ob- PRESSURE VESSEL DESIGN jectives. Propellant densification require- Access to space – 143 The 1994 NASA/USRA/ADP Design ments definition – 139 Advanced Electric Propulsion for RLV Projects – 140 PROPELLANT TANKS Launched Geosynchronous Space- PRESSURE VESSELS Composite,Cryogenic, Conformal, Com- craft – 90 Development of a Pressure Box to Evalu- mon Bulkhead, Aerogel-Insulated Tank Advanced Low-Cost O2/H2 Engines for ate Reusable-Launch-Vehicle (CBAT) – 74 the SSTO Application – 123 Cryogenic-Tank Panels – 132 Cyclic Cryogenic Thermal-Mechanical Advances in hypersonic vehicle synthe- Evaluation of Microcracking in Two Testing of an X-33/RLV Liquid Oxygen sis with application to studies of ad- Carbon-Fiber/Epoxy-Matrix Composite Tank Concept – 94 vanced thermal protection sys- 72 Cryogenic Tanks – Detection of Micro-Leaks Through Com- tem – 138 PRESSURE plex Geometries Under Mechanical Load Affordable Flight Demonstration of the Advanced Control Surface Seal Develop- and at Cryogenic Temperature – 79 GTX Air-Breathing SSTO Vehicle Con- ment at NASA GRC for 3rd Generation Findings from the X-33 Hydrogen Tank cept – 46 RLV – 29 Failure Investigation – 60 Alternate Propulsion Subsystem Con- PROBABILITY THEORY Milestone 4: Test plan for Reusable Hy- cepts Tripropellant Comparison An Overview of PRA with Applications to drogen Composite Tank System Study – 132 Aerospace Systems – 51 (RHCTS). Task 3: Composite tank mate- Analysis of a New Rocket-Based NASA Overview – 35 rials – 142 Combined-Cycle Engine Concept at Low Speed – 94 System Level Uncertainty Assessment Prediction of the ullage gas thermal for Collaborative RLV Design – 46 stratification in a NASP vehicle propellant Analysis of Parallel Burn Without Cross- tank experimental simulation using feed TSTO RLV Architectures and Com- System Risk Assessment and Allocation FLOW-3D – 177 parison to Parallel Burn With Crossfeed in Conceptual Design – 27 Reusable Launch Vehicle Technology and Series Burn Architectures – 19 PROCEDURES Program – 127 Design and Analysis of a Two-Stage-to- Flush Airdata Sensing (FADS) System Testing of Densified Liquid Hydrogen Orbit Airbreathing Hypersonic Vehicle Calibration Procedures and Results for Stratification in a Scale Model Propellant Concept – 129 Blunt Forebodies – 93 Tank – 67 Dual-Fuel Propulsion in Single-Stage Ad- Reusable Rocket Engine Operability Thermal Structures Technology Develop- vanced Manned Launch System Ve- Modeling and Analysis – 101 125 ment for Reusable Launch Vehicle Cryo- hicle – PRODUCT DEVELOPMENT genic Propellant Tanks – 107 Evaluation of the Linear Aerospike SR-71 Experiment (LASRE) Oxygen Sen- 2000 NASA Seal/Secondary Air System Trade study plan for Reusable Hydrogen 62 sor – 92 Workshop – Composite Tank System Advanced High Temperature Structural (RHCTS) – 145 Hypersonic vehicle model and control Seals – 43 law development using H(infinity) and X-33 LH2 Tank Failure Investigation micron synthesis – 144 Selection and Evaluation of An Alloy for Findings – 74 Nozzle Application – 22 Integrated propulsion technology demon- X-33 Tank Failure During Autoclave Fab- strator. Program plan – 147 PROGRAM VERIFICATION (COMPUT- rication – 60 ERS) Linear Aerospike SR-71 Experiment Survey of NASA V and V PROPELLANT TESTS (LASRE): Aerospace Propulsion Hazard Processes/Methods – 58 Engineering the Future of Full-Scale Pro- Mitigation Systems – 102 pulsion Testing – 7 V and V of Advanced Systems at Magnetohydrodynamic Augmented Pro- NASA – 58 PROPELLANT TRANSFER pulsion Experiment – 52 PROJECT MANAGEMENT Slush hydrogen propellant production, Operationally Efficient Propulsion Sys- Managing MDO Software Development transfer, and expulsion studies at the tem Study (OEPSS) data book. Volume Projects – 37 NASA K-Site Facility – 173 2: Ground operations problems – 170 NASA’s New Orbital Space Plane: A PROPELLANTS Performance Evaluation of the NASA Bridge to the Future – 19 NASA’s Space Launch Initiative Targets GTX RBCC Flowpath – 70 Toxic Propellants – 64 PROJECT PLANNING Propulsion module design review. Mile- stone 8: Integrated Propulsion Technol- Background, current status, and progno- PROPORTIONAL CONTROL ogy Demonstrator – 150 sis of the ongoing slush hydrogen tech- Reconfigurable Flight Control Designs nology development program for the With Application to the X-33 Ve- Reusable launch vehicle: Technology de- NASP – 177 hicle – 95 velopment and test program – 135

A-29 Rocket Based Combined Cycle (RBCC) Operationally Efficient Propulsion Sys- The future challenge for aeropropul- Propulsion Technology Workshop. Vol- tem Study (OEPSS) data book. Volume sion – 167 ume 1: Executive summary – 166 2: Ground operations problems – 170 PROPULSIVE EFFICIENCY Rocket Based Combined Cycle (RBCC) Operationally efficient propulsion system Technology Roadmap for Dual-Mode Propulsion Workshop, volume 2 – 166 study (OEPSS) data book – 110 Scramjet Propulsion to Support Space- Selection process for trade study: Graph- Operationally Efficient Propulsion Sys- Access Vision Vehicle Develop- ite Composite Primary Structure tem Study (OEPSS): OEPSS Video ment – 38 (GCPS) – 148 Script – 110 Thrust augmentation options for the Beta Status of the RBCC Direct-Connect Performance Evaluation of the NASA 2 two-stage-to-orbit vehicle – 154 Mixer Combustor Experiment – 54 GTX RBCC Flowpath – 70 PROTECTIVE COATINGS Structural Sizing of a 25,000-lb Payload, Performance Tests of a Liquid Hydrogen Evaluation of Thermal Control Coatings Air-Breathing Launch Vehicle for Single- Propellant Densification Ground System for Flexible Ceramic Thermal Protection Stage-to-Orbit – 85 for the X33/RLV – 118 Systems – 117 Test Plan. GCPS Task 4, subtask 4.2 Propulsion module design review. Mile- PROTEINS thrust structure development – 147 stone 8: Integrated Propulsion Technol- Advanced High Brilliance X-Ray ogy Demonstrator – 150 Source – 111 The future challenge for aeropropulsion – 167 Rapid near-optimal trajectory generation PROTONS and guidance law development for Durability of Membrane Electrode As- The rationale/benefits of nuclear thermal single-stage-to-orbit airbreathing ve- semblies (MEAs) in PEM Fuel Cells Op- rocket propulsion for NASA’s lunar space hicles – 177 erated on Pure Hydrogen and Oxy- transportation system – 141 gen – 6 Reusable Launch Vehicle Technology Thrust augmentation options for the Beta Program – 127 Performance Evaluation of Electro- 2 two-stage-to-orbit vehicle – 154 chem’s PEM Fuel Cell Power Plant for Rocket Based Combined Cycle (RBCC) NASA’s 2nd Generation Reusable PROPULSION SYSTEM PERFORMANCE Propulsion Technology Workshop. Vol- Launch Vehicle – 6 A conceptual design of an unmanned test ume 1: Executive summary – 166 PROTOTYPES vehicle using an airbreathing propulsion Rocket Based Combined Cycle (RBCC) system – 154 Propulsion Workshop, volume 2 – 166 Main Engine Prototype Development for 2nd Generation RLV RS-83 – 43 A parametric sensitivity study for single- Status of the RBCC Direct-Connect Selection process for trade study: Reus- stage-to-orbit hypersonic vehicles using Mixer Combustor Experiment – 54 trajectory optimization – 153 able Hydrogen Composite Tank System Technology Roadmap for Dual-Mode (RHCTS) – 148 Advanced Electric Propulsion for RLV Scramjet Propulsion to Support Space- PROVING Launched Geosynchronous Space- Access Vision Vehicle Develop- craft – 90 ment – 38 Design, Development,and Testing of Um- billical System Mechanisms for the X-33 Affordable Flight Demonstration of the Test Report for NASA MSFC Support of Advanced Technology Demonstra- GTX Air-Breathing SSTO Vehicle Con- the Linear Aerospike SR-71 Experiment tor – 90 cept – 46 (LASRE) – 89 PULSED LASERS Alternate Propulsion Subsystem Con- The future challenge for aeropropul- Evaluation and Characterization Study of cepts Tripropellant Comparison sion – 167 Dual Pulse Laser-Induced Spark (DPLIS) Study – 132 The StarBooster System: A Cargo Air- for Rocket Engine Ignition System Appli- An Approach to Establishing System craft for Space – 92 cation – 14 Benefits for Technology in NASA’s Hy- PROPULSION PUMPS personics Investment Area – 67 Water Flow Performance of a Superscale Advanced Transportation System Stud- Model of the Fastrac Liquid Oxygen Analysis of a New Rocket-Based ies Technical Area 2 (TA-2) Heavy Lift Pump – 55 Combined-Cycle Engine Concept at Low Launch Vehicle Development Con- Speed – 94 tract – 129 PURGING Analysis of Parallel Burn Without Cross- Human lunar mission capabilities using Evaluation of the Linear Aerospike SR-71 feed TSTO RLV Architectures and Com- SSTO, ISRU and LOX-augmented NTR Experiment (LASRE) Oxygen Sen- parison to Parallel Burn With Crossfeed technologies: A preliminary assess- sor – 91 and Series Burn Architectures – 19 ment – 137 QUADRATIC PROGRAMMING Effect of Aeropropulsive Interactions and National Rocket Propulsion Materials Two Reconfigurable Flight-Control De- Design Sensitivities on Optimal Hyper- Plan: A NASA, Department of Defense, sign Methods: Robust Servomechanism sonic Ascent Trajectories – 125 and Industry Partnership – 73 and Control Allocation – 76 Hot dynamic test rig for measuring hyper- Preliminary Assessment of Using Gelled RADIO TELEMETRY sonic engine seal flow and durabil- and Hybrid Propellant Propulsion for X-33 Integrated Test Facility Extended ity – 151 VTOL/SSTO Launch Systems – 104 Range Simulation – 100 Launch Vehicle Sizing Benefits Utilizing Research Reports: 1996 NASA/ASEE RAMAN SPECTROSCOPY Main Propulsion System Crossfeed and Summer Faculty Fellowship Pro- An Experimental Study of Characteristic Project Status – 34 gram – 102 Combustion-Driven Flow for CFD Valida- Linear Aerospike SR-71 Experiment Supercharged Ejector Ramjet with Ma- tion – 103 (LASRE): Aerospace Propulsion Hazard glifter Launch Assist (for the Argus Con- RAMJET ENGINES Mitigation Systems – 102 cept) – 120 Analysis of a New Rocket-Based Near-Optimal Operation of Dual-Fuel Support to 2nd Generation RLV Propul- Combined-Cycle Engine Concept at Low Launch Vehicles – 119 sion Project Office – 60 Speed – 94

A-30 Supercharged Ejector Ramjet with Ma- Evaluation of Thermal Control Coatings New V and V Tools for Diagnostic Mod- glifter Launch Assist (for the Argus Con- for Flexible Ceramic Thermal Protection eling Environment (DME) – 58 120 117 cept) – Systems – O/S analysis of conceptual space ve- The design and evolution of the beta REENTRY PHYSICS hicles – 134 two-stage-to-orbit horizontal takeoff and Infrared Sensing Aeroheating Flight Ex- System Risk Assessment and Allocation 159 landing launch system – periement: STS-96 Flight Results – 84 in Conceptual Design – 27 RAMPS (STRUCTURES) REENTRY SHIELDING RELIABILITY Ceramic Matrix Composite Cooled Task 4 supporting technology. Part 1: A Quantitative Reliability, Maintainability Nozzle Material Development Pro- Detailed test plan for leading edge tile and Supportability Approach for NASA’s gram – 21 development. Leading edge material de- Second Generation Reusable Launch velopment and testing – 139 RANGE SAFETY Vehicle – 41 Task 4 supporting technology. Part 2: Advanced Range Safety System for High Analysis of space concepts enabled by Detailed test plan for thermal seals. Ther- Energy Vehicles – 15 new transportation (ASCENT) mal seals evaluation, improvement and study – 28 REACTION KINETICS test. CAN8-1, Reusable Launch Vehicle An Investigation of the Effect of Surface (RLV), advanced technology demonstra- National Rocket Propulsion Materials Impurities on the Adsorption Kinetics of tor: X-33. Leading edge and seals ther- Plan: A NASA, Department of Defense, Hydrogen Chemisorbed onto Iron – 101 mal protection system technology dem- and Industry Partnership – 73 REAL TIME OPERATION onstration – 139 Reusable launch vehicle reliability, main- NASA IVHM Technology Experiment for REENTRY TRAJECTORIES tainability and operability assess- X-vehicles (NITEX) – 67 An Automated Method to Compute Or- ment – 150 Rapid near-optimal aerospace plane tra- bital Re-entry Trajectories with Heating Rocket Engine Health Management: jectory generation and guidance – 171 Constraints – 39 Early Definition of Critical Flight Mea- Real-Time Sensor Validation, Signal Re- An Automated Method to Compute Or- surements – 19 construction, and Feature Detection for bital Re-Entry Trajectories with Heating Systems Engineering Approach to Tech- an RLV Propulsion Testbed – 98 Constraints – 53 nology Integration for NASA’s 2nd Gen- Real-time trajectory optimization on par- Minimum Heating Re-Entry Trajectories eration Reusable Launch Vehicle – 23 allel processors – 158 for Advanced Hypersonic Launch Ve- REMOTE SENSING hicles – 119 Robust intelligent flight control for hyper- Design and Calibration of the X-33 Flush sonic vehicles – 166 Near-Optimal Re-Entry Trajectories for Airdata Sensing (FADS) System – 113 Reusable Launch Vehicles – 115 Using Neural Networks in Decision Mak- Director’s Discretionary Fund Report for ing for a Reconfigurable Electro Me- REENTRY VEHICLES Fiscal Year 1996 – 120 chanical Actuator (EMA) – 68 Reduction of Base Drag on Launch Ve- hicles – 50 REMOTE SENSORS RECOMBINATION COEFFICIENT Development of Structural Health Man- REFRACTORY COATINGS Surface Catalysis and Characterization agement Technology for Aerospace Ve- of Proposed Candidate TPS for Access- Current Technology for Thermal Protec- hicles – 1 to-Space Vehicles – 118 tion Systems – 165 REQUIREMENTS Surface Catalytic Efficiency of Advanced REFRACTORY MATERIALS Carbon Carbon Candidate Thermal Pro- Current Technology for Thermal Protec- Integrated propulsion technology demon- 147 tection Materials for SSTO Ve- tion Systems – 165 strator. Program plan – hicles – 133 Development of Processing Techniques RESCUE OPERATIONS RECOVERABILITY for Advanced Thermal Protection Materi- A Collection of Technical Papers – 122 als – 112 Conceptual design of two-stage-to-orbit RESEARCH AIRCRAFT hybrid launch vehicle – 169 Fractographic Analysis of HfB2-SiC and A conceptual design of an unmanned test 57 RECOVERABLE LAUNCH VEHICLES ZrB2-SiC Composites – vehicle using an airbreathing propulsion Optimal technology investment strate- High temperature NASP engine seal de- system – 154 153 gies for a reusable launch vehicle – 134 velopment – NASA aeronautics: Research and tech- Propulsion module design review. Mile- High Temperature Polyimide Materials in nology program highlights – 164 Extreme Temperature Environ- stone 8: Integrated Propulsion Technol- The F-18 systems research aircraft facil- ments – 73 ogy Demonstrator – 150 ity – 163 The X-33 Extended Flight Test Tantalum Addition to Zirconium Diboride Range – 99 for Improved Oxidation Resistance – 18 RESEARCH AND DEVELOPMENT Advanced High Temperature Structural RECOVERABLE SPACECRAFT REFUELING Seals – 43 Advanced transportation system studies. A Cryogenic Propellant Production Depot Technical area 2: Heavy lift launch ve- for Low Earth Orbit – 81 Fuel Cell Activities at the NASA Glenn Research Center – 49 hicle development – 129 REINFORCED PLATES Main Engine Prototype Development for REDUNDANT COMPONENTS Analytical comparison of three stiffened 2nd Generation RLV RS-83 – 43 Using Neural Networks in Decision Mak- panel concepts – 135 ing for a Reconfigurable Electro Me- RELIABILITY ANALYSIS On-Orbit Propulsion System Project Overview – 44 chanical Actuator (EMA) – 68 Access to space study – 153 Operations Analysis of Space Shuttle REENTRY EFFECTS Inlet Development for a Rocket Based System – 50 Effective Thermal Conductivity of High Combined Cycle, Single Stage to Orbit Temperature Insulations for Reusable Vehicle Using Computational Fluid Dy- Pathfinder Technologies Specialist, Launch Vehicles – 97 namics – 95 X-37 – 69

A-31 Second Generation Reusable Launch Adaptive Critic Neural Network-Based Composite,Cryogenic, Conformal, Com- Vehicle Development and Global Com- Terminal Area Energy Management and mon Bulkhead, Aerogel-Insulated Tank petitiveness of US Space Transportation Approach and Landing Guidance – 9 (CBAT) – 74 Industry: Critical Success Factors As- Advanced Control Surface Seal Develop- Cost Per Pound From Orbit – 47 sessment – 175 ment at NASA GRC for 3rd Generation Cryopumping in Cryogenic Insulations Technology Roadmap for Dual-Mode RLV – 29 for a Reusable Launch Vehicle – 2 Scramjet Propulsion to Support Space- Advanced Electric Propulsion for RLV Design and Analysis of a Two-Stage-to- Access Vision Vehicle Develop- Launched Geosynchronous Space- 38 Orbit Airbreathing Hypersonic Vehicle ment – craft – 90 Concept – 129 The Center for Aerospace Research: A Advanced Guidance and Control Meth- NASA Center of Excellence at North ods for Reusable Launch Vehicles: Test Design and Testing of Non-Toxic RCS Carolina Agricultural and Technical State Results – 30 Thrusters for Second Generation Reus- University – 162 able Launch Vehicle – 14 Advanced High Temperature Structural RESEARCH FACILITIES Seals – 43 Development and Capabilities of Unique John C. Stennis Space Center over- Structural Seal Test Rigs – 41 view – 152 Advanced Propulsion System Test and Tell Capability – 133 Development of a Pressure Box to Evalu- Overview of Fluid Dynamic Activities at ate Reusable-Launch-Vehicle the Marshall Space Flight Center – 56 Affordable Flight Demonstration of the Cryogenic-Tank Panels – 132 GTX Air-Breathing SSTO Vehicle Con- Overview of MSFC’s Applied Fluid Dy- cept – 26 Development of Metallic Thermal Protec- namics Analysis Group Activities – 24 tion Systems for the Reusable Launch Affordable In-Space Transportation Vehicle – 124 Slush hydrogen technology pro- Phase 2: An Advanced Concepts gram – 149 Project – 113 Discrete Roughness Effects on Shuttle Orbiter at Mach 6 – 37 The F-18 systems research aircraft facil- Affordable In-Space Transporta- ity – 163 tion – 106 Dynamics and Control of Orbiting Space Structures NASA Advanced Design Pro- Using CFD as Rocket Injector Design An Advanced Video Sensor for Auto- gram (ADP) – 128 Tool: Recent Progress at Marshall Space mated Docking – 61 Flight Center – 3 Earth-to-Orbit Rocket Propulsion – 17 An Approach to Establishing System RESEARCH MANAGEMENT Benefits for Technologies In NASA’s Effective Thermal Conductivity of High Overview of Fluid Dynamic Activities at Spaceliner Investment Area – 77 Temperature Insulations for Reusable 56 Launch Vehicles – 97 the Marshall Space Flight Center – An Automated Method to Compute Or- RESEARCH VEHICLES bital Re-entry Trajectories with Heating Estimate of avoidance maneuver rate for Configuration development study of the Constraints – 39 HASTOL tether boost facility – 28 OSU 1 hypersonic research ve- An Automated Method to Compute Or- Estimating the Cost of NASA’s Space hicle – 154 bital Re-Entry Trajectories with Heating Launch Initiative: How SLI Cost Stack Up Hyper-X Wind Tunnel Program – 109 Constraints – 53 Against the Shuttle – 51 Numerical investigations in three- Analysis of Parallel Burn Without Cross- Evaluation of Thermal Control Coatings dimensional internal flows – 171 feed TSTO RLV Architectures and Com- for Flexible Ceramic Thermal Protection parison to Parallel Burn With Crossfeed Systems – 117 RESEARCH and Series Burn Architectures – 19 An Investigation of the Effect of Surface Experimental Performance Evaluation of Analysis of space concepts enabled by Impurities on the Adsorption Kinetics of a Supersonic Turbine for Rocket Engine new transportation (ASCENT) Hydrogen Chemisorbed onto Iron – 101 Applications – 10 study – 28 REUSABLE HEAT SHIELDING Fabrication and Testing of Ceramic Ma- Analysis of the staging maneuver and trix Composite Rocket Propulsion Com- Task 4 supporting technology. Part 1: booster glideback guidance for a two- ponents – 75 Detailed test plan for leading edge tile stage, winged, fully reusable launch ve- development. Leading edge material de- hicle – 159 Flight Testing the Linear Aerospike velopment and testing – 139 SR-71 Experiment (LASRE) – 100 Analytical comparison of three stiffened REUSABLE LAUNCH VEHICLES panel concepts – 135 Fractographic Analysis of HfB2-SiC and 2nd Generation Reusable Launch Ve- ZrB2-SiC Composites – 57 Automation of NDE on RSRM Metal hicle (2G RLV) – 82 Components – 57 Genetic Algorithm Tuned Fuzzy Logic for 2nd Generation Reusable Launch Ve- Gliding Return Trajectories – 12 Boeing 2nd Generation Reusable hicle Potential Commercial Development Launch Vehicle Architecture – 21 GTX Reference Vehicle Structural Verifi- Scenarios – 79 cation Methods and Weight Sum- Bonding and Sealing Evaluations for 2nd Generation RLV: Program Goals and mary – 44 Cryogenic Tanks – 116 Acquisition Strategy – 79 High Temperature Polyimide Materials in Calculation of Supersonic Combustion A Plan for Advanced Guidance and Con- Extreme Temperature Environ- Using Implicit Schemes – 30 trol Technology for 2nd Generation Re- ments – 73 52 CFD-Based Design Optimization for usable Launch Vehicles – Highly reusable space transportation: Single Element Rocket Injector – 25 A Quantitative Reliability, Maintainability Approaches for reducing ETO launch and Supportability Approach for NASA’s Characterization of C/SiC Ceramic Ma- costs to $100 - $200 per pound of pay- Second Generation Reusable Launch trix Composites (CMCs) with Novel Inter- load – 135 Vehicle – 41 face Fiber Coatings – 48 Highly Reusable Space Transportation Accelerating the Design of Space Ve- Composite Development and Applica- System Concept Evaluation (The Argus hicles – 10 tions for RLV Tankage – 16 Launch Vehicle) – 103

A-32 Impedance-Based Structural Health Operationally Efficient Propulsion Sys- Structural Sizing of a 25,000-lb Payload, Monitoring for Composite Laminates at tem Study (OEPSS) data book. Volume Air-Breathing Launch Vehicle for Single- Cryogenic Environments – 7 2: Ground operations problems – 170 Stage-to-Orbit – 85 Improved Re-Configurable Sliding Mode Operations Analysis of Space Shuttle Subsonic aerodynamic characteristics of Controller for Reusable Launch Vehicle System – 50 a proposed advanced manned launch of Second Generation Addressing Aero- Operations Analysis of the 2nd Genera- system orbiter configuration – 160 dynamic Surface Failures and Thrust De- tion Reusable Launch Vehicle – 43 ficiencies – 34 Supercharged Ejector Ramjet with Ma- Orbital Space Plane (OSP) Program at glifter Launch Assist (for the Argus Con- Large Composite Structures Processing Lockheed Martin – 7 cept) – 120 Technologies for Reusable Launch Ve- hicles – 75 Orbital Space Plane Program Sta- Support to 2nd Generation RLV Propul- tus – 5 sion Project Office – 60 Launch Vehicle Sizing Benefits Utilizing Main Propulsion System Crossfeed and Overview of GRC’s Advanced Sensor Systems Engineering Approach to Tech- Project Status – 34 and Instrumentation Development – 41 nology Integration for NASA’s 2nd Gen- eration Reusable Launch Vehicle – 23 Life Cycle Systems Engineering Ap- Performance Evaluation of Electro- proach to NASA’s 2nd Generation Reus- chem’s PEM Fuel Cell Power Plant for Task 4 supporting technology. Densifica- able Launch Vehicle – 39 NASA’s 2nd Generation Reusable tion requirements definition and test ob- Launch Vehicle – 6 Liquid flyback booster pre-phase: A study jectives. Propellant densification require- 139 assessment – 146 Progress and Status of the 2nd Genera- ments definition – tion Reusable Launch Vehicle Pro- Task 4 supporting technology. Part 1: Main Engine Prototype Development for gram – 76 2nd Generation RLV RS-83 – 43 Detailed test plan for leading edge tile Requirements report for SSTO vertical development. Leading edge material de- Manufacturing Process Simulation of take-off/horizontal landing ve- velopment and testing – 139 Large-Scale Cryotanks – 36 hicle – 149 Task 4 supporting technology. Part 2: Minimum Heating Re-Entry Trajectories Reusable Launch Vehicle Attitude Con- Detailed test plan for thermal seals. Ther- for Advanced Hypersonic Launch Ve- trol Using a Time-Varying Sliding Mode mal seals evaluation, improvement and hicles – 119 Control Technique – 40 test. CAN8-1, Reusable Launch Vehicle Mission and sizing analysis for the Beta 2 Reusable launch vehicle development (RLV), advanced technology demonstra- two-stage-to-orbit vehicle – 168 research – 140 tor: X-33. Leading edge and seals thermal protection system technology dem- Multi-Disciplinary Analysis for Future Reusable Launch Vehicle Re- onstration – 139 Launch Systems Using NASA’s Ad- search – 119 Technology Roadmap for Dual-Mode vanced Engineering Environment Reusable launch vehicle: Technology de- Scramjet Propulsion to Support Space- (AEE) – 17 velopment and test program – 135 Access Vision Vehicle Develop- NASA 2nd Generation RLV Program In- RLV Turbine Performance Optimiza- ment – 38 troduction, Status and Future tion – 83 Plans – 31 Test Report for NASA MSFC Support of Rocket Engine Health Management: the Linear Aerospike SR-71 Experiment NASA Alternate Access to Station Ser- Early Definition of Critical Flight Mea- (LASRE) – 89 vice Concept – 65 surements – 19 NASA EPSCoR Preparation Grant – 49 Tether Transportation System Second Generation Reusable Launch Study – 109 NASA IVHM Technology Experiment for Vehicle Development and Global Com- X-vehicles (NITEX) – 67 petitiveness of US Space Transportation The design and evolution of the beta Industry: Critical Success Factors As- two-stage-to-orbit horizontal takeoff and NASA’s Hypersonic Investment sessment – 51 landing launch system – 159 Area – 47 Second Generation RLV Program – 71 The rationale/benefits of nuclear thermal NASA’s Orbital Space Plane Risk Re- rocket propulsion for NASA’s lunar space duction Strategy – 13 Second Generation RLV Space Vehicle Concept – 65 transportation system – 141 NASA’s Space Launch Initiative Targets The StarBooster System: A Cargo Air- Toxic Propellants – 64 Selection of Lockheed Martin’s Preferred TSTO Configurations for the Space craft for Space – 92 Natural Atmospheric Environment Model Launch Initiative – 12 Development for the National Aeronau- The X-33 Extended Flight Test tics and Space Administration’s Second Space Launch Initiative: New Capabili- Range – 99 Generation Reusable Launch Ve- ties ... New Horizons – 36 Thermal Protection Materials for Reentry hicle – 57 Space Launch Initiative: New Capabili- and Planetary Applications – 68 ties - New Horizons – 38 Near-Optimal Re-Entry Trajectories for Thermal Protection Materials for Reentry Reusable Launch Vehicles – 115 Space Ops 2002: Bringing Space Opera- Applications – 59 tions into the 21st Century. Track 3: Op- New V and V Tools for Diagnostic Mod- Thermal Structures Technology Develop- eling Environment (DME) – 58 erations, Mission Planning and Control. 2nd Generation Reusable Launch ment for Reusable Launch Vehicle Cryo- On-Board Entry Trajectory Planning Ex- Vehicle-Concepts for Flight Opera- genic Propellant Tanks – 107 panded to Sub-orbital Flight – 12 tions – 35 Thermal/Fluid Analysis of a Composite On-Line Computation of a Local Attain- Space Transportation Systems Tech- Heat Exchanger for Use on the RLV able Moment Set for Reusable Launch nologies – 70 Rocket Engine – 37 Vehicles – 42 Statistical Methods for Rapid Aerother- Third Generation RLV Structural Seal On-Orbit Propulsion System Project mal Analysis and Design Technol- Development Programs at NASA Overview – 44 ogy – 55 GRC – 3

A-33 Transportation Systems Analyses (TSA): Systems Engineering Approach to Tech- Systems Challenges for Hypersonic Ve- Highly Reusable Space Transportation nology Integration for NASA’s 2nd Gen- hicles – 117 (HRST). A preliminary programmatic as- eration Reusable Launch Vehicle – 23 Thermal/Fluid Analysis of a Composite sessment – 131 Test Plan. GCPS Task 4, subtask 4.2 Heat Exchanger for Use on the RLV Update on the Development and Capa- thrust structure development – 147 Rocket Engine – 37 bilities of Unique Structural Seal Test ROBUSTNESS (MATHEMATICS) Water Flow Performance of a Superscale Rigs – 3 Hypersonic vehicle model and control Model of the Fastrac Liquid Oxygen V and V of Advanced Systems at law development using H(infinity) and Pump – 55 NASA – 58 micron synthesis – 144 ROCKET NOZZLES Wireless Subsurface Sensors for Health Reusable Launch Vehicle Attitude Con- Ceramic Matrix Composite Cooled Monitoring of Thermal Protection Sys- trol Using a Time-Varying Sliding Mode tems on Hypersonic Vehicles – 68 Nozzle Material Development Pro- Control Technique – 40 gram – 21 REUSABLE ROCKET ENGINES ROCKET ENGINE DESIGN Multidisciplinary Approach to Aerospike Earth-to-Orbit Rocket Propulsion – 17 2nd Generation Reusable Launch Ve- Nozzle Design – 120 Reusable Rocket Engine Operability hicle (2G RLV) – 82 Modeling and Analysis – 101 ROCKET-BASED COMBINED-CYCLE EN- Experimental Evaluation of a Subscale GINES Gaseous Hydrogen/Gaseous Oxygen REUSABLE SPACECRAFT A Thermal Management Systems Model Coaxial Rocket Injector – 34 An Approach to Establishing System for the NASA GTX RBCC Concept – 53 Benefits for Technologies In NASA’s Integrated propulsion technology demon- Spaceliner Investment Area – 77 strator. Program plan – 147 Affordable Flight Demonstration of the GTX Air-Breathing SSTO Vehicle Con- Conceptual design of two-stage-to-orbit Propulsion module design review. Mile- cept – 26 hybrid launch vehicle – 169 stone 8: Integrated Propulsion Technol- ogy Demonstrator – 150 Calculation of Supersonic Combustion Earth-to-Orbit Rocket Propulsion – 17 Using Implicit Schemes – 30 Highly Reusable Space Transportation Rocket Based Combined Cycle (RBCC) Propulsion Technology Workshop. Vol- Status of the RBCC Direct-Connect (HRST) Baseline Concepts and Analysis: Mixer Combustor Experiment – 54 Rocket/RBCC Options – 114 ume 1: Executive summary – 166 Structural Seal Development – 11 Ion engine propelled Earth-Mars cycler Rocket-Based Combined-Cycle (RBCC) with nuclear thermal propelled transfer Propulsion Technology Workshop. Tuto- ROLLING MOMENTS rial session – 168 vehicle, volume 2 – 144 Subsonic Aerodynamic Characteristics of Mission and sizing analysis for the Beta 2 Using CFD as Rocket Injector Design a Circular Body Earth-to-Orbit Ve- two-stage-to-orbit vehicle – 168 Tool: Recent Progress at Marshall Space hicle – 128 Flight Center – 3 Structures for the 3rd Generation Reus- ROTATING ENVIRONMENTS able Concept Vehicle – 69 ROCKET ENGINES Ion engine propelled Earth-Mars cycler Transportation Systems Analyses (TSA): Analysis of a New Rocket-Based with nuclear thermal propelled transfer Highly Reusable Space Transportation Combined-Cycle Engine Concept at Low vehicle, volume 2 – 144 (HRST). A preliminary programmatic as- Speed – 94 ROTATION sessment – 131 CFD-Based Design Optimization for Progress in Unsteady Turbopump Flow Single Element Rocket Injector – 25 X-33 Phase 2 – 115 Simulations Using Overset Grid Sys- X-33 – 105 Evaluation and Characterization Study of tems – 30 Dual Pulse Laser-Induced Spark (DPLIS) REYNOLDS AVERAGING for Rocket Engine Ignition System Appli- SAFETY FACTORS Inlet Development for a Rocket Based cation – 14 Systems Engineering Approach to Tech- Combined Cycle, Single Stage to Orbit nology Integration for NASA’s 2nd Gen- Vehicle Using Computational Fluid Dy- Human lunar mission capabilities using eration Reusable Launch Vehicle – 24 namics – 95 SSTO, ISRU and LOX-augmented NTR technologies: A preliminary assess- SAFETY REYNOLDS NUMBER ment – 137 A Quantitative Reliability, Maintainability Computational/Experimental Aeroheat- Operationally efficient propulsion system and Supportability Approach for NASA’s ing Predictions for X-33 Phase 2 Ve- study (OEPSS) data book – 110 Second Generation Reusable Launch hicle – 110 Vehicle – 41 X-33 Turbulent Aeroheating Measure- Options for flight testing rocket-based combined-cycle (RBCC) engines – 130 NASA’s Hypersonic Investment ments and Predictions – 32 Area – 47 Oxidation Behavior of Copper Alloy Can- RISK didates for Rocket Engine Applications National Rocket Propulsion Materials An Overview of PRA with Applications to (Technical Poster) – 48 Plan: A NASA, Department of Defense, Aerospace Systems – 51 and Industry Partnership – 73 Rocket Engine Health Management: NASA Overview – 35 Early Definition of Critical Flight Mea- Statistical Methods for Rapid Aerother- NASA’s Orbital Space Plane Risk Re- surements – 19 mal Analysis and Design Technol- duction Strategy – 13 ogy – 55 Rocket-Based Combined-Cycle (RBCC) On-Orbit Propulsion System Project Propulsion Technology Workshop. Tuto- Systems Engineering Approach to Tech- Overview – 44 rial session – 168 nology Integration for NASA’s 2nd Gen- eration Reusable Launch Vehicle – 23 STARS: The Space Transportation Archi- Rocket-Induced Magnetohydrodynamic tecture Risk System – 116 Ejector: A Single-Stage-to-Orbit Ad- SALT SPRAY TESTS System Risk Assessment and Allocation vanced Propulsion Concept – 131 Lightweight durable TPS: Milestone in Conceptual Design – 27 Second Generation RLV Program – 71 4–151

A-34 SANDWICH STRUCTURES Third Generation RLV Structural Seal SILICON CARBIDES X-33 Tank Failure During Autoclave Fab- Development Programs at NASA Ceramic Matrix Composites (CMC) Life rication – 64 GRC – 41 Prediction Method Development – 88 SATELLITE CONTROL SEMICONDUCTOR LASERS Characterization of C/SiC Ceramic Ma- Subsonic Aerodynamic Characteristics of Director’s Discretionary Fund Report for trix Composites (CMCs) with Novel Inter- a Circular Body Earth-to-Orbit Ve- Fiscal Year 1996 – 120 face Fiber Coatings – 45 hicle – 128 SENSITIVITY ANALYSIS Demonstration of Advanced C/SiC Cooled Ramp – 56 SCALE MODELS Investigation of Fundamental Modeling Discrete Roughness Effects on Shuttle and Thermal Performance Issues for a X-37 C-SiC CMC Control Surface Com- Orbiter at Mach 6 – 37 Metallic Thermal Protection System De- ponents Development: Status of the Testing of Densified Liquid Hydrogen sign – 33 NASA/Boeing/USAF Orbital Vehicle and 20 Stratification in a Scale Model Propellant Preliminary Thermal-Mechanical Sizing Related Efforts – Tank – 67 of Metallic TPS: Process Development SILICONES Water Flow Performance of a Superscale and Sensitivity Studies – 33 Task 4 supporting technology. Part 1: Model of the Fastrac Liquid Oxygen SENSORS Detailed test plan for leading edge tile development. Leading edge material de- Pump – 55 Mosad and Stream Vision For A Telero- velopment and testing – 139 X-33 Rev-F Turbulent Aeroheating Re- botic, Flying Camera System – 59 SIMULATION sults From Test 6817 in NASA Langley Overview of GRC’s Advanced Sensor 20-Inch Mach 6 Air Tunnel and Compari- and Instrumentation Development – 41 Advanced Control Surface Seal Develop- sons With Computations – 26 ment at NASA GRC for 3rd Generation Real-Time Sensor Validation, Signal Re- RLV – 29 SCHEDULES construction, and Feature Detection for Integrated propulsion technology demon- an RLV Propulsion Testbed – 98 Inlet Development for a Rocket Based strator. Program plan – 147 Combined Cycle, Single Stage to Orbit SEPARATED FLOW Vehicle Using Computational Fluid Dy- SEALERS Analysis of Aerospike Plume Induced namics – 95 Task 4 supporting technology. Part 2: Base-Heating Environment – 109 Integrated G and C Implementation Detailed test plan for thermal seals. Ther- Hypersonic flows as related to the na- within IDOS: A Simulink Based Reusable mal seals evaluation, improvement and tional aerospace plane – 162 Launch Vehicle Simulation – 11 test. CAN8-1, Reusable Launch Vehicle (RLV), advanced technology demonstra- Space Shuttle Main Engine Off-Nominal Manufacturing Process Simulation of tor: X-33. Leading edge and seals ther- Low Power Level Operation – 117 Large-Scale Cryotanks – 36 mal protection system technology dem- SERVOCONTROL Opportunities for Breakthroughs in onstration – 139 NASP aeroservothermoelasticity stud- Large-Scale Computational Simulation SEALING ies – 174 and Design – 54 Bonding and Sealing Evaluations for Progress in Unsteady Turbopump Flow Cryogenic Tanks – 116 SERVOMECHANISMS Simulations Using Overset Grid Sys- Reconfigurable Flight Control Designs tems – 30 High temperature NASP engine seals: A With Application to the X-33 Ve- technology review – 173 hicle – 95 Progress in Unsteady Turbopump Flow Simulations – 51 Third Generation RLV Structural Seal Two Reconfigurable Flight-Control De- Development Programs at NASA sign Methods: Robust Servomechanism SINGLE STAGE TO ORBIT VEHICLES GRC – 3 and Control Allocation – 76 A Collaborative Analysis Tool for Thermal Update on the Development and Capa- Protection Systems for Single Stage to SHAPES bilities of Unique Structural Seal Test Orbit Launch Vehicles – 56 Rigs – 3 Wingless Flight: The Lifting Body A Collection of Technical Papers – 122 Story – 105 SEALS (STOPPERS) A conceptual design of an unmanned test SHARP LEADING EDGES 2000 NASA Seal/Secondary Air System vehicle using an airbreathing propulsion Workshop – 62 Design and Analysis of UHTC Leading system – 154 Edge Attachment – 44 Advanced Control Surface Seal Develop- A guidance and control assessment of ment at NASA GRC for 3rd Generation SHOCK WAVE INTERACTION three vertical landing options for RLV – 29 Hypersonic flows as related to the Na- RLV – 136 tional Aerospace Plane – 136 Advanced High Temperature Structural A parametric sensitivity study for single- Seals – 43 Hypersonic flows as related to the Na- stage-to-orbit hypersonic vehicles using tional Aerospace plane – 172 trajectory optimization – 153 Development and Capabilities of Unique A Procedure for Structural Weight Esti- Structural Seal Test Rigs – 41 SHROUDED PROPELLERS mation of Single Stage to Orbit Launch Development of a Pressure Box to Evalu- Unshrouded Impeller Technology Task Vehicles (Interim User’s Manual) – 44 ate Reusable-Launch-Vehicle Status – 81 Cryogenic-Tank Panels – 132 A Two-Stage-to-Orbit Spaceplane Con- SIDESLIP cept With Growth Potential – 80 High temperature NASP engine seal de- Transonic aerodynamic characteristics of velopment – 153 Abort performance for a winged-body a proposed wing-body reusable launch single-stage to orbit vehicle – 137 High temperature NASP engine seals: A vehicle concept – 138 Access to space study – 153 technology review – 173 SIGNAL PROCESSING Access to space – 143 Hot dynamic test rig for measuring hyper- Real-Time Sensor Validation, Signal Re- sonic engine seal flow and durabil- construction, and Feature Detection for Advanced Low-Cost O2/H2 Engines for ity – 151 an RLV Propulsion Testbed – 98 the SSTO Application – 123

A-35 Advanced transportation system studies. Human lunar mission capabilities using Selection process for trade study: Graph- Technical area 2: Heavy lift launch ve- SSTO, ISRU and LOX-augmented NTR ite Composite Primary Structure hicle development – 129 technologies: A preliminary assess- (GCPS) – 148 ment – 137 Advanced Transportation System Stud- Selection process for trade study: Reus- ies Technical Area 2 (TA-2) Heavy Lift Hypersonic vehicle control law develop- able Hydrogen Composite Tank System Launch Vehicle Development Con- ment using H infinity and mu- (RHCTS) – 148 tract – 129 synthesis – 164 Single-stage-to-orbit: Meeting the chal- Advances in hypersonic vehicle synthe- Hypersonic vehicle model and control lenge – 136 sis with application to studies of ad- law development using H(infinity) and SRM-Assisted Trajectory for the GTX vanced thermal protection sys- micron synthesis – 144 Reference Vehicle – 47 tem – 138 Integrated propulsion technology demon- Structural Arrangement Trade Study. Vol- Aerodynamic Characteristics of a Vertical strator. Program plan – 147 ume 1: Reusable Hydrogen Composite Takeoff Vertical Landing (VTVL) Single Milestone 4: Test plan for Reusable Hy- Tank System (RHCTS) and Graphite Stage to Orbit Vehicle from M(sub infin- Composite Primary Structures (GCPS). ity) = 0.1 to 10 – 122 drogen Composite Tank System (RHCTS). Task 3: Composite tank mate- Executive summary – 141 Affordable Flight Demonstration of the rials – 142 Structural arrangement trade study. Vol- GTX Air-Breathing SSTO Vehicle Con- ume 3: Reusable Hydrogen Composite cept – 26 Milestone 4: Thrust structure concepts and IHM screening Graphite Composite Tank System (RHCTS) and Graphite Air-breathing hypersonic vehicle guid- Primary Structure (GCPS) – 143 Composite Primary Structures (GCPS). ance and control studies; An integrated Addendum – 141 trajectory/control analysis methodology: Multidisciplinary design techniques ap- Subsonic-to-Hypersonic Aerodynamic Phase 1 – 171 plied to conceptual aerospace vehicle Characteristics for a Winged, Circular- design – 156 Air-breathing hypersonic vehicle guid- Body, Single-Stage-to-Orbit Spacecraft ance and control studies: An integrated Near-Optimal Operation of Dual-Fuel Configuration – 124 trajectory/control analysis methodology, Launch Vehicles – 119 System Sensitivity Analysis Applied to phase 2 – 165 Optimal technology investment strate- the Conceptual Design of a Dual-Fuel Airbreathing/Rocket Single-Stage-to- gies for a reusable launch vehicle – 134 Rocket SSTO – 125 Orbit Design Matrix – 124 Preliminary Assessment of Using Gelled Test Plan. GCPS Task 4, subtask 4.2 An overview of aeroelasticity studies for and Hybrid Propellant Propulsion for thrust structure development – 147 the National Aerospace Plane – 160 VTOL/SSTO Launch Systems – 104 Test plan. GCPS task 7, subtask 7.1: IHM Analysis of a New Rocket-Based Preliminary Sizing of Vertical Take-Off development – 149 Combined-Cycle Engine Concept at Low Rocket-Based Combined-Cycle Pow- Test plan. Task 5, subtask 5.2: Early Speed – 94 ered Launch Vehicles – 85 on-orbit TPSdebris impact tests – 148 Comparison of Two Multidisciplinary Op- Project Antares: A low cost modular Test Plans. Lightweight Durable TPS: timization Strategies for Launch-Vehicle launch vehicle for the future – 168 Tasks 1,2,4,5, and 6 – 146 Design – 123 Rapid near-optimal trajectory generation The 1994 NASA/USRA/ADP Design Computational Aerodynamics Analysis of and guidance law development for Projects – 140 Future Launch Vehicle Configura- single-stage-to-orbit airbreathing ve- tions – 122 hicles – 177 The Control System for the X-33 Linear Aerospike Engine – 105 Computational Study of a McDonnell Requirements report for SSTO vertical Douglas Single-Stage-to-Orbit Vehicle take-off and horizontal landing ve- The Reusable Launch Vehicle Technol- Concept for Aerodynamic Analy- hicle – 143 ogy Program and the X-33 Advanced sis – 126 Technology Demonstrator – 128 Requirements report for SSTO vertical Trade study plan for Graphite Composite Configuration development study of the take-off/horizontal landing ve- Primary Structure (GCPS) – 146 OSU 1 hypersonic research vehicle – 149 hicle – 154 Trade study plan for Reusable Hydrogen Reusable launch vehicle development Composite Tank System Damage tolerance of candidate thermo- research – 140 set composites for use on single stage to (RHCTS) – 145 orbit vehicles – 148 Reusable launch vehicle reliability, main- Transonic aerodynamic characteristics of tainability and operability assess- a proposed wing-body reusable launch Development of an integrated BEM ap- ment – 150 proach for hot fluid structure interac- vehicle concept – 138 tion – 169 Reusable Launch Vehicle Technology Trim drag reduction concepts for horizon- Program – 127 Dual-Fuel Propulsion in Single-Stage Ad- tal takeoff single-stage-to-Orbit ve- vanced Manned Launch System Ve- Robust intelligent flight control for hyper- hicles – 170 hicle – 125 sonic vehicles – 166 Various advanced design projects pro- Effect of Aeropropulsive Interactions and Rocket Based Combined Cycle (RBCC) moting engineering education – 144 Design Sensitivities on Optimal Hyper- Propulsion Technology Workshop. Vol- X-33 Phase 2 – 115 sonic Ascent Trajectories – 125 ume 1: Executive summary – 166 SLIDING High Altitude Launch for a Practical Rocket Based Combined Cycle (RBCC) SSTO – 29 Propulsion Workshop, volume 2 – 166 Improved Re-Configurable Sliding Mode Controller for Reusable Launch Vehicle Highly Reusable Space Transportation Rocket-Induced Magnetohydrodynamic of Second Generation Addressing Aero- (HRST) Baseline Concepts and Analysis: Ejector: A Single-Stage-to-Orbit Ad- dynamic Surface Failures and Thrust De- Rocket/RBCC Options – 114 vanced Propulsion Concept – 131 ficiencies – 34

A-36 SLUSH HYDROGEN Highly reusable space transportation: Propellant Densification for Shuttle: The A summary of the slush hydrogen tech- Approaches for reducing ETO launch SSME Perspective – 32 costs to $100 - $200 per pound of pay- nology program for the National Aero- Space Shuttle Main Engine Off-Nominal load – 135 Space Plane – 140 Low Power Level Operation – 117 Background, current status, and progno- International Space Station Evolution Space Shuttle Main Engine Start with sis of the ongoing slush hydrogen tech- Data Book – 87 Off-Nominal Propellant Inlet Pres- nology development program for the NASA Alternate Access to Station Ser- sures – 116 NASP – 177 vice Concept – 65 The design and evolution of the beta Slush hydrogen propellant production, Second Generation Reusable Launch two-stage-to-orbit horizontal takeoff and transfer, and expulsion studies at the Vehicle Development and Global Com- landing launch system – 159 NASA K-Site Facility – 173 petitiveness of US Space Transportation Slush hydrogen technology pro- Industry: Critical Success Factors As- SPACE SHUTTLE ORBITERS gram – 149 sessment – 51 Development of the X-33 Aerodynamic Uncertainty Model – 108 Slush hydrogen transfer studies at the SPACE DEBRIS NASA K-Site Test Facility – 167 Estimate of avoidance maneuver rate for Discrete Roughness Effects on Shuttle HASTOL tether boost facility – 28 Orbiter at Mach 6 – 37 Technology issues associated with using densified hydrogen for space ve- Results of Two-Stage Light-Gas Gun De- Infrared Sensing Aeroheating Flight Ex- hicles – 167 velopment Efforts and Hypervelocity Im- periment: STS-96 Flight Results – 86 pact Tests of Advanced Thermal Protec- Orbital Space Plane Program Sta- SLUSH tion Materials – 100 A summary of the slush hydrogen tech- tus – 5 SPACE ERECTABLE STRUCTURES nology program for the National Aero- The Evolution of Flexible Insulation as Space Plane – 140 Various advanced design projects pro- Thermal Protection Systems for Reus- moting engineering education – 144 able Launch Vehicles: AFRSI (Advanced SMART MATERIALS SPACE FLIGHT Flexible Reusable Surface Insulation) to Wireless Subsurface Microsensors for General Public Space Travel and Tour- CRI (Conformal Reusable Insula- Health Monitoring of Thermal Protection ism – 96 tion) – 67 Systems on Hypersonic Vehicles – 69 SPACE HABITATS SPACE SHUTTLES SMART STRUCTURES General Public Space Travel and Tour- Access to space study – 153 X-33/RLV System Health Management/ ism – 96 Vehicle Health Management – 107 An Approach to Establishing System SPACE INDUSTRIALIZATION Benefits for Technology in NASA’s Hy- SOFTWARE DEVELOPMENT TOOLS International Space Station Evolution personics Investment Area – 67 A Monte Carlo Dispersion Analysis of the Data Book – 87 Boeing 2nd Generation Reusable X-33 Simulation Software – 72 SPACE LOGISTICS Launch Vehicle Architecture – 21 New V and V Tools for Diagnostic Mod- A Collection of Technical Papers – 122 Current Technology for Thermal Protec- eling Environment (DME) – 58 Advanced transportation system study: tion Systems – 165 SOFTWARE ENGINEERING Manned launch vehicle concepts for two Estimating the Cost of NASA’s Space Managing MDO Software Development way transportation system payloads to Launch Initiative: How SLI Cost Stack Up Projects – 37 LEO – 152 Against the Shuttle – 51 Survey of NASA V and V SPACE MISSIONS High Voltage Design Guidelines: A Processes/Methods – 58 Access to space study – 153 Timely Update – 61 V and V of Advanced Systems at SPACE PROCESSING Operations Analysis of Space Shuttle NASA – 58 Research Reports: 1996 NASA/ASEE System – 50 SOLAR POWER SATELLITES Summer Faculty Fellowship Pro- Orbital Space Plane Program Sta- gram – 102 A Collection of Technical Papers – 122 tus – 5 Various advanced design projects pro- Progress and Status of the 2nd Genera- SOLID PROPELLANT ROCKET ENGINES moting engineering education – 144 Liquid flyback booster pre-phase: A study tion Reusable Launch Vehicle Pro- 76 assessment – 146 SPACE PROGRAMS gram – High Voltage Design Guidelines: A Space Ops 2002: Bringing Space Opera- National Rocket Propulsion Materials Timely Update – 61 Plan: A NASA, Department of Defense, tions into the 21st Century. Track 3: Op- and Industry Partnership – 73 SPACE SHUTTLE BOOSTERS erations, Mission Planning and Control. 2nd Generation Reusable Launch SRM-Assisted Trajectory for the GTX Liquid flyback booster pre-phase: A study assessment – 146 Vehicle-Concepts for Flight Opera- Reference Vehicle – 47 tions – 35 SPACE SHUTTLE MAIN ENGINE SONIC BOOMS 2nd Generation Reusable Launch Ve- SPACE STATION FREEDOM The 1995 NASA High-Speed Research hicle (2G RLV) – 82 John C. Stennis Space Center over- Program Sonic Boom Workshop – 126 view – 152 CFD-Based Design Optimization for SOUND PROPAGATION Single Element Rocket Injector – 25 Shuttle to Space Station. Heart assist The 1995 NASA High-Speed Research implant. Hubble update. X-30 mock- Program Sonic Boom Workshop – 126 Dual-Fuel Propulsion in Single-Stage Ad- up – 138 vanced Manned Launch System Ve- SPACE COMMERCIALIZATION hicle – 125 SPACE STATION MODULES Access to space study – 153 Overview of the Beta II Two-Stage-To- International Space Station Evolution Access to space – 143 Orbit vehicle design – 172 Data Book – 87

A-37 SPACE STATION STRUCTURES A guidance and control assessment of SPACECRAFT COMPONENTS International Space Station Evolution three vertical landing options for Automation of NDE on RSRM Metal Data Book – 87 RLV – 136 Components – 57 Access to space study – 153 SPACE STATIONS SPACECRAFT CONFIGURATIONS Advanced transportation system studies Access to space – 143 Analysis of a New Rocket-Based technical area 2(TA-2): Heavy lift launch Combined-Cycle Engine Concept at Low vehicle development – 130 Advanced Transportation System Stud- ies Technical Area 2 (TA-2) Heavy Lift Speed – 94 General Public Space Travel and Tour- Launch Vehicle Development Con- Orbital Space Plane (OSP) Program at ism – 96 131 tract – Lockheed Martin – 7 SPACE TRANSPORTATION SYSTEM Advanced transportation system studies Reusable launch vehicle: Technology de- FLIGHTS technical area 2(TA-2): Heavy lift launch velopment and test program – 135 A Hazardous Gas Detection System for vehicle development – 130 Aerospace and Commercial Applica- Selection of Lockheed Martin’s Preferred Air-breathing aerospace plane develop- tions – 98 TSTO Configurations for the Space ment essential: Hypersonic propulsion Launch Initiative – 12 SPACE TRANSPORTATION SYSTEM flight tests – 142 Subsonic-to-Hypersonic Aerodynamic A Hazardous Gas Detection System for An Approach to Establishing System Characteristics for a Winged, Circular- Aerospace and Commercial Applica- Benefits for Technology in NASA’s Hy- Body, Single-Stage-to-Orbit Spacecraft tions – 98 personics Investment Area – 67 Configuration – 124 A Two-Stage-to-Orbit Spaceplane Con- Analysis of space concepts enabled by cept With Growth Potential – 80 new transportation (ASCENT) SPACECRAFT CONSTRUCTION MATERI- ALS Advanced transportation system studies study – 28 Damage tolerance of candidate thermo- technical area 2 (TA-2): Heavy lift launch General Public Space Travel and Tour- vehicle development – 130 set composites for use on single stage to ism – 96 orbit vehicles – 148 Advanced transportation system study: Highly reusable space transportation: Impact Characteristics of Candidate Ma- Manned launch vehicle concepts for two Approaches for reducing ETO launch way transportation system payloads to terials for Single-Stage-to-Orbit (SSTO) costs to $100 - $200 per pound of pay- Technology – 123 LEO – 152 load – 135 National aerospace plane – 151 Affordable In-Space Transportation Highly Reusable Space Transportation Phase 2: An Advanced Concepts (HRST) Baseline Concepts and Analysis: Thermoviscoplastic response of Ti-15-3 113 Project – Rocket/RBCC Options – 114 under various loading conditions – 172 Affordable In-Space Transporta- Highly Reusable Space Transportation tion – 106 SPACECRAFT CONTROL System Concept Evaluation (The Argus A guidance and control assessment of An Approach to Establishing System Launch Vehicle) – 103 three vertical landing options for Benefits for Technologies In NASA’s Human Transportation System (HTS) RLV – 136 Spaceliner Investment Area – 77 study, volume 2 – 155 Air-breathing hypersonic vehicle guid- Boeing 2nd Generation Reusable ance and control studies: An integrated Launch Vehicle Architecture – 21 Ion engine propelled Earth-Mars cycler with nuclear thermal propelled transfer trajectory/control analysis methodology, Flight Demonstrations of Orbital Space vehicle, volume 2 – 144 phase 2 – 165 Plane (OSP) Technologies – 16 NASA’s Advanced Space Transportation An inverse dynamics approach to trajec- NASA’s Space Launch Initiative Targets Program: RTA Project Summary – 42 tory optimization for an aerospace Toxic Propellants – 64 plane – 163 NASA’s Hypersonic Investment Space Launch Initiative: New Capabili- Area – 47 Analysis of the staging maneuver and ties ... New Horizons – 36 booster glideback guidance for a two- NASA’s Spaceliner Investment Area stage, winged, fully reusable launch ve- Space Launch Initiative: New Capabili- Technology Activities – 81 hicle – 159 ties - New Horizons – 38 Orbital Space Plane (OSP) Program at Hypersonic vehicle control law develop- Space Transportation Systems Tech- Lockheed Martin – 7 ment using H infinity and mu- nologies – 70 Orbital Space Plane (OSP) Pro- synthesis – 164 STARS: The Space Transportation Archi- gram – 5 tecture Risk System – 116 Robust intelligent flight control for hyper- Progress and Status of the 2nd Genera- sonic vehicles – 166 Subsonic aerodynamic characteristics of tion Reusable Launch Vehicle Pro- a proposed advanced manned launch SPACECRAFT DESIGN gram – 76 system orbiter configuration – 160 Access to space – 143 Single-stage-to-orbit: Meeting the chal- The rationale/benefits of nuclear thermal lenge – 136 Advanced transportation system studies. rocket propulsion for NASA’s lunar space Technical area 2: Heavy lift launch ve- transportation system – 141 Structural Sizing of a 25,000-lb Payload, hicle development – 129 Transportation Systems Analyses (TSA): Air-Breathing Launch Vehicle for Single- Stage-to-Orbit – 85 Advanced Transportation System Stud- Highly Reusable Space Transportation ies Technical Area 2 (TA-2) Heavy Lift (HRST). A preliminary programmatic as- Systems Engineering Approach to Tech- Launch Vehicle Development Con- sessment – 131 nology Integration for NASA’s 2nd Gen- tract – 131 eration Reusable Launch Vehicle – 24 SPACE TRANSPORTATION Advanced transportation system studies 2nd Generation RLV: Program Goals and The StarBooster System: A Cargo Air- technical area 2 (TA-2): Heavy lift launch Acquisition Strategy – 79 craft for Space – 92 vehicle development – 130

A-38 Air-breathing hypersonic vehicle guid- SPACECRAFT INSTRUMENTS Space Launch Initiative: New Capabili- ance and control studies: An integrated Perspective on the National Aero-Space ties ... New Horizons – 36 trajectory/control analysis methodology, Plane Program instrumentation develop- Space Launch Initiative: New Capabili- 165 phase 2 – ment – 156 ties - New Horizons – 38 An integrated aerodynamic/propulsion SPACECRAFT LANDING The design and evolution of the beta study for generic aero-space planes Personnel launch system autoland de- two-stage-to-orbit horizontal takeoff and based on waverider concepts – 173 velopment study – 175 landing launch system – 159 Boeing 2nd Generation Reusable Testing the Preliminary X-33 Navigation Towers for Earth Launch – 29 Launch Vehicle Architecture – 21 System – 84 SPACECRAFT MAINTENANCE Dynamics and Control of Orbiting Space SPACECRAFT LAUNCHING O/S analysis of conceptual space ve- Structures NASA Advanced Design Pro- Access to space study – 153 hicles – 134 gram (ADP) – 128 Advanced transportation system study: Reusable launch vehicle reliability, main- Highly Reusable Space Transportation Manned launch vehicle concepts for two tainability and operability assess- System Concept Evaluation (The Argus way transportation system payloads to ment – 150 Launch Vehicle) – 103 LEO – 152 Test plan. GCPS task 7, subtask 7.1: IHM Ion engine propelled Earth-Mars cycler Analysis of the staging maneuver and development – 149 with nuclear thermal propelled transfer booster glideback guidance for a two- vehicle, volume 2 – 144 stage, winged, fully reusable launch ve- SPACECRAFT MANEUVERS Mission and sizing analysis for the Beta 2 hicle – 159 Analysis of the staging maneuver and booster glideback guidance for a two- two-stage-to-orbit vehicle – 168 Design, Development,and Testing of Um- stage, winged, fully reusable launch ve- billical System Mechanisms for the X-33 NASA 2nd Generation RLV Program In- hicle – 159 troduction, Status and Future Advanced Technology Demonstra- Plans – 31 tor – 90 SPACECRAFT ORBITS Orbital Space Plane (OSP) Program at NASA Overview – 35 Estimate of avoidance maneuver rate for HASTOL tether boost facility – 28 Lockheed Martin – 7 O/S analysis of conceptual space ve- High Altitude Launch for a Practical SPACECRAFT PERFORMANCE hicles – 134 SSTO – 29 Design and Testing of Non-Toxic RCS Project Antares: A low cost modular Thrusters for Second Generation Reus- Highly reusable space transportation: launch vehicle for the future – 168 able Launch Vehicle – 14 Approaches for reducing ETO launch Requirements report for SSTO vertical costs to $100 - $200 per pound of pay- SPACECRAFT POWER SUPPLIES take-off and horizontal landing ve- load – 135 Fuel Cell Activities at the NASA Glenn hicle – 143 Liquid flyback booster pre-phase: A study Research Center – 49 Reusable launch vehicle: Technology de- assessment – 146 SPACECRAFT PROPULSION velopment and test program – 135 Lockheed Martin Response to the OSP Advanced Electric Propulsion for RLV Selection of Lockheed Martin’s Preferred Challenge – 20 Launched Geosynchronous Space- craft – 90 TSTO Configurations for the Space Mission and sizing analysis for the Beta 2 Launch Initiative – 12 two-stage-to-orbit vehicle – 168 Air-breathing aerospace plane development essential: Hypersonic propulsion Structures for the 3rd Generation Reus- NASA 2nd Generation RLV Program In- flight tests – 138 able Concept Vehicle – 69 troduction, Status and Future Plans – 31 An integrated aerodynamic/propulsion The 1994 NASA/USRA/ADP Design study for generic aero-space planes Projects – 140 NASA Glenn Research Center’s Hyper- based on waverider concepts – 173 sonic Propulsion Program – 95 Trade study plan for Reusable Hydrogen Development and Capabilities of Unique NASA’s Next Generation Launch Tech- Composite Tank System Structural Seal Test Rigs – 41 (RHCTS) – 145 nology Program - Strategy and Plans – 4 Development of an integrated BEM ap- SPACECRAFT DOCKING proach for hot fluid structure interac- Operationally Efficient Propulsion Sys- tion – 169 An Advanced Video Sensor for Auto- tem Study (OEPSS) data book. Volume mated Docking – 61 2: Ground operations problems – 170 Integrated Technology Assessment Cen- ter (ITAC) Update – 42 Lockheed Martin Response to the OSP Operationally Efficient Propulsion Sys- Challenge – 20 tem Study (OEPSS): OEPSS Video JTEC panel report on space and transat- Script – 110 mospheric propulsion technology – 175 SPACECRAFT GUIDANCE Air-breathing hypersonic vehicle guid- Operations Analysis of the 2nd Genera- Linear Aerospike SR-71 Experiment ance and control studies: An integrated tion Reusable Launch Vehicle – 43 (LASRE): Aerospace Propulsion Hazard Mitigation Systems – 102 trajectory/control analysis methodology, Progress and Status of the 2nd Genera- phase 2 – 165 tion Reusable Launch Vehicle Pro- NASA’s Advanced Space Transportation Personnel launch system autoland de- gram – 76 Program: RTA Project Summary – 42 velopment study – 175 Reusable Launch Vehicle Re- NASA’s Spaceliner Investment Area Rapid near-optimal trajectory generation search – 119 Technology Activities – 81 and guidance law development for Second Generation RLV Space Vehicle On-Orbit Propulsion System Project single-stage-to-orbit airbreathing ve- Concept – 65 Overview – 44 hicles – 177 Selection of Lockheed Martin’s Preferred Operationally Efficient Propulsion Sys- Trajectory optimization for the National TSTO Configurations for the Space tem Study (OEPSS) data book. Volume aerospace plane – 160 Launch Initiative – 12 2: Ground operations problems – 170

A-39 Overview of MSFC’s Applied Fluid Dy- Analysis of space concepts enabled by STORAGE TANKS namics Analysis Group Activities – 24 new transportation (ASCENT) Cyclic Cryogenic Thermal-Mechanical 28 Rapid near-optimal trajectory generation study – Testing of an X-33/RLV Liquid Oxygen and guidance law development for High Altitude Launch for a Practical Tank Concept – 94 single-stage-to-orbit airbreathing ve- SSTO – 32 Development of a Pressure Box to Evalu- hicles – 177 SPACECREW TRANSFER ate Reusable-Launch-Vehicle Cryogenic-Tank Panels – 132 Second Generation RLV Program – 71 STS-102 Photo-op/Suit-up/Depart Technology issues associated with using O&C/Launch Discovery On Evaluation of Microcracking in Two densified hydrogen for space ve- Orbit/Landing/Crew Egress – 27 Carbon-Fiber/Epoxy-Matrix Composite Cryogenic Tanks – 72 hicles – 167 SPACECREWS Vision for CFD-Based Combustion Insta- STS-102 Photo-op/Suit-up/Depart Thermal-Mechanical Cyclic Test of a bility Predictions – 23 O&C/Launch Discovery On Composite Cryogenic Tank for Reusable Orbit/Landing/Crew Egress – 27 Launch Vehicles – 23 SPACECRAFT REENTRY STRATIFICATION Subsonic aerodynamic characteristics of SPARK IGNITION a proposed advanced manned launch Evaluation and Characterization Study of Testing of Densified Liquid Hydrogen system orbiter configuration – 160 Dual Pulse Laser-Induced Spark (DPLIS) Stratification in a Scale Model Propellant for Rocket Engine Ignition System Appli- Tank – 67 Thermal Protection Materials for Reentry cation – 14 Applications – 59 STRATOSPHERE SPECIFIC IMPULSE Stratospheric turbulence measurements SPACECRAFT RELIABILITY Advanced Low-Cost O2/H2 Engines for and models for aerospace plane de- O/S analysis of conceptual space ve- the SSTO Application – 123 sign – 164 hicles – 134 SR-71 AIRCRAFT STRESS ANALYSIS Test plan. GCPS task 7, subtask 7.1: IHM Evaluation of the Linear Aerospike SR-71 National Aerospace Plane Thermal De- development – 149 Experiment (LASRE) Oxygen Sen- velopment. (Latest Citations from the Test Results for an Off-the-Shelf sor – 92 Aerospace Database) – 113 GPS/INS During Approach and Landing Flight Stability and Control and Perfor- STRUCTURAL ANALYSIS Testing of the X-40A – 63 mance Results from the Linear Aerospike SR-71 Experiment (LASRE) – 101 Development of Metallic Thermal Protec- SPACECRAFT STABILITY tion Systems for the Reusable Launch Transonic aerodynamic characteristics of Flight Testing the Linear Aerospike Vehicle – 124 SR-71 Experiment (LASRE) – 100 a proposed wing-body reusable launch Preliminary Thermal-Mechanical Sizing vehicle concept – 138 Linear Aerospike SR-71 Experiment of Metallic TPS: Process Development SPACECRAFT STRUCTURES (LASRE): Aerospace Propulsion Hazard and Sensitivity Studies – 33 Mitigation Systems – 102 Dynamics and Control of Orbiting Space Structures for the 3rd Generation Reus- Structures NASA Advanced Design Pro- Test Report for NASA MSFC Support of able Concept Vehicle – 69 gram (ADP) – 128 the Linear Aerospike SR-71 Experiment (LASRE) – 89 The future challenge for aeropropul- Milestone 4: Thrust structure concepts sion – 167 and IHM screening Graphite Composite STABILITY TESTS Thermal Structures Technology Develop- Primary Structure (GCPS) – 143 An Improved RLV Stability Analysis Via a Continuation Approach – 9 ment for Reusable Launch Vehicle Cryo- Structural Arrangement Trade Study. Vol- genic Propellant Tanks – 107 ume 1: Reusable Hydrogen Composite STABILITY Tank System (RHCTS) and Graphite Dynamics and Stability and Control STRUCTURAL DESIGN CRITERIA Composite Primary Structures (GCPS). Characteristics of the X-37 – 62 Structures and materials technology for Executive summary – 141 hypersonic aerospacecraft – 178 N-Factor Computations for the X-33 Ve- Structural arrangement trade study. Vol- hicle – 76 Thermal Management Design for the ume 3: Reusable Hydrogen Composite X-33 Lifting Body – 104 Tank System (RHCTS) and Graphite STABILIZATION STRUCTURAL DESIGN Composite Primary Structures (GCPS). Advanced aeroservoelastic stabilization Addendum – 141 techniques for hypersonic flight ve- Access to space – 143 hicles – 163 Advanced High Temperature Structural SPACECRAFT TRAJECTORIES Seals – 87 Air-breathing hypersonic vehicle guid- STAGE SEPARATION ance and control studies: An integrated Stage Separation CFD Tool Develop- GTX Reference Vehicle Structural Verifi- trajectory/control analysis methodology, ment and Evaluation – 38 cation Methods and Weight Sum- mary – 44 phase 2 – 165 STATISTICAL ANALYSIS An inverse dynamics approach to trajec- Investigation of Fundamental Modeling International Space Station Evolution tory optimization for an aerospace and Thermal Performance Issues for a Data Book – 87 plane – 163 Metallic Thermal Protection System De- Milestone 4: Thrust structure concepts sign – 33 Rapid near-optimal trajectory generation and IHM screening Graphite Composite and guidance law development for Statistical Methods for Rapid Aerother- Primary Structure (GCPS) – 143 single-stage-to-orbit airbreathing ve- mal Analysis and Design Technol- Requirements report for SSTO vertical hicles – 177 ogy – 55 take-off and horizontal landing ve- SPACECRAFT STIFFENING hicle – 143 Advanced spacecraft designs in support Effects of Stiffening and Mechanical Selection process for trade study: Graph- of human missions to earth’s neighbor- Load on Thermal Buckling of Stiffened ite Composite Primary Structure hood – 28 Cylindrical Shells – 121 (GCPS) – 148

A-40 Structures and materials technology for SUPERSONIC COMBUSTION RAMJET Current Technology for Thermal Protec- hypersonic aerospacecraft – 178 ENGINES tion Systems – 165 Structures for the 3rd Generation Reus- Advances in hypersonic vehicle synthe- High Temperature Polyimide Materials in able Concept Vehicle – 69 sis with application to studies of ad- Extreme Temperature Environ- vanced thermal protection sys- ments – 73 The 1994 NASA/USRA/ADP Design tem – 138 Projects – 140 Materials Research for High Speed Civil Air-breathing aerospace plane develop- Transport and Generic Hypersonics- Thermal Management Design for the ment essential: Hypersonic propulsion Metals Durability – 127 X-33 Lifting Body – 104 flight tests – 138 Neural Network and Regression Approxi- Third Generation RLV Structural Seal Calculation of Supersonic Combustion Development Programs at NASA mations in High Speed Civil Transport Using Implicit Schemes – 30 106 GRC – 3 Aircraft Design Optimization – Computational parametric study of Various advanced design projects pro- sidewall-compression scramjet inlet per- SUPERSONIC TURBINES moting engineering education – 144 formance at Mach 10 – 161 Experimental Performance Evaluation of a Supersonic Turbine for Rocket Engine STRUCTURAL ENGINEERING Design and Analysis of a Two-Stage-to- Applications – 10 Orbit Airbreathing Hypersonic Vehicle High Altitude Launch for a Practical SUPPORT SYSTEMS SSTO – 32 Concept – 129 A Collection of Technical Papers – 122 Research Reports: 1996 NASA/ASEE Effect of Aeropropulsive Interactions and A study of facilities and fixtures for testing Summer Faculty Fellowship Pro- Design Sensitivities on Optimal Hyper- of a high speed civil transport wing com- gram – 102 sonic Ascent Trajectories – 125 ponent – 127 Structures and materials technology for Hyper-X Wind Tunnel Program – 109 O/S analysis of conceptual space ve- hypersonic aerospacecraft – 178 Options for flight testing rocket-based hicles – 134 combined-cycle (RBCC) engines – 130 Thermal Structures Technology Develop- Overview of Fluid Dynamic Activities at ment for Reusable Launch Vehicle Cryo- Technology Roadmap for Dual-Mode the Marshall Space Flight Center – 56 genic Propellant Tanks – 107 Scramjet Propulsion to Support Space- Access Vision Vehicle Develop- SURFACE DEFECTS STRUCTURAL FAILURE ment – 38 An Investigation of the Effect of Surface A Collaborative Analysis Tool for Thermal SUPERSONIC COMBUSTION Impurities on the Adsorption Kinetics of Protection Systems for Single Stage to Hydrogen Chemisorbed onto Iron – 101 Orbit Launch Vehicles – 56 Computation of H2/air reacting flowfields in drag-reduction external combus- SURFACE PROPERTIES X-33 LH2 Tank Failure Investigation tion – 161 Surface Catalysis and Characterization Findings – 14 The aerospace plane design challenge: of Proposed Candidate TPS for Access- STRUCTURAL MEMBERS Credible computational fluid dynamics to-Space Vehicles – 118 Acoustic Emission Monitoring of the results – 176 Surface Catalytic Efficiency of Advanced DC-XA Composite Liquid Hydrogen Tank SUPERSONIC FLOW Carbon Carbon Candidate Thermal Pro- 126 During Structural Testing – PNS predictions of external/internal hy- tection Materials for SSTO Ve- Analytical comparison of three stiffened personic flows for NASP propulsion ap- hicles – 133 panel concepts – 135 plications – 157 SURFACE ROUGHNESS STRUCTURAL RELIABILITY SUPERSONIC FLUTTER Comparison of Methods for Determining Boundary Layer Edge Conditions for Test plan. GCPS task 7, subtask 7.1: IHM Aerothermoelastic analysis of a NASP Transition Correlations – 23 development – 149 demonstrator model – 155 SURFACE TEMPERATURE STRUCTURAL STABILITY SUPERSONIC SPEED Improvements in Thermal Protection Siz- Advanced aeroservoelastic stabilization Optimal trajectories for an aerospace ing Capabilities for TCAT: Conceptual techniques for hypersonic flight ve- plane. Part 1: Formulation, results, and Design for Advanced Space Transporta- hicles – 163 analysis – 176 tion Systems – 61 Optimal trajectories for an aerospace STRUCTURAL WEIGHT plane. Part 2: Data, tables, and Infrared Sensing Aeroheating Flight Ex- A Procedure for Structural Weight Esti- graphs – 176 periement: STS-96 Flight Results – 84 mation of Single Stage to Orbit Launch Vehicles (Interim User’s Manual) – 44 Panel Flutter and Sonic Fatigue Analysis Near-Optimal Re-Entry Trajectories for for RLV – 77 Reusable Launch Vehicles – 115 SUBSONIC FLOW Subsonic-to-Hypersonic Aerodynamic SURVEYS Subsonic Aerodynamic Characteristics of Characteristics for a Winged, Circular- a Circular Body Earth-to-Orbit Ve- Survey of NASA V and V Body, Single-Stage-to-Orbit Spacecraft Processes/Methods – 58 hicle – 128 Configuration – 124 SURVIVAL SUBSONIC SPEED Supersonic aeroelastic instability results Subsonic-to-Hypersonic Aerodynamic for a NASP-like wing model – 159 Test plan. Task 5, subtask 5.2: Early Characteristics for a Winged, Circular- on-orbit TPSdebris impact tests – 148 SUPERSONIC TRANSPORTS Body, Single-Stage-to-Orbit Spacecraft SWEEP ANGLE Configuration – 124 A study of facilities and fixtures for testing of a high speed civil transport wing com- Computational parametric study of Transonic aerodynamic characteristics of ponent – 127 sidewall-compression scramjet inlet per- a proposed wing-body reusable launch formance at Mach 10 – 161 vehicle concept – 138 Aircraft Emission Scenarios Projected in Year 2015 for the NASA Technology Con- SWEPT WINGS SUPERCOMPUTERS cept Aircraft (TCA) High Speed Civil Supersonic aeroelastic instability results NASA Exhibits – 59 Transport – 108 for a NASP-like wing model – 159

A-41 SYNCHRONOUS PLATFORMS Multidisciplinary Optimization Branch Ex- TANKS (CONTAINERS) Advanced Electric Propulsion for RLV perience Using iSIGHT Software – 91 Composite Development and Applica- Launched Geosynchronous Space- Overview of MSFC’s Applied Fluid Dy- tions for RLV Tankage – 16 craft – 90 namics Analysis Group Activities – 24 TANTALUM SYNCHRONOUS SATELLITES Second Generation RLV Program – 71 Tantalum Addition to Zirconium Diboride Advanced Electric Propulsion for RLV for Improved Oxidation Resistance – 18 Second Generation RLV Space Vehicle Launched Geosynchronous Space- Concept – 65 TECHNOLOGICAL FORECASTING craft – 90 System Risk Assessment and Allocation National Rocket Propulsion Materials SYSTEM EFFECTIVENESS in Conceptual Design – 27 Plan: A NASA, Department of Defense, Airbreathing/Rocket Single-Stage-to- and Industry Partnership – 73 Orbit Design Matrix – 124 Systems Engineering Approach to Tech- nology Integration for NASA’s 2nd Gen- TECHNOLOGIES SYSTEMS ANALYSIS eration Reusable Launch Vehicle – 24 Reusable Launch Vehicle Technology Program – 127 Autocommander: A Supervisory Control- Third Generation RLV Structural Seal ler for Integrated Guidance and Control Development Programs at NASA The Reusable Launch Vehicle Technol- for the 2nd Generation Reusable Launch GRC – 41 ogy Program and the X-33 Advanced Vehicle – 40 Technology Demonstrator – 128 SYSTEMS HEALTH MONITORING Improvements in Thermal Protection Siz- TECHNOLOGY ASSESSMENT Active Wireless Temperature Sensors for ing Capabilities for TCAT: Conceptual A Damage Resistance Comparison Be- Design for Advanced Space Transporta- Aerospace Thermal Protection Sys- tems – 46 tween Candidate Polymer Matrix Com- tion Systems – 61 posite Feedline Materials – 88 Integrated Technology Assessment Cen- Advanced Propulsion System Test and Tell Capability – 133 A guidance and control assessment of ter (ITAC) Update – 42 three vertical landing options for Multi-Disciplinary Analysis for Future Development of Structural Health Man- RLV – 136 agement Technology for Aerospace Ve- Launch Systems Using NASA’s Ad- Advanced Transportation System Stud- hicles – 1 vanced Engineering Environment ies Technical Area 2 (TA-2) Heavy Lift (AEE) – 17 Flight Demonstration of X-33 Vehicle Launch Vehicle Development Con- NASA’s Hypersonic Investment Health Management System Compo- tract – 129 Area – 47 nents on the F/A-18 Systems Research Affordable In-Space Transportation Aircraft – 15 Phase 2: An Advanced Concepts NASA’s Spaceliner Investment Area Health Monitoring Technology for Ther- Project – 113 Technology Activities – 81 mal Protection Systems on Reusable Hy- An Approach to Establishing System Orbital Space Plane (OSP) Pro- personic Vehicles – 71 Benefits for Technology in NASA’s Hy- gram – 5 NASA EPSCoR Preparation Grant – 49 personics Investment Area – 67 Progress in Unsteady Turbopump Flow Real-Time Sensor Validation, Signal Re- Composite,Cryogenic, Conformal, Com- Simulations – 51 construction, and Feature Detection for mon Bulkhead, Aerogel-Insulated Tank Research Reports: 1996 NASA/ASEE an RLV Propulsion Testbed – 98 (CBAT) – 74 Summer Faculty Fellowship Pro- Research Reports: 1996 NASA/ASEE Design, Development,and Testing of Um- gram – 102 Summer Faculty Fellowship Pro- billical System Mechanisms for the X-33 System Sensitivity Analysis Applied to gram – 102 Advanced Technology Demonstra- the Conceptual Design of a Dual-Fuel tor – 90 The NASA Integrated Vehicle Health Rocket SSTO – 125 Management Technology Experiment for Integrated Technology Assessment Cen- Technology Roadmap for Dual-Mode X-37 – 53 ter (ITAC) Update – 42 Scramjet Propulsion to Support Space- SYSTEMS INTEGRATION JTEC panel report on space and transat- Access Vision Vehicle Develop- mospheric propulsion technology – 175 ment – 38 Airbreathing/Rocket Single-Stage-to- Orbit Design Matrix – 124 Launch Vehicle Sizing Benefits Utilizing SYSTEMS ENGINEERING Main Propulsion System Crossfeed and Assessment of flying-quality criteria for Design and Calibration of the X-33 Flush Project Status – 34 air-breathing aerospacecraft – 167 Airdata Sensing (FADS) System – 113 Lightweight Chambers for Thrust Assem- Design and Calibration of the X-33 Flush Development of Structural Health Man- blies – 82 Airdata Sensing (FADS) System – 113 agement Technology for Aerospace Ve- NASA IVHM Technology Experiment for hicles – 1 Flight Demonstration of X-33 Vehicle X-vehicles (NITEX) – 67 Health Management System Compo- Human Transportation System (HTS) nents on the F/A-18 Systems Research Optimal technology investment strate- study, volume 2 – 155 Aircraft – 15 gies for a reusable launch vehicle – 134 Improvements in Thermal Protection Siz- Integrated Technology Assessment Cen- Recent Advances and Applications in ing Capabilities for TCAT: Conceptual ter (ITAC) Update – 42 Cryogenic Propellant Densification Tech- Design for Advanced Space Transporta- nology – 89 tion Systems – 61 Research Reports: 1996 NASA/ASEE Summer Faculty Fellowship Pro- Single-stage-to-orbit: Meeting the chal- Life Cycle Systems Engineering Ap- gram – 102 lenge – 136 proach to NASA’s 2nd Generation Reus- able Launch Vehicle – 39 Survey of NASA V and V Support to X-33/Reusable Launch Ve- Processes/Methods – 58 hicle Technology Program – 86 Multi-Disciplinary Analysis for Future Launch Systems Using NASA’s Ad- Test Results for an Off-the-Shelf Systems Engineering Approach to Tech- vanced Engineering Environment GPS/INS During Approach and Landing nology Integration for NASA’s 2nd Gen- (AEE) – 17 Testing of the X-40A – 63 eration Reusable Launch Vehicle – 24

A-42 Task 4 supporting technology. Part 2: TEMPERATURE CONTROL Experimental Evaluation of a Subscale Detailed test plan for thermal seals. Ther- A Thermal Management Systems Model Gaseous Hydrogen/Gaseous Oxygen mal seals evaluation, improvement and for the NASA GTX RBCC Concept – 53 Coaxial Rocket Injector – 34 test. CAN8-1, Reusable Launch Vehicle Current Technology for Thermal Protec- High Heat Flux Facility – 151 (RLV), advanced technology demonstration Systems – 165 tor: X-33. Leading edge and seals ther- John C. Stennis Space Center over- mal protection system technology dem- Thermal Management Design for the view – 152 139 104 onstration – X-33 Lifting Body – Liquid Engine Test Facilities Assess- The NASA Integrated Vehicle Health TEMPERATURE DISTRIBUTION ment – 1 Management Technology Experiment for Aerothermoelastic analysis of a NASP Overview of GRC’s Advanced Sensor X-37 – 53 demonstrator model – 155 and Instrumentation Development – 41 TECHNOLOGY TRANSFER TEMPERATURE EFFECTS Slush hydrogen technology pro- Research Reports: 1996 NASA/ASEE Acoustic Emission Monitoring of the gram – 149 Summer Faculty Fellowship Pro- DC-XA Composite Liquid Hydrogen Tank X-33 Integrated Test Facility Extended gram – 102 126 During Structural Testing – Range Simulation – 100 TECHNOLOGY UTILIZATION Hot dynamic test rig for measuring hyper- Ceramic Matrix Composites (CMC) Life sonic engine seal flow and durabil- TEST FIRING Prediction Method Development – 88 ity – 151 Advanced Propulsion System Test and Tell Capability – 133 Characterization of Polyimide Foams for Investigation of Fundamental Modeling Ultra-Lightweight Space Struc- and Thermal Performance Issues for a TEST STANDS tures – 22 Metallic Thermal Protection System De- Propulsion module design review. Mile- Composite,Cryogenic, Conformal, Com- sign – 33 stone 8: Integrated Propulsion Technol- ogy Demonstrator – 150 mon Bulkhead, Aerogel-Insulated Tank TEMPERATURE MEASUREMENT (CBAT) – 74 Active Wireless Temperature Sensors for TETHERING Development of Structural Health Man- Aerospace Thermal Protection Sys- Tether Transportation System agement Technology for Aerospace Ve- tems – 46 Study – 109 hicles – 1 Infrared Sensing Aeroheating Flight Ex- THERMAL ANALYSIS Life Cycle Systems Engineering Ap- periement: STS-96 Flight Results – 84 Advanced High Temperature Structural proach to NASA’s 2nd Generation Reus- Seals – 43 able Launch Vehicle – 39 TEMPERATURE SENSORS Active Wireless Temperature Sensors for Lightweight Chambers for Thrust Assem- Evaluation of Advanced Thermal Protec- Aerospace Thermal Protection Sys- blies – 82 tion Techniques for Future Reusable tems – 46 Launch Vehicles – 62 NASA 2nd Generation RLV Program In- Reusable Cryogenic Tank VHM Using Investigation of Fundamental Modeling troduction, Status and Future Fiber Optic Distributed Sensing Technol- and Thermal Performance Issues for a Plans – 31 ogy – 108 Metallic Thermal Protection System De- NASA’s Advanced Space Transportation sign – 33 Program: RTA Project Summary – 42 TEMPERATURE Advanced Control Surface Seal Develop- Lightweight durable TPS: Milestone NASA’s Spaceliner Investment Area ment at NASA GRC for 3rd Generation 4–151 Technology Activities – 81 RLV – 29 Preliminary Thermal-Mechanical Sizing Natural Atmospheric Environment Model TENSILE STRESS of Metallic TPS: Process Development Development for the National Aeronau- and Sensitivity Studies – 33 tics and Space Administration’s Second Development of a Pressure Box to Evalu- Statistical Methods for Rapid Aerother- Generation Reusable Launch Ve- ate Reusable-Launch-Vehicle mal Analysis and Design Technol- hicle – 57 Cryogenic-Tank Panels – 132 ogy – 55 Optimal technology investment strate- TERMINAL AREA ENERGY MANAGE- Thermal/Fluid Analysis of a Composite gies for a reusable launch vehicle – 134 MENT Adaptive Critic Neural Network-Based Heat Exchanger for Use on the RLV Pathfinder Technologies Specialist, Terminal Area Energy Management and Rocket Engine – 37 X-37 – 69 Approach and Landing Guidance – 9 THERMAL BUCKLING Space Launch Initiative: New Capabili- Dynamics and Stability and Control Effects of Stiffening and Mechanical ties - New Horizons – 38 Characteristics of the X-37 – 62 Load on Thermal Buckling of Stiffened TEFLON (TRADEMARK) Fuzzy Logic Trajectory Design and Guid- Cylindrical Shells – 121 Development of a Pressure Box to Evalu- ance for Terminal Area Energy Manage- THERMAL CONDUCTIVITY ate Reusable-Launch-Vehicle ment – 8 Effective Thermal Conductivity of High Cryogenic-Tank Panels – 133 TEST CHAMBERS Temperature Insulations for Reusable 97 TELECOMMUNICATION A study of facilities and fixtures for testing Launch Vehicles – Analysis of space concepts enabled by of a high speed civil transport wing com- THERMAL CONTROL COATINGS new transportation (ASCENT) ponent – 127 Evaluation of Thermal Control Coatings study – 28 TEST FACILITIES for Flexible Ceramic Thermal Protection 117 X-33 Integrated Test Facility Extended A study of facilities and fixtures for testing Systems – Range Simulation – 100 of a high speed civil transport wing com- THERMAL CYCLING TESTS ponent – 127 TELEROBOTICS Hot dynamic test rig for measuring hyper- Mosad and Stream Vision For A Telero- Engineering the Future of Full-Scale Pro- sonic engine seal flow and durabil- botic, Flying Camera System – 59 pulsion Testing – 7 ity – 151

A-43 Milestone 4: Test plan for Reusable Hy- Health Monitoring Technology for Ther- Thermal Structures Technology Develop- drogen Composite Tank System mal Protection Systems on Reusable Hy- ment for Reusable Launch Vehicle Cryo- (RHCTS). Task 3: Composite tank mate- personic Vehicles – 71 genic Propellant Tanks – 107 rials – 142 Hypervelocity Impact Test Results for a Wireless Subsurface Microsensors for Thermal-Mechanical Cyclic Test of a Metallic Thermal Protection Sys- Health Monitoring of Thermal Protection Composite Cryogenic Tank for Reusable tem – 10 Systems on Hypersonic Vehicles – 69 Launch Vehicles – 23 Improvements in Thermal Protection Siz- Wireless Subsurface Sensors for Health ing Capabilities for TCAT: Conceptual Monitoring of Thermal Protection Sys- THERMAL INSULATION Design for Advanced Space Transporta- tems on Hypersonic Vehicles – 68 Current Technology for Thermal Protec- tion Systems – 61 tion Systems – 165 THERMAL STRESSES Investigation of Fundamental Modeling Development of a Pressure Box to Evalu- Effective Thermal Conductivity of High and Thermal Performance Issues for a ate Reusable-Launch-Vehicle Temperature Insulations for Reusable Metallic Thermal Protection System De- Cryogenic-Tank Panels – 132 Launch Vehicles – 97 sign – 33 Development of an integrated BEM ap- Reusable Cryogenic Tank VHM Using Lightweight durable TPS: Milestone proach for hot fluid structure interac- Fiber Optic Distributed Sensing Technol- 4–151 tion – 169 ogy – 108 Materials Testing on the DC-X and DC- National Aerospace Plane Thermal De- Task 4 supporting technology. Part 2: XA – 121 velopment. (Latest Citations from the Detailed test plan for thermal seals. Ther- Aerospace Database) – 113 mal seals evaluation, improvement and Panel Flutter and Sonic Fatigue Analysis THERMOCHEMICAL PROPERTIES test. CAN8-1, Reusable Launch Vehicle for RLV – 77 Surface Catalysis and Characterization (RLV), advanced technology demonstra- Parametric Weight Comparison of Ad- of Proposed Candidate TPS for Access- tor: X-33. Leading edge and seals ther- vanced Metallic, Ceramic Tile, and Ce- to-Space Vehicles – 118 mal protection system technology dem- ramic Blanket Thermal Protection Sys- onstration – 139 tems – 88 THERMODYNAMIC CYCLES Test Plans. Lightweight Durable TPS: Preliminary Thermal-Mechanical Sizing Advanced Low-Cost O2/H2 Engines for Tasks 1,2,4,5, and 6 – 146 of Metallic TPS: Process Development the SSTO Application – 123 33 The Evolution of Flexible Insulation as and Sensitivity Studies – THERMODYNAMIC PROPERTIES Thermal Protection Systems for Reus- Requirements report for SSTO vertical National Aerospace Plane Thermal De- able Launch Vehicles: AFRSI (Advanced take-off and horizontal landing ve- velopment. (Latest citations from the Flexible Reusable Surface Insulation) to hicle – 143 Aerospace Database) – 114 CRI (Conformal Reusable Insula- Results of Two-Stage Light-Gas Gun De- THERMODYNAMICS tion) – 67 velopment Efforts and Hypervelocity Im- Preliminary Thermal-Mechanical Sizing of Metallic TPS: Process Development THERMAL PROTECTION pact Tests of Advanced Thermal Protec- tion Materials – 100 and Sensitivity Studies – 33 A Collaborative Analysis Tool for Thermal Protection Systems for Single Stage to Reusable launch vehicle: Technology de- THERMOELASTICITY Orbit Launch Vehicles – 56 velopment and test program – 135 Development of an integrated BEM approach for hot fluid structure interac- Surface Catalysis and Characterization A Method of Integrating Aeroheating into tion – 169 Conceptual Reusable Launch Vehicle of Proposed Candidate TPS for Access- Design: Evaluation of Advanced Thermal to-Space Vehicles – 118 THERMOGRAPHY Protection Techniques for Future Reus- Surface Catalytic Efficiency of Advanced Thermographic Analysis of Composite able Launch Vehicles – 77 Carbon Carbon Candidate Thermal Pro- Cobonds on the X-33 – 75 Advanced Metallic Thermal Protection tection Materials for SSTO Ve- X-33 Turbulent Aeroheating Measure- System Development – 64 hicles – 133 ments and Predictions – 32 Advances in hypersonic vehicle synthe- Task 4 supporting technology. Part 1: THERMOPLASTICITY sis with application to studies of ad- Detailed test plan for leading edge tile Thermoviscoplastic response of Ti-15-3 vanced thermal protection sys- development. Leading edge material de- under various loading conditions – 172 tem – 138 velopment and testing – 139 THIN FILMS Task 4 supporting technology. Part 2: Configuration development study of the Characterization of Thin Film Polymers Detailed test plan for thermal seals. Ther- OSU 1 hypersonic research ve- Through Dynamic Mechanical Analysis mal seals evaluation, improvement and hicle – 154 and Permeation – 22 test. CAN8-1, Reusable Launch Vehicle Current Technology for Thermal Protec- (RLV), advanced technology demonstra- Heat flux sensor research and develop- tion Systems – 165 tor: X-33. Leading edge and seals ther- ment: The cool film calorimeter – 170 mal protection system technology dem- Development of Metallic Thermal Protec- THREE DIMENSIONAL FLOW onstration – 139 tion Systems for the Reusable Launch Analysis of Aerospike Plume Induced Vehicle – 124 Thermal Management Design for the Base-Heating Environment – 109 X-33 Lifting Body – 104 Development of Processing Techniques Computational Aerodynamics Analysis of for Advanced Thermal Protection Materi- Thermal Protection Materials for Reentry Future Launch Vehicle Configura- als – 112 and Planetary Applications – 68 tions – 122 Director’s Discretionary Fund Report for Thermal Protection Materials for Reentry Numerical investigations in three- Fiscal Year 1996 – 120 Applications – 59 dimensional internal flows – 171 Evaluation of Thermal Control Coatings Thermal Protection System Evaluation PNS predictions of external/internal hy- for Flexible Ceramic Thermal Protection Using Arc-jet Flows: Flight Simulation or personic flows for NASP propulsion ap- Systems – 117 Research Tool? – 31 plications – 157

A-44 THREE DIMENSIONAL MODELS TILES TRAJECTORY CONTROL Analysis of X-33 Linear Aerospike Plume Parametric Weight Comparison of Ad- Air-breathing hypersonic vehicle guid- Induced Base-Heating Environment Dur- vanced Metallic, Ceramic Tile, and Ce- ance and control studies; An integrated ing Power-Pack Out – 29 ramic Blanket Thermal Protection Sys- trajectory/control analysis methodology: tems – 88 Phase 1 – 171 On-Board Generation of Three- Dimensional Constrained Entry Trajecto- Task 4 supporting technology. Part 1: Air-breathing hypersonic vehicle guid- ries – 35 Detailed test plan for leading edge tile ance and control studies: An integrated development. Leading edge material de- trajectory/control analysis methodology, THROTTLING velopment and testing – 139 phase 2 – 165 The Control System for the X-33 Linear An inverse dynamics approach to trajec- Aerospike Engine – 105 TILT ROTOR AIRCRAFT tory optimization for an aerospace Director’s Discretionary Fund Report for plane – 163 THRUST AUGMENTATION Fiscal Year 1996 – 120 Magnetohydrodynamic Augmented Pro- Genetic Algorithm Tuned Fuzzy Logic for pulsion Experiment – 52 TIME DEPENDENCE Gliding Return Trajectories – 12 Time Dependent Simulation of Tur- Operationally efficient propulsion system Rapid near-optimal aerospace plane tra- bopump Flows – 55 study (OEPSS) data book – 110 jectory generation and guidance – 171 TIME LAG Rocket-Induced Magnetohydrodynamic TRAJECTORY OPTIMIZATION Structural Seal Development – 11 Ejector: A Single-Stage-to-Orbit Ad- A parametric sensitivity study for single- vanced Propulsion Concept – 131 TIME TEMPERATURE PARAMETER stage-to-orbit hypersonic vehicles using Thrust augmentation options for the Beta Near-Optimal Re-Entry Trajectories for trajectory optimization – 153 2 two-stage-to-orbit vehicle – 154 Reusable Launch Vehicles – 115 Advances in hypersonic vehicle synthesis with application to studies of ad- THRUST CHAMBERS TITANIUM ALLOYS vanced thermal protection sys- An Experimental Study of Characteristic Materials Research for High Speed Civil tem – 138 Combustion-Driven Flow for CFD Valida- Transport and Generic Hypersonics- tion – 103 Metals Durability – 127 An Automated Method to Compute Or- bital Re-entry Trajectories with Heating Development of Aluminum Composites Thermoviscoplastic response of Ti-15-3 Constraints – 39 for a Rocket Engine’s Lightweight Thrust under various loading conditions – 172 Cell – 80 An Automated Method to Compute Or- TOTAL QUALITY MANAGEMENT bital Re-Entry Trajectories with Heating Lightweight Chambers for Thrust Assem- A Collection of Technical Papers – 122 Constraints – 53 blies – 66 Design of launch systems using continu- An inverse dynamics approach to trajec- THRUST CONTROL ous improvement process – 134 tory optimization for an aerospace The Control System for the X-33 Linear plane – 163 TOWERS Aerospike Engine – 105 Analytical solutions to constrained hyper- High Altitude Launch for a Practical sonic flight trajectories – 162 THRUST VECTOR CONTROL SSTO – 29 Effect of Aeropropulsive Interactions and The Control System for the X-33 Linear Towers for Earth Launch – 29 Aerospike Engine – 105 Design Sensitivities on Optimal Hyper- TRAJECTORIES sonic Ascent Trajectories – 125 THRUSTORS Air-breathing hypersonic vehicle guid- Design and Testing of Non-Toxic RCS Near-Optimal Operation of Dual-Fuel ance and control studies; An integrated Thrusters for Second Generation Reus- Launch Vehicles – 119 trajectory/control analysis methodology: able Launch Vehicle – 18 Phase 1 – 171 Optimal trajectories for an aerospace plane. Part 2: Data, tables, and THRUST Fuzzy Logic Trajectory Design and Guid- graphs – 176 Design and Testing of Non-Toxic RCS ance for Terminal Area Energy Manage- Thrusters for Second Generation Reus- ment – 8 Rapid near-optimal aerospace plane tra- able Launch Vehicle – 14 jectory generation and guidance – 171 On-Board Generation of Three- Improved Re-Configurable Sliding Mode Dimensional Constrained Entry Trajecto- Rapid near-optimal trajectory generation Controller for Reusable Launch Vehicle ries – 35 and guidance law development for of Second Generation Addressing Aero- single-stage-to-orbit airbreathing ve- Optimal trajectories for an aerospace dynamic Surface Failures and Thrust De- hicles – 177 plane. Part 1: Formulation, results, and ficiencies – 34 analysis – 176 Real-time trajectory optimization on par- Thrust Enhancement in Hypervelocity allel processors – 158 Nozzles by Chemical Catalysis – 102 Parametric Weight Comparison of Ad- vanced Metallic, Ceramic Tile, and Ce- SRM-Assisted Trajectory for the GTX Reference Vehicle – 48 THRUST-WEIGHT RATIO ramic Blanket Thermal Protection Sys- Selection process for trade study: Graph- tems – 88 Trajectory optimization for the national aerospace plane – 156 ite Composite Primary Structure TRAJECTORY ANALYSIS (GCPS) – 148 Air-breathing hypersonic vehicle guid- Trajectory optimization for the National Test Plan. GCPS Task 4, subtask 4.2 ance and control studies: An integrated aerospace plane – 160 thrust structure development – 147 trajectory/control analysis methodology, TRAJECTORY PLANNING phase 2 – 165 The rationale/benefits of nuclear thermal A parametric sensitivity study for single- rocket propulsion for NASA’s lunar space Analytical solutions to constrained hyper- stage-to-orbit hypersonic vehicles using transportation system – 141 sonic flight trajectories – 162 trajectory optimization – 153 Trade study plan for Graphite Composite High Altitude Launch for a Practical On-Board Entry Trajectory Planning Ex- Primary Structure (GCPS) – 146 SSTO – 29 panded to Sub-orbital Flight – 12

A-45 Robust intelligent flight control for hyper- TURBULENCE MODELS UNSTRUCTURED GRIDS (MATHEMAT- sonic vehicles – 166 Hypersonic flows as related to the Na- ICS) TRANSATMOSPHERIC VEHICLES tional Aerospace Plane – 136 An Inviscid Computational Study of an X-33 Configuration at Hypersonic Access to space – 143 Turbulence modeling for hypersonic Speeds – 94 161 Conceptual design of two-stage-to-orbit flight – Computational Study of a McDonnell hybrid launch vehicle – 169 TURBULENT BOUNDARY LAYER Douglas Single-Stage-to-Orbit Vehicle Effects of Stiffening and Mechanical Infrared Sensing Aeroheating Flight Ex- Concept for Aerodynamic Analy- Load on Thermal Buckling of Stiffened periment: STS-96 Flight Results – 86 sis – 126 Cylindrical Shells – 121 TURBULENT FLOW UPLINKING Hyper-X Wind Tunnel Program – 109 Computational/Experimental Aeroheat- ing Predictions for X-33 Phase 2 Ve- X-33 Integrated Test Facility Extended JTEC panel report on space and transat- hicle – 110 Range Simulation – 100 mospheric propulsion technology – 175 Computational/Experimental Aeroheat- USER MANUALS (COMPUTER PRO- TRANSFERRING ing Predictions for X-33 – 112 GRAMS) Slush hydrogen transfer studies at the A Procedure for Structural Weight Esti- NASA K-Site Test Facility – 167 Director’s Discretionary Fund Report for Fiscal Year 1996 – 120 mation of Single Stage to Orbit Launch TRANSITION FLOW Vehicles (Interim User’s Manual) – 44 X-33 Rev-F Turbulent Aeroheating Re- Computational/Experimental Aeroheat- VELOCITY MEASUREMENT ing Predictions for X-33 Phase 2 Ve- sults From Test 6817 in NASA Langley hicle – 110 20-Inch Mach 6 Air Tunnel and Compari- An Experimental Study of Characteristic sons With Computations – 26 Combustion-Driven Flow for CFD Valida- TRANSONIC SPEED 115 X-33 Turbulent Aeroheating Measure- tion – Computation of H2/air reacting flowfields ments and Predictions – 32 in drag-reduction external combus- VENTURESTAR LAUNCH VEHICLE tion – 161 TWO DIMENSIONAL FLOW An Inviscid Computational Study of an X-33 Configuration at Hypersonic Transonic aerodynamic characteristics of PNS predictions of external/internal hy- Speeds – 94 a proposed wing-body reusable launch personic flows for NASP propulsion applications – 157 vehicle concept – 138 VERTICAL LANDING TRANSPORTATION TWO PHASE FLOW A guidance and control assessment of Analysis of space concepts enabled by Current Technology for Thermal Protec- three vertical landing options for new transportation (ASCENT) tion Systems – 165 RLV – 136 study – 28 ULLAGE Advanced Transportation System Stud- TRAVEL Prediction of the ullage gas thermal ies Technical Area 2 (TA-2) Heavy Lift General Public Space Travel and Tour- stratification in a NASP vehicle propellant Launch Vehicle Development Con- ism – 96 tank experimental simulation using tract – 129 FLOW-3D – 177 TURBINE BLADES Aerodynamic Characteristics of a Vertical UNIVERSITY PROGRAM Experimental Performance Evaluation of Takeoff Vertical Landing (VTVL) Single a Supersonic Turbine for Rocket Engine Research Reports: 1996 NASA/ASEE Stage to Orbit Vehicle from M(sub infin- Applications – 10 Summer Faculty Fellowship Pro- ity) = 0.1 to 10 – 122 gram – 102 TURBINE ENGINES Design, Development,and Testing of Um- The 1994 NASA/USRA/ADP Design billical System Mechanisms for the X-33 NASA’s Advanced Space Transportation Projects – 140 Advanced Technology Demonstra- Program: RTA Project Summary – 42 The Center for Aerospace Research: A tor – 90 TURBINE PUMPS NASA Center of Excellence at North VERTICAL TAKEOFF AIRCRAFT Progress in Unsteady Turbopump Flow Carolina Agricultural and Technical State Preliminary Assessment of Using Gelled Simulations Using Overset Grid Sys- University – 162 tems – 30 and Hybrid Propellant Propulsion for UNSTEADY AERODYNAMICS VTOL/SSTO Launch Systems – 104 Progress in Unsteady Turbopump Flow A method of predicting quasi-steady Simulations – 51 aerodynamics for flutter analysis of high VERTICAL TAKEOFF Space Shuttle Main Engine Off-Nominal speed vehicles using steady CFD calcu- Advanced Transportation System Stud- Low Power Level Operation – 117 lations – 157 ies Technical Area 2 (TA-2) Heavy Lift Launch Vehicle Development Con- Time Dependent Simulation of Tur- An overview of aeroelasticity studies for tract – 129 bopump Flows – 55 the National Aerospace Plane – 160 Time-Dependent Simulations of Tur- NASP aeroservothermoelasticity stud- Aerodynamic Characteristics of a Vertical bopump Flows – 38 ies – 174 Takeoff Vertical Landing (VTVL) Single Stage to Orbit Vehicle from M(sub infin- Unshrouded Impeller Technology Task UNSTEADY FLOW ity) = 0.1 to 10 – 122 Status – 81 Progress in Unsteady Turbopump Flow Analysis of the staging maneuver and TURBOFANS Simulations Using Overset Grid Sys- booster glideback guidance for a two- The future challenge for aeropropul- tems – 30 stage, winged, fully reusable launch ve- sion – 167 Progress in Unsteady Turbopump Flow hicle – 159 TURBOMACHINERY Simulations – 51 Highly Reusable Space Transportation Design and Analysis of Turbines for Time Dependent Simulation of Tur- (HRST) Baseline Concepts and Analysis: Space Applications – 16 bopump Flows – 55 Rocket/RBCC Options – 114 Overview of MSFC’s Applied Fluid Dy- Time-Dependent Simulations of Tur- Magnetohydrodynamic Augmented Pro- namics Analysis Group Activities – 24 bopump Flows – 38 pulsion Experiment – 52

A-46 VIBRATION Research Reports: 1996 NASA/ASEE WIND VELOCITY MEASUREMENT An implementation of a chemical and Summer Faculty Fellowship Pro- A Compendium of Wind Statistics and thermal nonequilibrium flow solver on gram – 102 Models for the NASA Space Shuttle and unstructured meshes and application to Other Aerospace Vehicle Pro- blunt bodies – 145 WIND EFFECTS grams – 99 A Compendium of Wind Statistics and VIDEO TAPES Stratospheric turbulence measurements Models for the NASA Space Shuttle and Operationally Efficient Propulsion Sys- and models for aerospace plane de- Other Aerospace Vehicle Pro- sign – 164 tem Study (OEPSS): OEPSS Video grams – 99 Script – 110 WING OSCILLATIONS VISCOPLASTICITY WIND PROFILES Evaluation of Computational Aeroelastic- Thermoviscoplastic response of Ti-15-3 A Compendium of Wind Statistics and ity Codes for Loads and Flutter – 82 under various loading conditions – 172 Models for the NASA Space Shuttle and WINGED VEHICLES Other Aerospace Vehicle Pro- VISCOUS DRAG Abort performance for a winged-body grams – 99 137 Reduction of Base Drag on Launch Ve- single-stage to orbit vehicle – hicles – 50 WIND TUNNEL MODELS WINGS VISCOUS FLOW Access to space – 143 A method of predicting quasi-steady aerodynamics for flutter analysis of high An Inviscid Computational Study of an Subsonic Aerodynamic Characteristics of speed vehicles using steady CFD calcu- X-33 Configuration at Hypersonic a Circular Body Earth-to-Orbit Ve- lations – 157 Speeds – 94 hicle – 128 A study of facilities and fixtures for testing Comparison of Methods for Determining of a high speed civil transport wing com- Boundary Layer Edge Conditions for WIND TUNNEL TESTS ponent – 127 Transition Correlations – 23 Computational/Experimental Aeroheat- ing Predictions for X-33 – 112 Heat pipes for wing leading edges of Development of an integrated BEM ap- hypersonic vehicles – 178 proach for hot fluid structure interac- Discrete Roughness Effects on Shuttle tion – 169 Orbiter at Mach 6 – 37 Wingless Flight: The Lifting Body Story – 105 VOIDS Flight Stability and Control and Perfor- Thermographic Analysis of Composite mance Results from the Linear Aerospike WIRELESS COMMUNICATION Cobonds on the X-33 – 75 SR-71 Experiment (LASRE) – 101 Wireless Subsurface Microsensors for Health Monitoring of Thermal Protection WATER FLOW Hypersonic Boundary-Layer Transition Systems on Hypersonic Vehicles – 69 Water Flow Performance of a Superscale for X-33 Phase 2 Vehicle – 111 X RAY ANALYSIS Model of the Fastrac Liquid Oxygen Hypersonic flows as related to the Na- Development of Processing Techniques Pump – 55 tional Aerospace Plane – 136 for Advanced Thermal Protection Materi- WATER INJECTION NASP aeroservothermoelasticity stud- als – 112 Thrust augmentation options for the Beta ies – 174 X RAY SOURCES 2 two-stage-to-orbit vehicle – 154 Subsonic Aerodynamic Characteristics of Advanced High Brilliance X-Ray WAVERIDERS a Circular Body Earth-to-Orbit Ve- Source – 111 An integrated aerodynamic/propulsion 128 hicle – X-29 AIRCRAFT study for generic aero-space planes based on waverider concepts – 173 Subsonic aerodynamic characteristics of NASA aeronautics: Research and tech- a proposed advanced manned launch nology program highlights – 164 WEIGHT REDUCTION system orbiter configuration – 160 X-30 VEHICLE Development of Aluminum Composites Supersonic aeroelastic instability results A study of facilities and fixtures for testing for a Rocket Engine’s Lightweight Thrust for a NASP-like wing model – 159 of a high speed civil transport wing com- Cell – 80 ponent – 127 Dual-Fuel Propulsion in Single-Stage Ad- Transonic aerodynamic characteristics of Aircraft Emission Scenarios Projected in vanced Manned Launch System Ve- a proposed wing-body reusable launch 138 Year 2015 for the NASA Technology Con- hicle – 125 vehicle concept – cept Aircraft (TCA) High Speed Civil Wind-Tunnel Investigations of Blunt- WELD TESTS Transport – 108 Body Drag Reduction Using Forebody Fatigue Crack Growth Rate Test Results Surface Roughness – 85 An overview of aeroelasticity studies for for Al-Li 2195 Parent Metal, Variable Po- the National Aerospace Plane – 160 larity Plasma Arc Welds and Friction Stir X-33 (Rev-F) Aeroheating Results of Test Welds – 89 6770 in NASA Langley 20-Inch Mach 6 Shuttle to Space Station. Heart assist Air Tunnel – 96 implant. Hubble update. X-30 mock- WELDABILITY up – 138 X-33 Rev-F Turbulent Aeroheating Re- Selection and Evaluation of An Alloy for X-31 AIRCRAFT Nozzle Application – 22 sults From Test 6817 in NASA Langley 20-Inch Mach 6 Air Tunnel and Compari- Aero-Space Plane: Flexible access to WELDED JOINTS sons With Computations – 26 space – 150 Fatigue Crack Growth Rate Test Results X-33 Turbulent Aeroheating Measure- X-33 REUSABLE LAUNCH VEHICLE for Al-Li 2195 Parent Metal, Variable Po- ments and Predictions – 32 A Strategy for Integrating a Large Finite larity Plasma Arc Welds and Friction Stir Element Model Using MSC Welds – 89 WIND TUNNELS NASTRAN/PATRAN: X-33 Lessons Learned – 97 WELDING Winds of change: Expanding the frontiers Development of Metal Matrix Compos- of flight. Langley Research Center’s 75 An Experimental Study of Characteristic ites for NASA’s Advanced Propulsion years of accomplishment, 1917- Combustion-Driven Flow for CFD Valida- Systems – 74 1992 – 165 tion – 103

A-47 An Improved RLV Stability Analysis Via a N-Factor Computations for the X-33 Ve- X-33 (Rev-F) Aeroheating Results of Test Continuation Approach – 9 hicle – 76 6770 in NASA Langley 20-Inch Mach 6 Air Tunnel – 96 An Inviscid Computational Study of an Predictor-Corrector Entry Guidance for X-33 Configuration at Hypersonic Reusable Launch Vehicles – 74 X-33 Rev-F Turbulent Aeroheating Re- Speeds – 94 sults From Test 6817 in NASA Langley Reconfigurable Control Design for the 20-Inch Mach 6 Air Tunnel and Compari- Analysis of Aerospike Plume Induced Full X-33 Flight Envelope – 72 sons With Computations – 26 Base-Heating Environment – 109 Reconfigurable Flight Control Designs X-33 Tank Failure During Autoclave Fab- Analysis of X-33 Linear Aerospike Plume With Application to the X-33 Ve- rication – 60 Induced Base-Heating Environment Dur- hicle – 95 ing Power-Pack Out – 29 X-33 XRS-2200 Linear Aerospike Engine Reusable Cryogenic Tank VHM Using Sea Level Plume Radiation – 66 Base-Bleed Effect on X-33 Aerospike Fiber Optic Distributed Sensing Technol- X-33/RLV System Health Management/ Plume Induced Base-Heating Environ- ogy – 108 ment During Power-Pack Out – 25 Vehicle Health Management – 107 Support to X-33/Resusable Launch Ve- X-33 – 105 Ceramic Matrix Composite Cooled hicle Technology Program – 86 Nozzle Material Development Pro- X-34 REUSABLE LAUNCH VEHICLE Support to X-33/Reusable Launch Ve- gram – 21 Options for flight testing rocket-based hicle Technology Program – 86 Computational/Experimental Aeroheat- combined-cycle (RBCC) engines – 130 ing Predictions for X-33 Phase 2 Ve- System Identification of X-33 Neural Net- X-37 VEHICLE work – 8 hicle – 110 Dynamics and Stability and Control Computational/Experimental Aeroheat- Testing the Preliminary X-33 Navigation Characteristics of the X-37 – 62 ing Predictions for X-33 – 112 System – 84 Flight Demonstrations of Orbital Space Design and Calibration of the X-33 Flush The Control System for the X-33 Linear Plane (OSP) Technologies – 16 Airdata Sensing (FADS) System – 113 Aerospike Engine – 105 Pathfinder Technologies Specialist, Design, Development,and Testing of Um- The Reusable Launch Vehicle Technol- X-37 – 69 billical System Mechanisms for the X-33 ogy Program and the X-33 Advanced Second Generation RLV Program – 71 128 Advanced Technology Demonstra- Technology Demonstrator – Test Results for an Off-the-Shelf tor – 90 The X-33 Extended Flight Test GPS/INS During Approach and Landing Development of the X-33 Aerodynamic Range – 99 Testing of the X-40A – 63 Uncertainty Model – 108 Thermal Management Design for the The NASA Integrated Vehicle Health Evaluation of Microcracking in Two X-33 Lifting Body – 104 Management Technology Experiment for Carbon-Fiber/Epoxy-Matrix Composite X-37 – 53 Cryogenic Tanks – 72 Third Generation RLV Structural Seal Development Programs at NASA Update on the Development and Capa- Findings from the X-33 Hydrogen Tank GRC – 41 bilities of Unique Structural Seal Test Failure Investigation – 60 Rigs – 3 Two Reconfigurable Flight-Control De- Flight Demonstration of X-33 Vehicle X-37 C-SiC CMC Control Surface Com- sign Methods: Robust Servomechanism ponents Development: Status of the Health Management System Compo- and Control Allocation – 76 nents on the F/A-18 Systems Research NASA/Boeing/USAF Orbital Vehicle and Aircraft – 15 Unic Unstructured CFD Methodology De- Related Efforts – 20 velopment – 82 Flight Testing the Linear Aerospike X-38 CREW RETURN VEHICLE SR-71 Experiment (LASRE) – 100 X-33 Integrated Test Facility Extended Update on the Development and Capa- Range Simulation – 100 bilities of Unique Structural Seal Test Full Envelope Reconfigurable Control Rigs – 3 72 Design for the X-33 Vehicle – X-33 LH2 Tank Failure Investigation ZIRCONIUM Findings – 14 Infrared Sensing Aeroheating Flight Ex- Tantalum Addition to Zirconium Diboride periment: STS-96 Flight Results – 86 X-33 Phase 2 – 115 for Improved Oxidation Resistance – 18

A-48 Personal Author Index

Aarnio, Steve Andrews, Jason Auslender, Aaron H. Project Antares: A low cost modular The StarBooster System: A Cargo Air- Technology Roadmap for Dual-Mode launch vehicle for the future – 168 craft for Space – 92 Scramjet Propulsion to Support Space- Access Vision Vehicle Develop- Abtahi, A. Ankney, W. ment – 38 Heat flux sensor research and develop- Liquid flyback booster pre-phase: A study ment: The cool film calorimeter – 170 assessment – 146 Austin, R. Eugene Reusable Launch Vehicle Technology Abu-Khajeel, Hasan Antol, Jeffrey Program – 127 Preliminary Thermal-Mechanical Sizing International Space Station Evolution of Metallic TPS: Process Development Data Book – 87 Austin, Robert Eugene and Sensitivity Studies – 33 Arbogast, Tim Single-stage-to-orbit: Meeting the challenge – 136 Achary, David C. 2nd Generation Reusable Launch Ve- Composite Development and Applica- hicle (2G RLV) – 82 Babai, Majid tions for RLV Tankage – 16 Ardema, M. D. Manufacturing Process Simulation of Large-Scale Cryotanks – 36 Adams, Michael J. Near-Optimal Operation of Dual-Fuel Liquid Engine Test Facilities Assess- Launch Vehicles – 119 Bahm, Cathy ment – 1 Near-Optimal Re-Entry Trajectories for Two Reconfigurable Flight-Control De- Adelfang, S. I. Reusable Launch Vehicles – 115 sign Methods: Robust Servomechanism A Compendium of Wind Statistics and and Control Allocation – 76 Ardema, Mark D. Models for the NASA Space Shuttle and Bahr, Alfred J. Other Aerospace Vehicle Pro- Advances in hypersonic vehicle synthe- Wireless Subsurface Microsensors for grams – 99 sis with application to studies of advanced thermal protection sys- Health Monitoring of Thermal Protection Aftosmis, Michael tem – 138 Systems on Hypersonic Vehicles – 69 Automated CFD Parameter Studies on Bailey, Michelle D. Distributed Parallel Computers – 36 Arnold, James Thermal Protection Materials for Reentry NASA Alternate Access to Station Ser- Aggarwal, Shiv Applications – 59 vice Concept – 65 System Identification of X-33 Neural Net- work – 8 Arnold, Jim Bailey, Michelle Active Wireless Temperature Sensors for Project Antares: A low cost modular Ahmad, Jasim Aerospace Thermal Protection Sys- launch vehicle for the future – 168 Automated CFD Parameter Studies on tems – 46 Distributed Parallel Computers – 36 Second Generation RLV Space Vehicle Fractographic Analysis of HfB2-SiC and Concept – 65 Alexander, Reginald ZrB2-SiC Composites – 57 A Collaborative Analysis Tool for Thermal Baize, Daniel G. Protection Systems for Single Stage to Health Monitoring Technology for Ther- The 1995 NASA High-Speed Research Orbit Launch Vehicles – 56 mal Protection Systems on Reusable Hy- Program Sonic Boom Workshop – 126 personic Vehicles – 71 Alexander, Stephen W. Bajwa, Anupa Mission and sizing analysis for the Beta 2 The Evolution of Flexible Insulation as NASA IVHM Technology Experiment for two-stage-to-orbit vehicle – 168 Thermal Protection Systems for Reus- X-vehicles (NITEX) – 67 able Launch Vehicles: AFRSI (Advanced Alexandrov, N. M. Flexible Reusable Surface Insulation) to Baker, Myles Multidisciplinary Approach to Aerospike CRI (Conformal Reusable Insula- Ion engine propelled Earth-Mars cycler Nozzle Design – 120 tion) – 67 with nuclear thermal propelled transfer vehicle, volume 2 – 144 Alexandrov, Natalia Thermal Protection Materials for Reentry Opportunities for Breakthroughs in and Planetary Applications – 68 Balakumar, P. Large-Scale Computational Simulation Wireless Subsurface Microsensors for N-Factor Computations for the X-33 Ve- and Design – 54 Health Monitoring of Thermal Protection hicle – 76 Alter, Stephen J. Systems on Hypersonic Vehicles – 69 Banerjee, Prasanta K. Opportunities for Breakthroughs in Wireless Subsurface Sensors for Health Development of an integrated BEM ap- Large-Scale Computational Simulation Monitoring of Thermal Protection Sys- proach for hot fluid structure interac- and Design – 54 tems on Hypersonic Vehicles – 68 tion – 169 X-33 Computational Arzaz, El Mehdi Bangham, M. E. Aeroheating/Aerodynamic Predictions Tether Transportation System and Comparisons With Experimental Project Antares: A low cost modular 168 Study – 109 Data – 25 launch vehicle for the future – Ambur, Damodar R. Ashkenas, Irving L. Bardina, Jorge E. Development of a Pressure Box to Evalu- Assessment of flying-quality criteria for Turbulence modeling for hypersonic ate Reusable-Launch-Vehicle air-breathing aerospacecraft – 167 flight – 161 Cryogenic-Tank Panels – 132 Atkins, Harold L. Barrett, Charles A. Anderson, Hobie Opportunities for Breakthroughs in Oxidation Behavior of Copper Alloy Can- Project Antares: A low cost modular Large-Scale Computational Simulation didates for Rocket Engine Applications launch vehicle for the future – 168 and Design – 54 (Technical Poster) – 48

B-1 Bartolotta, Paul Infrared Sensing Aeroheating Flight Ex- Blosser, M. L. NASA’s Advanced Space Transportation periement: STS-96 Flight Results – 84 Advanced Metallic Thermal Protection Program: RTA Project Summary – 42 Infrared Sensing Aeroheating Flight Ex- System Development – 64 Batts, Glen W. periment: STS-96 Flight Results – 86 Blosser, Max L. Natural Atmospheric Environment Model X-33 Computational Development of Metallic Thermal Protec- Development for the National Aeronau- Aeroheating/Aerodynamic Predictions tion Systems for the Reusable Launch tics and Space Administration’s Second and Comparisons With Experimental Vehicle – 124 Generation Reusable Launch Ve- Data – 25 Hypervelocity Impact Test Results for a hicle – 57 X-33 (Rev-F) Aeroheating Results of Test Metallic Thermal Protection Sys- Baughcum, Steven L. 6770 in NASA Langley 20-Inch Mach 6 tem – 10 Aircraft Emission Scenarios Projected in Air Tunnel – 96 Investigation of Fundamental Modeling Year 2015 for the NASA Technology Con- X-33 Rev-F Turbulent Aeroheating Re- and Thermal Performance Issues for a cept Aircraft (TCA) High Speed Civil sults From Test 6817 in NASA Langley Metallic Thermal Protection System De- Transport – 108 20-Inch Mach 6 Air Tunnel and Compari- sign – 33 Baumgarten, William J. sons With Computations – 26 Parametric Weight Comparison of Ad- Mission and sizing analysis for the Beta 2 X-33 Turbulent Aeroheating Measure- vanced Metallic, Ceramic Tile, and Ce- two-stage-to-orbit vehicle – 168 ments and Predictions – 32 ramic Blanket Thermal Protection Sys- tems – 88 Beason, Phil Bevacqua, Tim Bodan-Sanders, Patricia A Cryogenic Propellant Production Depot Integrated G and C Implementation for Low Earth Orbit – 81 within IDOS: A Simulink Based Reusable Reusable Cryogenic Tank VHM Using Launch Vehicle Simulation – 11 Fiber Optic Distributed Sensing Technol- Beaurain, Andre ogy – 108 Earth-to-Orbit Rocket Propulsion – 17 Bey, Kim S. Bogue, Rodney K. Beckham, Joanne Opportunities for Breakthroughs in Perspective on the National Aero-Space Lockheed Martin Response to the OSP Large-Scale Computational Simulation Plane Program instrumentation develop- Challenge – 20 and Design – 54 ment – 156 Beeghly, Jeff Bhutta, Bilal A. Boman, B. L. Project Antares: A low cost modular PNS predictions of external/internal hy- Heat pipes for wing leading edges of launch vehicle for the future – 168 personic flows for NASP propulsion ap- hypersonic vehicles – 178 plications – 157 Beil, Robert J. Book, Michael L. Propellant Feed System Leak Detection: Bibb, Karen L. An Advanced Video Sensor for Auto- Lessons Learned From the Linear Aero- Opportunities for Breakthroughs in mated Docking – 61 spike SR-71 Experiment (LASRE) – 93 Large-Scale Computational Simulation Borowski, Stanley K. and Design – 54 Bekey, Ivan Human lunar mission capabilities using General Public Space Travel and Tour- Biedron, R. T. SSTO, ISRU and LOX-augmented NTR ism – 96 Compute as Fast as the Engineers Can technologies: A preliminary assess- Bell, J. Think! ULTRAFAST COMPUTING TEAM ment – 137 Liquid flyback booster pre-phase: A study FINAL REPORT – 90 The rationale/benefits of nuclear thermal assessment – 146 Biedron, Robert T. rocket propulsion for NASA’s lunar space transportation system – 141 Bellini, Peter X. Opportunities for Breakthroughs in Highly Reusable Space Transportation Large-Scale Computational Simulation Borrer, Jerry System Concept Evaluation (The Argus and Design – 54 Flush Airdata Sensing (FADS) System Launch Vehicle) – 103 Calibration Procedures and Results for Biegel, Bryan Blunt Forebodies – 92 Belvin, H. L. NASA Exhibits – 59 Large Composite Structures Processing Bossi, J. A. Bienhoff, Dallas Technologies for Reusable Launch Ve- Personnel launch system autoland de- hicles – 75 Boeing 2nd Generation Reusable velopment study – 175 Launch Vehicle Architecture – 21 Berning, M. Boswell, Barry E. Liquid flyback booster pre-phase: A study Binder, Michael P. NASA’s New Orbital Space Plane: A assessment – 146 Real-Time Sensor Validation, Signal Re- Bridge to the Future – 19 construction, and Feature Detection for Bouquet, Clement Berry, Scott A. an RLV Propulsion Testbed – 98 Comparison of Methods for Determining Demonstration of Advanced C/SiC Boundary Layer Edge Conditions for Bird,R.K. Cooled Ramp – 56 Transition Correlations – 23 Advanced Metallic Thermal Protection Bouslog, S. System Development – 64 Computational/Experimental Aeroheat- Thermal Management Design for the ing Predictions for X-33 Phase 2 Ve- Blanchard, Robert C. X-33 Lifting Body – 104 hicle – 110 Infrared Sensing Aeroheating Flight Ex- Bouvier, Carl periement: STS-96 Flight Results – 84 Computational/Experimental Aeroheat- Reusable Cryogenic Tank VHM Using ing Predictions for X-33 – 112 Infrared Sensing Aeroheating Flight Ex- Fiber Optic Distributed Sensing Technol- Discrete Roughness Effects on Shuttle periment: STS-96 Flight Results – 86 ogy – 108 Orbiter at Mach 6 – 37 Bland, J. Bowditch, David N. Hypersonic Boundary-Layer Transition The 2002 NASA Faculty Fellowship Pro- The future challenge for aeropropul- for X-33 Phase 2 Vehicle – 111 gram Research Reports – 9 sion – 167

B-2 Bowles, J. V. Brownston, Lee Calvignac, Jacky Near-Optimal Operation of Dual-Fuel The NASA Integrated Vehicle Health Design and Testing of Non-Toxic RCS Launch Vehicles – 119 Management Technology Experiment for Thrusters for Second Generation Reus- 53 14 Near-Optimal Re-Entry Trajectories for X-37 – able Launch Vehicle – Reusable Launch Vehicles – 115 Bryan, Thomas C. Campbell, Jonathan Bowles, Jeffrey V. An Advanced Video Sensor for Auto- Rocket-Induced Magnetohydrodynamic A Two-Stage-to-Orbit Spaceplane Con- mated Docking – 61 Ejector: A Single-Stage-to-Orbit Ad- 80 vanced Propulsion Concept – 131 cept With Growth Potential – Bryant, L. Canabal, Francisco Bowman, L. M. Liquid flyback booster pre-phase: A study A study of facilities and fixtures for testing assessment – 146 Using CFD as Rocket Injector Design of a high speed civil transport wing com- Tool: Recent Progress at Marshall Space ponent – 127 Bufkin, A. Flight Center – 3 Liquid flyback booster pre-phase: A study Cano, R. J. Boxx, Dayna L. assessment – 146 Propellant Densification for Shuttle: The Large Composite Structures Processing SSME Perspective – 32 Bullock, Ellen P. Technologies for Reusable Launch Ve- Supersonic aeroelastic instability results hicles – 75 Bozung, Timothy J. for a NASP-like wing model – 159 Infrared Sensing Aeroheating Flight Ex- Caram, J. periement: STS-96 Flight Results – 84 Burchett, Bradley T. Liquid flyback booster pre-phase: A study assessment – 146 Infrared Sensing Aeroheating Flight Ex- Genetic Algorithm Tuned Fuzzy Logic for periment: STS-96 Flight Results – 86 Gliding Return Trajectories – 12 Card, Michael F. Effects of Stiffening and Mechanical Burchett, Bradley Bradford, John E. Load on Thermal Buckling of Stiffened System Level Uncertainty Assessment Fuzzy Logic Trajectory Design and Guid- Cylindrical Shells – 121 for Collaborative RLV Design – 46 ance for Terminal Area Energy Manage- ment – 8 Structures and materials technology for Bradley, Michael hypersonic aerospacecraft – 178 Space Shuttle Main Engine Off-Nominal Burkardt, Leo A. Carpenter, Mark H. Low Power Level Operation – 117 The design and evolution of the beta Opportunities for Breakthroughs in Space Shuttle Main Engine Start with two-stage-to-orbit horizontal takeoff and landing launch system – 159 Large-Scale Computational Simulation Off-Nominal Propellant Inlet Pres- and Design – 54 sures – 116 Burken, John J. Thrust Enhancement in Hypervelocity Brandon, Larry Full Envelope Reconfigurable Control Nozzles by Chemical Catalysis – 102 Predictor-Corrector Entry Guidance for Design for the X-33 Vehicle – 72 Carrington, Connie Reusable Launch Vehicles – 74 Reconfigurable Control Design for the A Cryogenic Propellant Production Depot Full X-33 Flight Envelope – 72 Braun, James for Low Earth Orbit – 81 Durability of Membrane Electrode As- Reconfigurable Flight Control Designs semblies (MEAs) in PEM Fuel Cells Op- With Application to the X-33 Ve- Carroll, Carol erated on Pure Hydrogen and Oxy- hicle – 95 Evaluation of Thermal Control Coatings gen – 6 for Flexible Ceramic Thermal Protection Two Reconfigurable Flight-Control De- Systems – 117 Braun, R. D. sign Methods: Robust Servomechanism Comparison of Two Multidisciplinary Op- and Control Allocation – 76 Materials Testing on the DC-X and DC- timization Strategies for Launch-Vehicle XA – 121 Butas, John P. Design – 123 Cartwright, Curt Rocket Engine Health Management: Project Antares: A low cost modular Optimal technology investment strate- Early Definition of Critical Flight Mea- launch vehicle for the future – 168 gies for a reusable launch vehicle – 134 surements – 19 Breen, Daniel P. Cazier, F. W., Jr. Butrica, Andrew J. Advanced Control Surface Seal Develop- Structural dynamic and aeroelastic con- X-33 Development History – 113 ment at NASA GRC for 3rd Generation siderations for hypersonic ve- RLV – 29 Cady, Edwin C. hicles – 173 Development and Capabilities of Unique Slush hydrogen technology pro- Cerro, J. A. gram – 149 Structural Seal Test Rigs – 41 A study of facilities and fixtures for testing Update on the Development and Capa- Cain, L. of a high speed civil transport wing com- bilities of Unique Structural Seal Test Liquid flyback booster pre-phase: A study ponent – 127 Rigs – 3 assessment – 146 Cerro, Jeffrey A. Breisacher, Kevin J. Calise, A. J. A Procedure for Structural Weight Esti- Experimental Evaluation of a Subscale Rapid near-optimal aerospace plane tra- mation of Single Stage to Orbit Launch Gaseous Hydrogen/Gaseous Oxygen jectory generation and guidance – 171 Vehicles (Interim User’s Manual) – 44 Coaxial Rocket Injector – 34 Rapid near-optimal trajectory generation Chamitoff, Gregory Errol Brinker, Stan and guidance law development for Robust intelligent flight control for hyper- Reusable Launch Vehicle Re- single-stage-to-orbit airbreathing ve- sonic vehicles – 166 search – 119 hicles – 177 Champion, R. Brown, Richard W. Calomino, Anthony M. Launch Vehicle Sizing Benefits Utilizing Design of launch systems using continu- Ceramic Matrix Composites (CMC) Life Main Propulsion System Crossfeed and ous improvement process – 134 Prediction Method Development – 88 Project Status – 34

B-3 Champion, Robert H. Chen, Y. S. Claxton, Jeffrey S. On-Orbit Propulsion System Project Unic Unstructured CFD Methodology De- Advanced Range Safety System for High Overview – 44 velopment – 82 Energy Vehicles – 15 Champion, Robert Cheng, Guangfeng Clinton, R. C., Jr. Design and Testing of Non-Toxic RCS Panel Flutter and Sonic Fatigue Analysis Fabrication and Testing of Ceramic Ma- Thrusters for Second Generation Reus- for RLV – 77 trix Composite Rocket Propulsion Com- able Launch Vehicle – 14 Cheng, Peter Y. ponents – 75 Chan, Samuel Y. Advanced aeroservoelastic stabilization Large Composite Structures Processing Advanced aeroservoelastic stabilization techniques for hypersonic flight ve- Technologies for Reusable Launch Ve- techniques for hypersonic flight vehicles – 163 hicles – 75 hicles – 163 Chenoweth, James Clinton, Raymond G., Jr. Chan, William Using CFD as Rocket Injector Design National Rocket Propulsion Materials Tool: Recent Progress at Marshall Space Plan: A NASA, Department of Defense, Progress in Unsteady Turbopump Flow Flight Center – 3 and Industry Partnership – 73 Simulations Using Overset Grid Sys- tems – 30 Childers, Dave Cobleigh, Brent R. Progress in Unsteady Turbopump Flow Test Results for an Off-the-Shelf Design and Calibration of the X-33 Flush Simulations – 51 GPS/INS During Approach and Landing Airdata Sensing (FADS) System – 113 Testing of the X-40A – 63 Time Dependent Simulation of Tur- Development of the X-33 Aerodynamic bopump Flows – 55 Chiles, Harry R. Uncertainty Model – 108 Evaluation of the Linear Aerospike SR-71 Flight Stability and Control and Perfor- Time-Dependent Simulations of Tur- Experiment (LASRE) Oxygen Sen- bopump Flows – 38 mance Results from the Linear Aerospike sor – 91 SR-71 Experiment (LASRE) – 101 Chandler, Frank Chojnacki, Kent T. Flush Airdata Sensing (FADS) System Launch Vehicle Sizing Benefits Utilizing Rocket Based Combined Cycle (RBCC) Calibration Procedures and Results for Main Propulsion System Crossfeed and Propulsion Technology Workshop. Vol- Blunt Forebodies – 92 Project Status – 34 ume 1: Executive summary – 166 Cockrell, B. Chapman, Jim Rocket Based Combined Cycle (RBCC) Liquid flyback booster pre-phase: A study Magnetohydrodynamic Augmented Pro- Propulsion Workshop, volume 2 – 166 assessment – 146 pulsion Experiment – 52 Chou, H. C. Cockrell, Charles E., Jr. Chappell, C. R. Near-Optimal Operation of Dual-Fuel Technology Roadmap for Dual-Mode Launch Vehicles – 119 Research Reports: 1996 NASA/ASEE Scramjet Propulsion to Support Space- Summer Faculty Fellowship Pro- Chou, H.-C. Access Vision Vehicle Develop- gram – 102 Near-Optimal Re-Entry Trajectories for ment – 38 Reusable Launch Vehicles – 115 Charania, A. C. Cole, John System Level Uncertainty Assessment Chowdhry, Rajiv S. Magnetohydrodynamic Augmented Pro- for Collaborative RLV Design – 46 Hypersonic vehicle control law develop- pulsion Experiment – 52 ment using H infinity and mu- Rocket-Induced Magnetohydrodynamic Chau, William synthesis – 164 Ejector: A Single-Stage-to-Orbit Ad- Project Antares: A low cost modular Hypersonic vehicle model and control vanced Propulsion Concept – 131 launch vehicle for the future – 168 law development using H(infinity) and Cole, Kevin D. 144 Chaudhary, Ashwani micron synthesis – Design of Experiments for the Thermal Dynamics and Stability and Control Chowdhry, Rajiv Characterization of Metallic Foam – 1 Characteristics of the X-37 – 62 Predictor-Corrector Entry Guidance for Cole, Stanley R. Reusable Launch Vehicles – 74 Cheatwood, F. McNeil Supersonic aeroelastic instability results Opportunities for Breakthroughs in Christensen, E. R. for a NASP-like wing model – 159 Large-Scale Computational Simulation Integrated Technology Assessment Cen- Collier, Craig and Design – 54 ter (ITAC) Update – 42 Design and Analysis of a Two-Stage-to- Chen, L. - Y. Christenson, R. L. Orbit Airbreathing Hypersonic Vehicle A Hazardous Gas Detection System for Reusable Rocket Engine Operability Concept – 129 Modeling and Analysis – 101 Aerospace and Commercial Applica- Conners, Timothy R. 98 tions – Christenson, Rick L. Flight Stability and Control and Perfor- Chen, L.-Y. Rocket Engine Health Management: mance Results from the Linear Aerospike A Hazardous Gas Detection System for Early Definition of Critical Flight Mea- SR-71 Experiment (LASRE) – 101 surements – 19 Aerospace and Commercial Applica- Cook, Stephen A. tions – 98 Cikanek, Harry A., III The Reusable Launch Vehicle Technol- Chen, R. R. Earth-to-Orbit Rocket Propulsion – 17 ogy Program and the X-33 Advanced Technology Demonstrator – 128 Advanced Metallic Thermal Protection Citrin, K. M. System Development – 64 Heat pipes for wing leading edges of Cook, Stephen hypersonic vehicles – 178 Chen, Victor L. An Approach to Establishing System Benefits for Technology in NASA’s Hy- X-37 C-SiC CMC Control Surface Com- Clancy, Daniel personics Investment Area – 67 ponents Development: Status of the The NASA Integrated Vehicle Health NASA/Boeing/USAF Orbital Vehicle and Management Technology Experiment for Magnetohydrodynamic Augmented Pro- Related Efforts – 20 X-37 – 53 pulsion Experiment – 52

B-4 Corban, J. E. Crumbly, Chris Dasappa, Srini Rapid near-optimal aerospace plane tra- NASA Alternate Access to Station Ser- An Overview of In-Stu Treatability Stud- jectory generation and guidance – 171 vice Concept – 65 ies at Marshall Space Flight Center, 78 Rapid near-optimal trajectory generation Cruse, T. A. Huntsville, Alabama – and guidance law development for Dynamics and Control of Orbiting Space Datta, A. single-stage-to-orbit airbreathing ve- Structures NASA Advanced Design Pro- Lee waves: Benign and malig- hicles – 177 gram (ADP) – 128 nant – 157 Corda, Stephen Cruse, Thomas Davis, Daniel J. Flight Testing the Linear Aerospike The 1994 NASA/USRA/ADP Design NASA’s Space Launch Initiative Targets SR-71 Experiment (LASRE) – 100 Projects – 140 Toxic Propellants – 64 Cornelison, C. J. Curran, F. M. Davis, Hubert P. Results of Two-Stage Light-Gas Gun De- Integrated Technology Assessment Cen- The StarBooster System: A Cargo Air- velopment Efforts and Hypervelocity Im- ter (ITAC) Update – 42 craft for Space – 92 pact Tests of Advanced Thermal Protec- Curry, D. tion Materials – 100 Davis, Stephan R. Liquid flyback booster pre-phase: A study NASA’s New Orbital Space Plane: A Corpening, Griffin P. assessment – 146 Bridge to the Future – 19 Evaluation of the Linear Aerospike SR-71 Curry, Robert E. Experiment (LASRE) Oxygen Sen- Dawdy, Andrew sor – 91 Wind-Tunnel Investigations of Blunt- Project Antares: A low cost modular Body Drag Reduction Using Forebody launch vehicle for the future – 168 Flight Testing the Linear Aerospike Surface Roughness – 85 SR-71 Experiment (LASRE) – 100 Dawson, Don Dagnostino, Mark Linear Aerospike SR-71 Experiment Infrared Sensing Aeroheating Flight Ex- Analysis of X-33 Linear Aerospike Plume periement: STS-96 Flight Results – 84 (LASRE): Aerospace Propulsion Hazard Induced Base-Heating Environment Dur- Mitigation Systems – 102 ing Power-Pack Out – 29 Infrared Sensing Aeroheating Flight Ex- periment: STS-96 Flight Results – 86 Cotting, M. Christopher DAgostino, Mark G. Full Envelope Reconfigurable Control X-33 XRS-2200 Linear Aerospike Engine Dean, P. Design for the X-33 Vehicle – 72 Sea Level Plume Radiation – 66 Heat flux sensor research and development: The cool film calorimeter – 170 Reconfigurable Control Design for the D’Agostino, Mark Full X-33 Flight Envelope – 72 Base-Bleed Effect on X-33 Aerospike Deardorff, Glenn Coughlin, D. Plume Induced Base-Heating Environ- NASA Exhibits – 59 A guidance and control assessment of ment During Power-Pack Out – 25 DeBonis, J. R. three vertical landing options for DAgostino, Mark Inlet Development for a Rocket Based 136 RLV – X-33 XRS-2200 Linear Aerospike Engine Combined Cycle, Single Stage to Orbit Cowart, Kris Sea Level Plume Radiation – 79 Vehicle Using Computational Fluid Dy- A Method of Integrating Aeroheating into D’Agostino, M. namics – 95 Conceptual Reusable Launch Vehicle Unic Unstructured CFD Methodology De- Dellacorte, Christopher Design: Evaluation of Advanced Thermal velopment – 82 High temperature NASP engine seals: A Protection Techniques for Future Reus- technology review – 173 able Launch Vehicles – 77 Dalbello, Teryn Evaluation of Advanced Thermal Protec- Computations of Internal and External DeMange, Jeff tion Techniques for Future Reusable Axisymmetric Nozzle Aerodynamics at Structural Seal Development – 11 Transonic Speeds – 4 Launch Vehicles – 62 DeMange, Jeffrey J. Dang, Lisa Cox, Michael Advanced Control Surface Seal Develop- Design and Testing of Non-Toxic RCS Development of Processing Techniques ment at NASA GRC for 3rd Generation Thrusters for Second Generation Reus- for Advanced Thermal Protection Materi- RLV – 29 able Launch Vehicle – 14 als – 112 Development and Capabilities of Unique Cox, Timothy H. Daniel, Charles Structural Seal Test Rigs – 41 Flight Stability and Control and Perfor- A Quantitative Reliability, Maintainability and Supportability Approach for NASA’s Third Generation RLV Structural Seal mance Results from the Linear Aerospike Development Programs at NASA SR-71 Experiment (LASRE) – 101 Second Generation Reusable Launch Vehicle – 41 GRC – 3 Flight Testing the Linear Aerospike Second Generation RLV Space Vehicle Update on the Development and Capa- SR-71 Experiment (LASRE) – 100 Concept – 65 bilities of Unique Structural Seal Test Creech, Stephen D. Rigs – 3 Daniels, Dan 2nd Generation Reusable Launch Ve- Reusable Launch Vehicle Attitude Con- Denis, Vincent hicle Potential Commercial Development trol Using a Time-Varying Sliding Mode High Altitude Launch for a Practical Scenarios – 79 Control Technique – 40 SSTO – 29 Creech, Steve Dargush, Gary F. Dennis, J. Orbital Space Plane Cost Credibility – 4 Development of an integrated BEM ap- Fabrication and Testing of Ceramic Ma- Crisalli, Dave proach for hot fluid structure interac- trix Composite Rocket Propulsion Com- Reusable Launch Vehicle Re- tion – 169 ponents – 75 search – 119 Daryabeigi, Kamran DePriest, Douglas Crittenden, Paul E. Effective Thermal Conductivity of High Statistical Methods for Rapid Aerother- Design of Experiments for the Thermal Temperature Insulations for Reusable mal Analysis and Design Technol- Characterization of Metallic Foam – 1 Launch Vehicles – 97 ogy – 55

B-5 Detert, Bruce Dukeman, Greg Ebeling, Charles E. Project Antares: A low cost modular An Automated Method to Compute Or- O/S analysis of conceptual space ve- launch vehicle for the future – 168 bital Re-entry Trajectories with Heating hicles – 134 Constraints – 39 Dewitt, R. L. Eckel, A. Background, current status, and progno- An Automated Method to Compute Or- Fabrication and Testing of Ceramic Ma- sis of the ongoing slush hydrogen tech- bital Re-Entry Trajectories with Heating trix Composite Rocket Propulsion Com- nology development program for the Constraints – 53 ponents – 75 NASP – 177 Dula, Arthur M. Eckel, Andrew Djomehri, M. Jahed The StarBooster System: A Cargo Air- Demonstration of Advanced C/SiC NASA Exhibits – 59 craft for Space – 92 Cooled Ramp – 56 Doehnert, Bill Dumbacher, D. L. Eckel, Andy 2nd Generation Reusable Launch Ve- 2nd Generation RLV: Program Goals and Ceramic Matrix Composite Cooled hicle (2G RLV) – 82 Acquisition Strategy – 79 Nozzle Material Development Pro- 21 Doggett, Robert V., Jr. Dumbacher, Dan L. gram – NASP aeroservothermoelasticity stud- NASA 2nd Generation RLV Program In- Effinger, M. R. troduction, Status and Future ies – 174 Fabrication and Testing of Ceramic Ma- Plans – 31 Structural dynamic and aeroelastic con- trix Composite Rocket Propulsion Com- siderations for hypersonic ve- NASA’s New Orbital Space Plane: A ponents – 75 Bridge to the Future – 19 hicles – 173 Effinger, Mike Domack, Marcia S. Dumbacher, Dan Lightweight Chambers for Thrust Assem- Fatigue Crack Growth Rate Test Results NASA’s Orbital Space Plane Risk Re- blies – 82 duction Strategy – 13 for Al-Li 2195 Parent Metal, Variable Po- Ehernberger, L. J. larity Plasma Arc Welds and Friction Stir Dumbacher, Daniel L. Stratospheric turbulence measurements 89 Welds – Orbital Space Plane Program Sta- and models for aerospace plane de- Dorney, Daniel J. tus – 5 sign – 164 Design and Analysis of Turbines for Progress and Status of the 2nd Genera- Elam,S.K. Space Applications – 16 tion Reusable Launch Vehicle Pro- Test Report for NASA MSFC Support of RLV Turbine Performance Optimiza- gram – 76 the Linear Aerospike SR-71 Experiment tion – 83 Space Launch Initiative: New Capabili- (LASRE) – 89 ties ... New Horizons – 36 Dorsey, J. T. Elam, Sandra K. Advanced Metallic Thermal Protection Dumbacher, Daniel Lightweight Chambers for Thrust Assem- System Development – 64 Space Launch Initiative: New Capabili- blies – 82 ties - New Horizons – 38 Douglas, Stan Elam, Sandy Analysis of X-33 Linear Aerospike Plume Duncan, Beverly S. Development of Aluminum Composites Induced Base-Heating Environment Dur- Computational analysis in support of the for a Rocket Engine’s Lightweight Thrust ing Power-Pack Out – 29 SSTO ﬂowpath test – 143 Cell – 80 Droege, Alan Dunlap, Patrick H., Jr. Elam, S. Analysis of X-33 Linear Aerospike Plume Advanced Control Surface Seal Develop- Development of Metal Matrix Compos- Induced Base-Heating Environment Dur- ment at NASA GRC for 3rd Generation ites for NASA’s Advanced Propulsion ing Power-Pack Out – 29 RLV – 29 Systems – 74 Base-Bleed Effect on X-33 Aerospike Development and Capabilities of Unique Fabrication and Testing of Ceramic Ma- Plume Induced Base-Heating Environ- Structural Seal Test Rigs – 41 trix Composite Rocket Propulsion Com- ment During Power-Pack Out – 25 Third Generation RLV Structural Seal ponents – 75 Stage Separation CFD Tool Develop- Development Programs at NASA Lightweight Chambers for Thrust Assem- ment and Evaluation – 38 GRC – 3 blies – 66 Update on the Development and Capa- Drummond, J. P. Ellerby, D. T. bilities of Unique Structural Seal Test Thrust Enhancement in Hypervelocity Rigs – 3 Fractographic Analysis of HfB2-SiC and Nozzles by Chemical Catalysis – 102 ZrB2-SiC Composites – 57 Dunn, H. J. Drummond, Philip J. Ellerby, Donald Multidisciplinary Approach to Aerospike Opportunities for Breakthroughs in Nozzle Design – 120 Thermal Protection Materials for Reentry Large-Scale Computational Simulation Applications – 59 and Design – 54 Multidisciplinary Optimization Branch Ex- perience Using iSIGHT Software – 91 Ellis, John R. Duffey, Jack Ceramic Matrix Composites (CMC) Life Transportation Systems Analyses (TSA): Dunn, P. W. Prediction Method Development – 88 Highly Reusable Space Transportation GTX Reference Vehicle Structural Veriﬁ- (HRST). A preliminary programmatic as- cation Methods and Weight Sum- Emanuel, G. sessment – 131 mary – 44 An integrated aerodynamic/propulsion study for generic aero-space planes Duffy, James B. Early, James based on waverider concepts – 173 Advanced transportation system study: Evaluation and Characterization Study of Manned launch vehicle concepts for two Dual Pulse Laser-Induced Spark (DPLIS) Emdee, Jeffery L. way transportation system payloads to for Rocket Engine Ignition System Appli- Liquid Engine Test Facilities Assess- LEO – 152 cation – 14 ment – 1

B-6 Engelund, W. C. Fisher, J. E. Frankhauser, Chris Hyper-X Wind Tunnel Program – 109 Autocommander: A Supervisory Control- Configuration development study of the Subsonic-to-Hypersonic Aerodynamic ler for Integrated Guidance and Control OSU 1 hypersonic research ve- 154 Characteristics for a Winged, Circular- for the 2nd Generation Reusable Launch hicle – Vehicle – 40 Body, Single-Stage-to-Orbit Spacecraft Franklin, W. Configuration – 124 Fisher, J. Large Composite Structures Processing Engelund, Walter C. Direct Fault Tolerant RLV Altitude Con- Technologies for Reusable Launch Ve- Aerothermoelastic analysis of a NASP trol: A Singular Perturbation Ap- hicles – 75 proach – 40 demonstrator model – 155 Frassanito, John Engler, Leah M. Fisher, Joseph E. The StarBooster System: A Cargo Air- NASA’s New Orbital Space Plane: A Integrated G and C Implementation craft for Space – 92 Bridge to the Future – 19 within IDOS: A Simulink Based Reusable Launch Vehicle Simulation – 11 Freeman, Delma C., Jr. Ennix, Kimberly A. Reusable Launch Vehicle Technology Fisher, Mark F. Evaluation of the Linear Aerospike SR-71 Program – 127 NASA’s New Orbital Space Plane: A Experiment (LASRE) Oxygen Sen- Bridge to the Future – 19 Single-stage-to-orbit: Meeting the chal- sor – 91 lenge – 136 Fisher, Mark Linear Aerospike SR-71 Experiment Freeman, Ken (LASRE): Aerospace Propulsion Hazard Main Engine Prototype Development for NASA Exhibits – 59 Mitigation Systems – 102 2nd Generation RLV RS-83 – 43 Ensworth, Clint Flandro, G. A. Freeman, M. High Voltage Design Guidelines: A Rapid near-optimal trajectory generation Research Reports: 1996 NASA/ASEE and guidance law development for Summer Faculty Fellowship Pro- Timely Update – 61 single-stage-to-orbit airbreathing ve- gram – 102 Enyinda, Chris I. hicles – 177 Freestone, Todd M. Second Generation Reusable Launch Fleming, M. M. Vehicle Development and Global Com- Testing the Preliminary X-33 Navigation Lightweight durable TPS: Milestone petitiveness of US Space Transportation System – 84 4–151 Industry: Critical Success Factors As- French, James R. sessment – 51 Fletcher, David Pathfinder Technologies Specialist, Erbland, Peter Advanced spacecraft designs in support X-37 – 69 of human missions to earth’s neighbor- Perspective on the National Aero-Space hood – 28 French, Jim Plane Program instrumentation develop- Reusable Launch Vehicle Re- 156 Florance, James R. ment – search – 119 Supersonic aeroelastic instability results Ervin, Miles for a NASP-like wing model – 159 Friedman, Mark Project Antares: A low cost modular An Improved RLV Stability Analysis Via a launch vehicle for the future – 168 Fogle, Frank R. Advanced Guidance and Control Meth- Continuation Approach – 9 Escher, D. ods for Reusable Launch Vehicles: Test Fulton, Christopher E. Integrated Technology Assessment Cen- Results – 30 ter (ITAC) Update – 42 Real-Time Sensor Validation, Signal Re- An Automated Method to Compute Or- construction, and Feature Detection for Espenschied, Erich bital Re-entry Trajectories with Heating an RLV Propulsion Testbed – 98 The Control System for the X-33 Linear Constraints – 39 Aerospike Engine – 105 Gallaher, M. An Automated Method to Compute Or- A guidance and control assessment of Falangas, Eric bital Re-Entry Trajectories with Heating three vertical landing options for Dynamics and Stability and Control Constraints – 53 RLV – 136 Characteristics of the X-37 – 62 Fogle, Frank Gambrel, Dave Farhangi, Shahram A Plan for Advanced Guidance and Con- NASA Exhibits – 59 Experimental Evaluation of a Subscale trol Technology for 2nd Generation Re- Gaseous Hydrogen/Gaseous Oxygen usable Launch Vehicles – 52 Garbos, Raymond J. 34 Coaxial Rocket Injector – Follett, W. W. X-33/RLV System Health Management/ 107 Operationally efficient propulsion system Multidisciplinary Approach to Aerospike Vehicle Health Management – study (OEPSS) data book – 110 Nozzle Design – 120 Garbos, Raymond Farr, John L., Jr. Ford, Robert Flight Demonstration of X-33 Vehicle A Thermal Management Systems Model Orbital Space Plane (OSP) Program at Health Management System Compo- for the NASA GTX RBCC Concept – 53 Lockheed Martin – 7 nents on the F/A-18 Systems Research Aircraft – 15 Fikes, John Forward, Robert L. A Cryogenic Propellant Production Depot Estimate of avoidance maneuver rate for Garcia, Robert for Low Earth Orbit – 81 HASTOL tether boost facility – 28 Using CFD as Rocket Injector Design Tool: Recent Progress at Marshall Space Finckenor, J. L. Foster, Lee D. Flight Center – 3 Composite,Cryogenic, Conformal, Com- Testing the Preliminary X-33 Navigation mon Bulkhead, Aerogel-Insulated Tank System – 84 Garcia, Roberto (CBAT) – 74 Fox, Charles H., Jr. Overview of Fluid Dynamic Activities at the Marshall Space Flight Center – 56 Findley, Benjamin Subsonic aerodynamic characteristics of Permeability of Impacted Coated Com- a proposed advanced manned launch Overview of MSFC’s Applied Fluid Dy- posite Laminates – 49 system orbiter configuration – 160 namics Analysis Group Activities – 24

B-7 Garn, Michelle Goracke, B. David Trade study plan for Reusable Hydrogen Multi-Disciplinary Analysis for Future Advanced Low-Cost O2/H2 Engines for Composite Tank System Launch Systems Using NASA’s Ad- the SSTO Application – 123 (RHCTS) – 145 vanced Engineering Environment Greenberg, Joel S. (AEE) – 17 Graham, J. Bart 2nd Generation RLV: Program Goals and STARS: The Space Transportation Archi- Garner, E. C. Acquisition Strategy – 79 tecture Risk System – 116 Heat pipes for wing leading edges of hypersonic vehicles – 178 Graham, Matthew Greene, William D. System Level Uncertainty Assessment Propellant Densification for Shuttle: The Gates, Thomas S. for Collaborative RLV Design – 46 SSME Perspective – 32 High Temperature Polyimide Materials in Extreme Temperature Environ- Grantham, Katie Testing of Densified Liquid Hydrogen ments – 73 Adaptive Critic Neural Network-Based Stratification in a Scale Model Propellant 67 Terminal Area Energy Management and Tank – Gaunce, M. T. Approach and Landing Guidance – 9 Human Transportation System (HTS) Gregory, Irene M. study, volume 2 – 155 Gray, Hugh R. Hypersonic vehicle control law development using H infinity and mu- Characterization of C/SiC Ceramic Ma- Gelderloos, Hendrik synthesis – 164 trix Composites (CMCs) with Novel Inter- Test Results for an Off-the-Shelf face Fiber Coatings – 45 Hypersonic vehicle model and control GPS/INS During Approach and Landing law development using H(infinity) and Testing of the X-40A – 63 Oxidation Behavior of Copper Alloy Can- didates for Rocket Engine Applications micron synthesis – 144 Genge, G. (Technical Poster) – 48 Trim drag reduction concepts for horizon- Fabrication and Testing of Ceramic Ma- tal takeoff single-stage-to-Orbit ve- trix Composite Rocket Propulsion Com- Green, Bryan hicles – 170 ponents – 75 NASA Exhibits – 59 Griffin, Brian George, Daweel J. Greenbauer-Seng, L. Manufacturing Process Simulation of Liquid Engine Test Facilities Assess- Oxidation-Reduction Resistance of Ad- Large-Scale Cryotanks – 36 ment – 1 vanced Copper Alloys – 11 Griffin, Lisa W. Georgiadis, Nicholas Greenbauer-Seng, Leslie Computations of Internal and External Design and Analysis of Turbines for Oxidation Behavior of Copper Alloy Can- Space Applications – 16 Axisymmetric Nozzle Aerodynamics at didates for Rocket Engine Applications Transonic Speeds – 4 (Technical Poster) – 48 Experimental Performance Evaluation of Gerry, Mark a Supersonic Turbine for Rocket Engine Greenberg, H. S. Applications – 10 A Cryogenic Propellant Production Depot Lightweight durable TPS: Milestone RLV Turbine Performance Optimiza- for Low Earth Orbit – 81 4–151 tion – 83 Geyer, M. S. Milestone 4: Test plan for Reusable Hy- Human Transportation System (HTS) drogen Composite Tank System Griffin, Lisa study, volume 2 – 155 (RHCTS). Task 3: Composite tank mate- Overview of Fluid Dynamic Activities at the Marshall Space Flight Center – 56 Gibson, Walter M. rials – 142 Advanced High Brilliance X-Ray Milestone 4: Thrust structure concepts Overview of MSFC’s Applied Fluid Dy- Source – 111 and IHM screening Graphite Composite namics Analysis Group Activities – 24 Primary Structure (GCPS) – 143 Giesing, Joe Griffin, Timothy P. Evaluation of Computational Aeroelastic- Requirements report for SSTO vertical Propellant Feed System Leak Detection: ity Codes for Loads and Flutter – 82 take-off and horizontal landing ve- Lessons Learned From the Linear Aero- hicle – 143 spike SR-71 Experiment (LASRE) – 93 Glasgow, J. K. Requirements report for SSTO vertical Grimsley, Brian W. An Overview of In-Stu Treatability Stud- take-off/horizontal landing ve- ies at Marshall Space Flight Center, hicle – 149 Cryopumping in Cryogenic Insulations Huntsville, Alabama – 78 for a Reusable Launch Vehicle – 2 Selection process for trade study: Graph- Glass, Christopher E. ite Composite Primary Structure Guitierrez, Sonia Infrared Sensing Aeroheating Flight Ex- (GCPS) – 148 A Cryogenic Propellant Production Depot periement: STS-96 Flight Results – 84 Selection process for trade study: Reus- for Low Earth Orbit – 81 Infrared Sensing Aeroheating Flight Ex- able Hydrogen Composite Tank System Guptill, James D. periment: STS-96 Flight Results – 86 (RHCTS) – 148 Neural Network and Regression Approxi- Glass, David E. Test Plan. GCPS Task 4, subtask 4.2 mations in High Speed Civil Transport Bonding and Sealing Evaluations for thrust structure development – 147 Aircraft Design Optimization – 106 Cryogenic Tanks – 116 Test plan. GCPS task 7, subtask 7.1: IHM Gusman, Mike development – 149 Gnoffo, Peter A. Thermal Protection Materials for Reentry Opportunities for Breakthroughs in Test plan. Task 5, subtask 5.2: Early Applications – 59 Large-Scale Computational Simulation on-orbit TPSdebris impact tests – 148 Guy, R. Wayne and Design – 54 Test Plans. Lightweight Durable TPS: Technology Roadmap for Dual-Mode Tasks 1,2,4,5, and 6 – 146 Gomez, Reynaldo Scramjet Propulsion to Support Space- Stage Separation CFD Tool Develop- Trade study plan for Graphite Composite Access Vision Vehicle Develop- ment and Evaluation – 38 Primary Structure (GCPS) – 146 ment – 38

B-8 Guzinski, Mike Hanson, John Henderson, Stephen C. Evaluation of Thermal Control Coatings An Automated Method to Compute Or- Aircraft Emission Scenarios Projected in for Flexible Ceramic Thermal Protection bital Re-entry Trajectories with Heating Year 2015 for the NASA Technology Con- Systems – 117 Constraints – 39 cept Aircraft (TCA) High Speed Civil Transport – 108 Haberbusch, Mark S. An Automated Method to Compute Or- A summary of the slush hydrogen tech- bital Re-Entry Trajectories with Heating Hendricks, Robert C. nology program for the National Aero- Constraints – 53 2000 NASA Seal/Secondary Air System Space Plane – 140 Hardy, T. L. Workshop – 62 Hadid, A. H. Background, current status, and progno- Henley, Mark Multidisciplinary Approach to Aerospike sis of the ongoing slush hydrogen tech- A Cryogenic Propellant Production Depot Nozzle Design – 120 nology development program for the for Low Earth Orbit – 81 NASP – 177 Haering, Edward A., Jr. Henry, John Hardy, Terry L. Flush Airdata Sensing (FADS) System Reusable Launch Vehicle Re- Calibration Procedures and Results for A summary of the slush hydrogen tech- search – 119 Blunt Forebodies – 92 nology program for the National Aero- Space Plane – 140 Henze, Chris Haering, Edward A. NASA Exhibits – 59 Design and Calibration of the X-33 Flush Prediction of the ullage gas thermal Airdata Sensing (FADS) System – 113 stratification in a NASP vehicle propellant Herring, Helen tank experimental simulation using Characterization of Thin Film Polymers Hafley, Robert A. FLOW-3D – 177 Through Dynamic Mechanical Analysis Fatigue Crack Growth Rate Test Results Slush hydrogen propellant production, and Permeation – 22 for Al-Li 2195 Parent Metal, Variable Po- transfer, and expulsion studies at the larity Plasma Arc Welds and Friction Stir NASA K-Site Facility – 173 Hillard, G. Barry Welds – 89 High Voltage Design Guidelines: A Slush hydrogen transfer studies at the Timely Update – 61 Hagopian, Jeff NASA K-Site Test Facility – 167 Space Ops 2002: Bringing Space Opera- Hillman, Keithan Technology issues associated with using tions into the 21st Century. Track 3: Op- Characterization of Polyimide Foams for densified hydrogen for space ve- erations, Mission Planning and Control. Ultra-Lightweight Space Struc- hicles – 167 2nd Generation Reusable Launch tures – 22 Vehicle-Concepts for Flight Opera- Harris, Charles E. Hinke, Thomas tions – 35 Advanced Durability and Damage Toler- NASA Exhibits – 59 Haimes, B. ance Design and Analysis Methods for A Geometry-Centered Approach For Composite Structures: Lessons Learned Hinkle, Karrie Two-Stage-To-Orbit Vehicle Synthe- from NASA Technology Development Task 4 supporting technology. Part 1: sis – 45 Programs – 26 Detailed test plan for leading edge tile Harris, S. development. Leading edge material de- Halbig, Michael C. velopment and testing – 139 Ceramic Matrix Composites (CMC) Life A Geometry-Centered Approach For Prediction Method Development – 88 Two-Stage-To-Orbit Vehicle Synthe- Hoberecht, Mark sis – 45 Durability of Membrane Electrode As- Hamaker, Joseph H. Hartman, John semblies (MEAs) in PEM Fuel Cells Op- Estimating the Cost of NASA’s Space Automation of NDE on RSRM Metal erated on Pure Hydrogen and Oxy- Launch Initiative: How SLI Cost Stack Up gen – 6 Against the Shuttle – 51 Components – 57 Hass, Neal Performance Evaluation of Electro- Hamilton, H. Harris, II chem’s PEM Fuel Cell Power Plant for Linear Aerospike SR-71 Experiment Computational/Experimental Aeroheat- NASA’s 2nd Generation Reusable (LASRE): Aerospace Propulsion Hazard ing Predictions for X-33 – 112 Launch Vehicle – 6 Mitigation Systems – 102 Discrete Roughness Effects on Shuttle Hodel, A. S. Orbiter at Mach 6 – 37 Propellant Feed System Leak Detection: Lessons Learned From the Linear Aero- Direct Fault Tolerant RLV Altitude Con- Hamilton, H. Harris spike SR-71 Experiment (LASRE) – 93 trol: A Singular Perturbation Ap- N-Factor Computations for the X-33 Ve- proach – 40 Hattis, Philip D. hicle – 76 Air-breathing hypersonic vehicle guid- On-Line Computation of a Local Attain- Hamilton, H. Harris,II ance and control studies; An integrated able Moment Set for Reusable Launch Computational/Experimental Aeroheat- trajectory/control analysis methodology: Vehicles – 42 ing Predictions for X-33 Phase 2 Ve- Phase 1 – 171 Hodge, A. J. hicle – 110 Air-breathing hypersonic vehicle guid- Evaluation of Microcracking in Two Hamilton, Harris H., II ance and control studies: An integrated Carbon-Fiber/Epoxy-Matrix Composite Hypersonic Boundary-Layer Transition trajectory/control analysis methodology, Cryogenic Tanks – 72 phase 2 – 165 for X-33 Phase 2 Vehicle – 111 Hodge, A. Hanson, John M. Heeg, Jennifer Damage tolerance of candidate thermo- A Plan for Advanced Guidance and Con- Aerothermoelastic analysis of a NASP set composites for use on single stage to trol Technology for 2nd Generation Re- demonstrator model – 155 orbit vehicles – 148 usable Launch Vehicles – 52 Heinemann, J. M. Hoffman, Daniel Advanced Guidance and Control Meth- Health Monitoring Technology for Ther- Materials Research for High Speed Civil ods for Reusable Launch Vehicles: Test mal Protection Systems on Reusable Hy- Transport and Generic Hypersonics- Results – 30 personic Vehicles – 71 Metals Durability – 127

B-9 Hogenson, P. A. Horvath, Thomas J. An Approach to Establishing System Task 4 supporting technology. Part 1: Comparison of Methods for Determining Benefits for Technology in NASA’s Hy- Detailed test plan for leading edge tile Boundary Layer Edge Conditions for personics Investment Area – 67 development. Leading edge material de- Transition Correlations – 23 NASA’s Hypersonic Investment velopment and testing – 139 Computational/Experimental Aeroheat- Area – 47 Task 4 supporting technology. Part 2: ing Predictions for X-33 Phase 2 Ve- NASA’s Next Generation Launch Tech- Detailed test plan for thermal seals. Ther- hicle – 110 nology Program - Strategy and mal seals evaluation, improvement and Computational/Experimental Aeroheat- Plans – 4 test. CAN8-1, Reusable Launch Vehicle ing Predictions for X-33 – 112 (RLV), advanced technology demonstra- NASA’s Spaceliner Investment Area tor: X-33. Leading edge and seals ther- Hypersonic Boundary-Layer Transition Technology Activities – 81 mal protection system technology dem- for X-33 Phase 2 Vehicle – 111 Hulcher, A. B. onstration – 139 X-33 Computational Large Composite Structures Processing Hoh, Roger H. Aeroheating/Aerodynamic Predictions Technologies for Reusable Launch Ve- Assessment of flying-quality criteria for and Comparisons With Experimental hicles – 75 air-breathing aerospacecraft – 167 Data – 25 Humphrey, D. L. X-33 (Rev-F) Aeroheating Results of Test Holland, R. Oxidation-Reduction Resistance of Ad- 6770 in NASA Langley 20-Inch Mach 6 vanced Copper Alloys – 11 Engineering the Future of Full-Scale Pro- Air Tunnel – 96 pulsion Testing – 7 Humphrey, Donald H. X-33 Rev-F Turbulent Aeroheating Re- Oxidation Behavior of Copper Alloy Can- Holland, S. D. sults From Test 6817 in NASA Langley didates for Rocket Engine Applications Hyper-X Wind Tunnel Program – 109 20-Inch Mach 6 Air Tunnel and Compari- (Technical Poster) – 48 sons With Computations – 26 Holland, Scott D. Hunt, James L. Computational parametric study of X-33 Turbulent Aeroheating Measure- Airbreathing/Rocket Single-Stage-to- sidewall-compression scramjet inlet per- ments and Predictions – 32 Orbit Design Matrix – 124 formance at Mach 10 – 161 Howard, Richard T. Design and Analysis of a Two-Stage-to- Hollis, Brian R. An Advanced Video Sensor for Auto- Orbit Airbreathing Hypersonic Vehicle mated Docking – 61 129 Comparison of Methods for Determining Concept – Boundary Layer Edge Conditions for Howell, Joe Systems Challenges for Hypersonic Ve- Transition Correlations – 23 A Cryogenic Propellant Production Depot hicles – 117 for Low Earth Orbit – 81 X-33 Computational Hunt, William Aeroheating/Aerodynamic Predictions Hrinda, Glenn A. Configuration development study of the and Comparisons With Experimental Structures for the 3rd Generation Reus- OSU 1 hypersonic research ve- Data – 25 able Concept Vehicle – 69 hicle – 154 X-33 Rev-F Turbulent Aeroheating Re- Hsu, Su-Yuen Hunter, G. W. sults From Test 6817 in NASA Langley Preliminary Thermal-Mechanical Sizing A Hazardous Gas Detection System for 20-Inch Mach 6 Air Tunnel and Compari- of Metallic TPS: Process Development Aerospace and Commercial Applica- sons With Computations – 26 and Sensitivity Studies – 33 tions – 98 X-33 Turbulent Aeroheating Measure- Huang, George Hunter, J. E. ments and Predictions – 32 Hypersonic flows as related to the Na- GTX Reference Vehicle Structural Verifi- Holmes, R. tional Aerospace Plane – 136 cation Methods and Weight Sum- mary – 44 Lightweight Chambers for Thrust Assem- Huang, P. G. blies – 66 Hypersonic flows as related to the na- Huntley, J. D. tional aerospace plane – 162 Wingless Flight: The Lifting Body Holmes, Richard Story – 105 Lightweight Chambers for Thrust Assem- Huber, Frank W. blies – 82 Design and Analysis of Turbines for Hurlbert, Eric Space Applications – 16 NASA’s Space Launch Initiative Targets Honkala, Keith A. Experimental Performance Evaluation of Toxic Propellants – 64 Development of an integrated BEM ap- a Supersonic Turbine for Rocket Engine Hutsell, Lawrence J. proach for hot fluid structure interac- Applications – 10 tion – 169 An overview of aeroelasticity studies for Huddleston, D. the National Aerospace Plane – 160 Hood, Robert Computational Aerodynamics Analysis of Hutt, John NASA Exhibits – 59 Future Launch Vehicle Configura- NASA’s Hypersonic Investment Hopkins, Dale A. tions – 122 Area – 47 Neural Network and Regression Approxi- Huebner, L. D. Huttsell, Larry mations in High Speed Civil Transport Hyper-X Wind Tunnel Program – 109 Evaluation of Computational Aeroelastic- Aircraft Design Optimization – 106 ity Codes for Loads and Flutter – 82 Huestis, David L. Hopkins, Joshua B. Wireless Subsurface Microsensors for Huttsell, Lawrence J. Selection of Lockheed Martin’s Preferred Health Monitoring of Thermal Protection An overview of aeroelasticity studies for TSTO Configurations for the Space Systems on Hypersonic Vehicles – 69 the National Aerospace Plane – 160 Launch Initiative – 12 Hueter, Uwe Iliff, Kenneth W. Horn, Albert E. An Approach to Establishing System Flight Stability and Control and Perfor- Advanced Propulsion System Test and Benefits for Technologies In NASA’s mance Results from the Linear Aerospike Tell Capability – 133 Spaceliner Investment Area – 77 SR-71 Experiment (LASRE) – 101

B-10 Izon, Stephen James Johnson, W. S. Kimble, Michael C. Improvements in Thermal Protection Siz- Permeability of Impacted Coated Com- Performance Evaluation of Electro- ing Capabilities for TCAT: Conceptual posite Laminates – 49 chem’s PEM Fuel Cell Power Plant for Design for Advanced Space Transporta- NASA’s 2nd Generation Reusable tion Systems – 61 Johnston, Donald E. Launch Vehicle – 6 Assessment of flying-quality criteria for Jackson, Jerry E. Kirby, Mark air-breathing aerospacecraft – 167 The Control System for the X-33 Linear Automation of NDE on RSRM Metal Aerospike Engine – 105 Johnston, N. J. Components – 57 Jackson, Scott Large Composite Structures Processing Kiris, Cetin C. Autocommander: A Supervisory Control- Technologies for Reusable Launch Ve- Progress in Unsteady Turbopump Flow ler for Integrated Guidance and Control hicles – 75 Simulations Using Overset Grid Sys- for the 2nd Generation Reusable Launch Jones, Robert E. tems – 30 Vehicle – 40 Advanced Guidance and Control Meth- Progress in Unsteady Turbopump Flow Direct Fault Tolerant RLV Altitude Con- ods for Reusable Launch Vehicles: Test Simulations – 51 trol: A Singular Perturbation Ap- Results – 30 proach – 40 Time Dependent Simulation of Tur- Jorgensen, Catherine A. bopump Flows – 55 On-Board Generation of Three- Dimensional Constrained Entry Trajecto- International Space Station Evolution Kiris, Cetin ries – 35 Data Book – 87 NASA Exhibits – 59 On-Line Computation of a Local Attain- Jurns, John M. Time-Dependent Simulations of Tur- able Moment Set for Reusable Launch Testing of Densified Liquid Hydrogen bopump Flows – 38 Vehicles – 42 Stratification in a Scale Model Propellant Kirkici, H. Reusable Launch Vehicle Attitude Con- Tank – 67 High Voltage Design Guidelines: A 61 trol Using a Time-Varying Sliding Mode Kabis, Hanee Timely Update – Control Technique – 40 Design and Analysis of a Two-Stage-to- Kiser, J. D. Jankovsky, Amy L. Orbit Airbreathing Hypersonic Vehicle Fabrication and Testing of Ceramic Ma- Real-Time Sensor Validation, Signal Re- Concept – 129 trix Composite Rocket Propulsion Com- construction, and Feature Detection for ponents – 75 an RLV Propulsion Testbed – 98 Kalia, Prince A Quantitative Reliability, Maintainability Kiser, J. Douglas Jarvis, Michele and Supportability Approach for NASA’s Characterization of C/SiC Ceramic Ma- Evaluation of the Linear Aerospike SR-71 Second Generation Reusable Launch trix Composites (CMCs) with Novel Inter- Experiment (LASRE) Oxygen Sen- Vehicle – 41 face Fiber Coatings – 45 sor – 91 Kiser, James Douglas Kamhawi, Hani Jaskowiak, M. H. Characterization of C/SiC Ceramic Ma- Status of the RBCC Direct-Connect Fabrication and Testing of Ceramic Ma- trix Composites (CMCs) with Novel Inter- Mixer Combustor Experiment – 54 trix Composite Rocket Propulsion Com- face Fiber Coatings – 48 ponents – 75 Karr, G. R. Kittredge, Sheryl Jensen, Brian J. Research Reports: 1996 NASA/ASEE Life Cycle Systems Engineering Ap- Cryopumping in Cryogenic Insulations Summer Faculty Fellowship Pro- proach to NASA’s 2nd Generation Reus- for a Reusable Launch Vehicle – 2 gram – 102 able Launch Vehicle – 39 Jensen, J. K. Karr, Katherine L. Systems Engineering Approach to Tech- A study of facilities and fixtures for testing Hypervelocity Impact Test Results for a nology Integration for NASA’s 2nd Gen- of a high speed civil transport wing com- Metallic Thermal Protection Sys- eration Reusable Launch Vehicle – 23 ponent – 127 tem – 10 Klem, Mark D. Johnson, S. M. Karunaratne, K. S. Experimental Evaluation of a Subscale Gaseous Hydrogen/Gaseous Oxygen Fractographic Analysis of HfB2-SiC and Health Monitoring Technology for Ther- Coaxial Rocket Injector – 34 ZrB2-SiC Composites – 57 mal Protection Systems on Reusable Hy- Johnson, Sylvia M. personic Vehicles – 71 Klop, Jeffrey Thermal Protection Materials for Reentry The Control System for the X-33 Linear Karunaratne, K. and Planetary Applications – 68 Aerospike Engine – 105 Active Wireless Temperature Sensors for Thermal Protection Materials for Reentry Aerospace Thermal Protection Sys- Klyde, David H. Applications – 59 tems – 46 Advanced aeroservoelastic stabilization techniques for hypersonic flight ve- Johnson, Theodore F. Keith, Amy hicles – 163 Cryopumping in Cryogenic Insulations An Overview of In-Stu Treatability Stud- for a Reusable Launch Vehicle – 2 Kmiec, Tom ies at Marshall Space Flight Center, 2nd Generation Reusable Launch Ve- Effects of Stiffening and Mechanical Huntsville, Alabama – 78 82 Load on Thermal Buckling of Stiffened hicle (2G RLV) – Cylindrical Shells – 121 Keith, Theo Knight, D. Computations of Internal and External High Temperature Polyimide Materials in A Hazardous Gas Detection System for Axisymmetric Nozzle Aerodynamics at Extreme Temperature Environ- Aerospace and Commercial Applica- Transonic Speeds – 4 ments – 73 tions – 98 Thermal Structures Technology Develop- Keller, Peter C. Kohout, Lisa L. ment for Reusable Launch Vehicle Cryo- Advanced High Temperature Structural Fuel Cell Activities at the NASA Glenn genic Propellant Tanks – 107 Seals – 43 Research Center – 49

B-11 Komar, D. R. Krupp, D Towers for Earth Launch – 29 Reusable Rocket Engine Operability A guidance and control assessment of Lang, J. Modeling and Analysis – 101 three vertical landing options for Fabrication and Testing of Ceramic Ma- RLV – 136 Kontinos, Dean trix Composite Rocket Propulsion Com- Thermal Protection Materials for Reentry Krupp, Don R. ponents – 75 Advanced Guidance and Control Meth- Applications – 59 Langehough, M. A. ods for Reusable Launch Vehicles: Test Korte,J.J. Results – 30 Personnel launch system autoland de- Multidisciplinary Approach to Aerospike velopment study – 175 Kudlac, Maureen T. Nozzle Design – 120 A summary of the slush hydrogen tech- Lansing, M. D. Multidisciplinary Optimization Branch Ex- nology program for the National Aero- Thermographic Analysis of Composite perience Using iSIGHT Software – 91 Space Plane – 140 Cobonds on the X-33 – 75 Kosanchick, Melvin, III Kussoy, Marvin Larson, Richard R. Configuration development study of the Hypersonic flows as related to the Na- Flight Testing the Linear Aerospike OSU 1 hypersonic research ve- tional Aerospace Plane – 136 SR-71 Experiment (LASRE) – 100 hicle – 154 Hypersonic flows as related to the na- Laruelle, Gerard Kosareo, Daniel N. tional aerospace plane – 162 Systems Challenges for Hypersonic Ve- Affordable Flight Demonstration of the Hypersonic flows as related to the Na- hicles – 117 GTX Air-Breathing SSTO Vehicle Con- tional Aerospace plane – 172 Latino, Carl D. cept – 46 Kwak, Dochan Using Neural Networks in Decision Mak- Structural Sizing of a 25,000-lb Payload, Automated CFD Parameter Studies on ing for a Reconfigurable Electro Me- Air-Breathing Launch Vehicle for Single- Distributed Parallel Computers – 36 chanical Actuator (EMA) – 68 Stage-to-Orbit – 85 Calculation of Supersonic Combustion Laue, Jay H. Kosareo, Daniel Using Implicit Schemes – 30 Space Transportation Systems Tech- SRM-Assisted Trajectory for the GTX nologies – 70 Progress in Unsteady Turbopump Flow Reference Vehicle – 47 Simulations Using Overset Grid Sys- Laue, Jay Kosnareo, Daniel N. tems – 30 Second Generation RLV Program – 71 Affordable Flight Demonstration of the Progress in Unsteady Turbopump Flow Laufenberg, Larry GTX Air-Breathing SSTO Vehicle Con- Simulations – 51 Accelerating the Design of Space Ve- cept – 26 Time Dependent Simulation of Tur- hicles – 10 Kotula, Paul bopump Flows – 55 Lavelle, Thomas M. Thermal Protection Materials for Reentry Time-Dependent Simulations of Tur- Neural Network and Regression Approxi- Applications – 59 bopump Flows – 38 mations in High Speed Civil Transport Kourtides, Demetrius Laganelli, Tony Aircraft Design Optimization – 106 Evaluation of Thermal Control Coatings A Thermal Management Systems Model Lawrence, D. A. for Flexible Ceramic Thermal Protection for the NASA GTX RBCC Concept – 53 Autocommander: A Supervisory Control- 117 Systems – Lai,H.T. ler for Integrated Guidance and Control for the 2nd Generation Reusable Launch Kovach, M. P. Computation of H2/air reacting flowfields Vehicle – 40 Composite,Cryogenic, Conformal, Com- in drag-reduction external combus- mon Bulkhead, Aerogel-Insulated Tank tion – 161 Direct Fault Tolerant RLV Altitude Con- trol: A Singular Perturbation Ap- (CBAT) – 74 Lai, Steven H.-Y. proach – 40 Kowalkowski, Matthew K. The Center for Aerospace Research: A NASA Center of Excellence at North Lawrence, Douglas A. X-33 (Rev-F) Aeroheating Results of Test Carolina Agricultural and Technical State 6770 in NASA Langley 20-Inch Mach 6 Integrated G and C Implementation University – 162 Air Tunnel – 96 within IDOS: A Simulink Based Reusable Laithier, Frederic Launch Vehicle Simulation – 11 Kremer, Steven E. Demonstration of Advanced C/SiC Lawrence, Tim The X-33 Extended Flight Test Cooled Ramp – 56 Ceramic Matrix Composite Cooled Range – 99 Lak, Tibor Nozzle Material Development Pro- Kren, Lawrence A. Task 4 supporting technology. Densifica- gram – 21 Hot dynamic test rig for measuring hyper- tion requirements definition and test ob- Lawrence, Timothy sonic engine seal flow and durabil- jectives. Propellant densification require- Demonstration of Advanced C/SiC ity – 151 139 ments definition – Cooled Ramp – 56 Krivanek, Thomas M. Lance, D. Lazaroff, Scott M. Affordable Flight Demonstration of the Damage tolerance of candidate thermo- Linear Aerospike SR-71 Experiment GTX Air-Breathing SSTO Vehicle Con- set composites for use on single stage to (LASRE): Aerospace Propulsion Hazard cept – 26 orbit vehicles – 148 Mitigation Systems – 102 Status of the RBCC Direct-Connect Lance, N. Mixer Combustor Experiment – 54 Human Transportation System (HTS) Leahy, Frank study, volume 2 – 155 Natural Atmospheric Environment Model Kroo, I. M. Development for the National Aeronau- Comparison of Two Multidisciplinary Op- Landis, Geoffrey A. tics and Space Administration’s Second timization Strategies for Launch-Vehicle High Altitude Launch for a Practical Generation Reusable Launch Ve- Design – 123 SSTO – 29 hicle – 57

B-12 Lee, Charles Levine, Stanley R. Love, Mike Testing the Preliminary X-33 Navigation Ceramic Matrix Composites (CMC) Life Evaluation of Computational Aeroelastic- System – 84 Prediction Method Development – 88 ity Codes for Loads and Flutter – 82 Lee, Howard Tantalum Addition to Zirconium Diboride Lovell, N. 18 Predictor-Corrector Entry Guidance for for Improved Oxidation Resistance – Integrated Technology Assessment Cen- Reusable Launch Vehicles – 74 Levy, Lionel ter (ITAC) Update – 42 Hypersonic flows as related to the Na- Lee, J. Lovell, T. Alan Development of Metal Matrix Compos- tional Aerospace plane – 172 A parametric sensitivity study for single- ites for NASA’s Advanced Propulsion Lewis, Clark H. stage-to-orbit hypersonic vehicles using Systems – 74 PNS predictions of external/internal hy- trajectory optimization – 153 Lightweight Chambers for Thrust Assem- personic flows for NASP propulsion ap- Effect of Aeropropulsive Interactions and blies – 66 plications – 157 Design Sensitivities on Optimal Hyper- Lee, Jonathan A. Liechty, Derek S. sonic Ascent Trajectories – 125 Development of Aluminum Composites Comparison of Methods for Determining for a Rocket Engine’s Lightweight Thrust Boundary Layer Edge Conditions for Lowrey, Alan Cell – 80 Transition Correlations – 23 Transportation Systems Analyses (TSA): Highly Reusable Space Transportation X-33 (Rev-F) Aeroheating Results of Test Lee, Jonathan (HRST). A preliminary programmatic as- 6770 in NASA Langley 20-Inch Mach 6 Lightweight Chambers for Thrust Assem- sessment – 131 blies – 82 Air Tunnel – 96 Lu, P. Lee, Seung-Hee Lin, Jeff Using CFD as Rocket Injector Design Direct Fault Tolerant RLV Altitude Con- Full Envelope Reconfigurable Control trol: A Singular Perturbation Ap- Design for the X-33 Vehicle – 72 Tool: Recent Progress at Marshall Space Flight Center – 3 proach – 40 Lee, Thomas J. Lineberry, John Lu, Ping Support to 2nd Generation RLV Propul- sion Project Office – 60 Magnetohydrodynamic Augmented Pro- An inverse dynamics approach to trajec- pulsion Experiment – 52 tory optimization for an aerospace Lee, W. Y. plane – 163 Linton, Donald F. Optimal trajectories for an aerospace Analytical solutions to constrained hyper- plane. Part 1: Formulation, results, and Advanced Range Safety System for High sonic flight trajectories – 162 analysis – 176 Energy Vehicles – 15 On-Board Entry Trajectory Planning Ex- Optimal trajectories for an aerospace Lister, Darlene panded to Sub-orbital Flight – 12 plane. Part 2: Data, tables, and Wingless Flight: The Lifting Body graphs – 176 Story – 105 On-Board Generation of Three- Lee, Young C. Litchford, Ron J. Dimensional Constrained Entry Trajecto- ries – 35 X-33 XRS-2200 Linear Aerospike Engine Magnetohydrodynamic Augmented Pro- Sea Level Plume Radiation – 66 pulsion Experiment – 52 Reconfigurable Flight Control Designs With Application to the X-33 Ve- Lee, Young-Ching Littlefield, Alan C. hicle – 95 Analysis of X-33 Linear Aerospike Plume Design, Development,and Testing of Um- Induced Base-Heating Environment Dur- billical System Mechanisms for the X-33 Trajectory optimization for the national ing Power-Pack Out – 29 Advanced Technology Demonstra- aerospace plane – 156 tor – 90 Base-Bleed Effect on X-33 Aerospike Trajectory optimization for the National Plume Induced Base-Heating Environ- Liu, C. C. aerospace plane – 160 ment During Power-Pack Out – 25 A Hazardous Gas Detection System for Two Reconfigurable Flight-Control De- Aerospace and Commercial Applica- Lepsch, R. A. sign Methods: Robust Servomechanism tions – 98 Comparison of Two Multidisciplinary Op- and Control Allocation – 76 timization Strategies for Launch-Vehicle Liu, J. Lu, Tina Design – 123 Unic Unstructured CFD Methodology De- velopment – 82 Task 4 supporting technology. Part 2: Lepsch, Roger A., Jr. Detailed test plan for thermal seals. Ther- Dual-Fuel Propulsion in Single-Stage Ad- Lockwood, Mary Kae mal seals evaluation, improvement and vanced Manned Launch System Ve- Design and Analysis of a Two-Stage-to- test. CAN8-1, Reusable Launch Vehicle hicle – 125 Orbit Airbreathing Hypersonic Vehicle (RLV), advanced technology demonstra- Concept – 129 Subsonic Aerodynamic Characteristics of tor: X-33. Leading edge and seals ther- a Circular Body Earth-to-Orbit Ve- Loehman, R. E. mal protection system technology demonstration – 139 hicle – 128 Fractographic Analysis of HfB2-SiC and Lessman, Craig ZrB2-SiC Composites – 57 Lyles, Garry M. Testing the Preliminary X-33 Navigation Loehman, Ron An Approach to Establishing System System – 84 Thermal Protection Materials for Reentry Benefits for Technologies In NASA’s Applications – 59 Spaceliner Investment Area – 77 Levack, Daniel J. H. Advanced Low-Cost O2/H2 Engines for Lomas, James J. Main Engine Prototype Development for the SSTO Application – 123 Testing the Preliminary X-33 Navigation 2nd Generation RLV RS-83 – 43 System – 84 Levack, Daniel NASA’s Spaceliner Investment Area Technology Activities – 81 Alternate Propulsion Subsystem Con- Lorenzini, E. cepts Tripropellant Comparison Tether Transportation System On-Orbit Propulsion System Project Study – 132 Study – 109 Overview – 44

B-13 Lyles, Garry Mammano, J. McClinton, Charles R. Launch Vehicle Sizing Benefits Utilizing Thermal Management Design for the Technology Roadmap for Dual-Mode Main Propulsion System Crossfeed and X-33 Lifting Body – 104 Scramjet Propulsion to Support Space- Project Status – 34 Access Vision Vehicle Develop- Mandl, William ment – 38 Lyon, Jeffery A. Mosad and Stream Vision For A Telero- Abort performance for a winged-body botic, Flying Camera System – 59 McClinton, Charles 137 NASA’s Hypersonic Investment single-stage to orbit vehicle – Mankins, John Area – 47 Lyons, Valerie J. General Public Space Travel and Tour- Towers for Earth Launch – 29 ism – 96 Mccomb, Harvey G., Jr. Structures and materials technology for Lyons, Valerie Markopoulos, N. hypersonic aerospacecraft – 178 Fuel Cell Activities at the NASA Glenn Rapid near-optimal aerospace plane tra- McConnaughey, P. Research Center – 49 jectory generation and guidance – 171 Computational Aerodynamics Analysis of High Altitude Launch for a Practical Marschall, Jochen Future Launch Vehicle Configura- SSTO – 29 Evaluation of Thermal Control Coatings tions – 122 for Flexible Ceramic Thermal Protection MacConochie, Ian O. Systems – 117 McCool, Alex Subsonic Aerodynamic Characteristics of Automation of NDE on RSRM Metal a Circular Body Earth-to-Orbit Ve- Materials Testing on the DC-X and DC- Components – 57 hicle – 128 XA – 121 McCue, Terry Mackall, Dale A. Martin, C. J., Jr. Characterization of C/SiC Ceramic Ma- The X-33 Extended Flight Test A study of facilities and fixtures for testing trix Composites (CMCs) with Novel Inter- Range – 99 of a high speed civil transport wing com- face Fiber Coatings – 48 ponent – 127 Magdaleno, Raymond E. McCurdy, D. R. Advanced aeroservoelastic stabilization Martin, Carl J. GTX Reference Vehicle Structural Verifi- techniques for hypersonic flight ve- Parametric Weight Comparison of Ad- cation Methods and Weight Sum- hicles – 163 vanced Metallic, Ceramic Tile, and Ce- mary – 44 ramic Blanket Thermal Protection Sys- Maguire, James F. tems – 88 McCurdy, David R. Development of a Pressure Box to Evalu- Preliminary Sizing of Vertical Take-Off Martinovic, Zoran N. ate Reusable-Launch-Vehicle Rocket-Based Combined-Cycle Pow- Cryogenic-Tank Panels – 132 A Procedure for Structural Weight Esti- ered Launch Vehicles – 85 mation of Single Stage to Orbit Launch Mahadevan, Sankaran Vehicles (Interim User’s Manual) – 44 McCurry, J. B. Advanced transportation system studies System Risk Assessment and Allocation Mathias, Donovan in Conceptual Design – 27 technical area 2 (TA-2): Heavy lift launch Multi-Disciplinary Analysis for Future vehicle development – 130 Mahoney, Michael Launch Systems Using NASA’s Ad- Integrated G and C Implementation vanced Engineering Environment McCurry, J. within IDOS: A Simulink Based Reusable (AEE) – 17 Advanced transportation system studies technical area 2(TA-2): Heavy lift launch Launch Vehicle Simulation – 11 Matlock, Steve vehicle development – 130 Majumdar, Alok 2nd Generation Reusable Launch Ve- A Novel Approach for Modeling Chemical hicle (2G RLV) – 82 McElroy, Bill An Overview of In-Stu Treatability Stud- Reaction in Generalized Fluid System Maul, William A., III Simulation Program – 13 ies at Marshall Space Flight Center, Real-Time Sensor Validation, Signal Re- Huntsville, Alabama – 78 Makel, D. B. construction, and Feature Detection for McGhee, D. S. A Hazardous Gas Detection System for an RLV Propulsion Testbed – 98 A Strategy for Integrating a Large Finite Aerospace and Commercial Applica- Mavris, Dimitri N. Element Model Using MSC tions – 98 Technology and Risk Assessment Using NASTRAN/PATRAN: X-33 Lessons Malchow, Harvey L. a Robust Design Simulation Methodol- Learned – 97 ogy – 121 Air-breathing hypersonic vehicle guid- McIver, K. ance and control studies; An integrated Mayers, Tom Large Composite Structures Processing trajectory/control analysis methodology: 2nd Generation Reusable Launch Ve- Technologies for Reusable Launch Ve- 171 Phase 1 – hicle (2G RLV) – 82 hicles – 75 Air-breathing hypersonic vehicle guid- Mazurkivich, P. McKenzie, Patrick M. ance and control studies: An integrated Launch Vehicle Sizing Benefits Utilizing trajectory/control analysis methodology, Orbital Space Plane (OSP) Pro- Main Propulsion System Crossfeed and gram – 5 phase 2 – 165 Project Status – 34 McLaughlin, Don Mallapragada, P. McBain, Michael C. The StarBooster System: A Cargo Air- Unic Unstructured CFD Methodology De- Composite Development and Applica- craft for Space – 92 velopment – 82 tions for RLV Tankage – 16 McMahon, W. M. Malone, John B. McCaleb, Rebecca Composite,Cryogenic, Conformal, Com- NASP aeroservothermoelasticity stud- An Overview of In-Stu Treatability Stud- mon Bulkhead, Aerogel-Insulated Tank ies – 174 ies at Marshall Space Flight Center, (CBAT) – 74 Huntsville, Alabama – 78 Maloney, Jill M. Large Composite Structures Processing Analytical comparison of three stiffened McClinton, C. R. Technologies for Reusable Launch Ve- panel concepts – 135 Hyper-X Wind Tunnel Program – 109 hicles – 75

B-14 Mcminn, John D. Menter, Florian Mizukami, Masashi Hypersonic vehicle control law develop- Hypersonic flows as related to the Na- Linear Aerospike SR-71 Experiment ment using H infinity and mu- tional Aerospace Plane – 136 (LASRE): Aerospace Propulsion Hazard synthesis – 164 Mitigation Systems – 102 Mercer, Carolyn Hypersonic vehicle model and control Overview of GRC’s Advanced Sensor Propellant Feed System Leak Detection: law development using H(infinity) and and Instrumentation Development – 41 Lessons Learned From the Linear Aero- micron synthesis – 144 spike SR-71 Experiment (LASRE) – 93 Merriam, M. L. McNeal, Curtis Cost Per Pound From Orbit – 47 Moerder, Daniel D. NASA’s Space Launch Initiative Targets Constrained minimization of smooth Toxic Propellants – 64 Merriam, Marshall functions using a genetic algo- New V and V Tools for Diagnostic Mod- rithm – 141 Mcnelis, Nancy B. eling Environment (DME) – 58 A summary of the slush hydrogen tech- Moes, Timothy R. nology program for the National Aero- Merski, N. R. A Base Drag Reduction Experiment on Space Plane – 140 Aerodynamic Characteristics of a Vertical the X-33 Linear Aerospike SR-71 Experi- Takeoff Vertical Landing (VTVL) Single ment (LASRE) Flight Program – 97 McNelis, Nancy Stage to Orbit Vehicle from M(sub infin- NASA’s Advanced Space Transportation ity) = 0.1 to 10 – 122 Flight Stability and Control and Perfor- Program: RTA Project Summary – 42 mance Results from the Linear Aerospike Merski, N. Ronald, Jr. SR-71 Experiment (LASRE) – 101 Mcruer, Duane T. Infrared Sensing Aeroheating Flight Ex- Flight Testing the Linear Aerospike Advanced aeroservoelastic stabilization periement: STS-96 Flight Results – 84 techniques for hypersonic flight ve- SR-71 Experiment (LASRE) – 100 hicles – 163 Infrared Sensing Aeroheating Flight Ex- periment: STS-96 Flight Results – 86 Monaghan, Richard C. Assessment of flying-quality criteria for Flight Testing the Linear Aerospike air-breathing aerospacecraft – 167 Messinger, Ross SR-71 Experiment (LASRE) – 100 Thermal-Mechanical Cyclic Test of a Monell, Donald Meador, Michael Composite Cryogenic Tank for Reusable Characterization of Polyimide Foams for Launch Vehicles – 23 Multi-Disciplinary Analysis for Future Ultra-Lightweight Space Struc- Launch Systems Using NASA’s Ad- tures – 22 Meyer, Claudia M. vanced Engineering Environment Real-Time Sensor Validation, Signal Re- (AEE) – 17 Mecholsky, J.J., Jr. construction, and Feature Detection for Moore, A. A. Fractographic Analysis of HfB2-SiC and an RLV Propulsion Testbed – 98 ZrB2-SiC Composites – 57 Optimal technology investment strate- Meyer, Rudolf X. gies for a reusable launch vehicle – 134 Megna, Thomas D. Ion engine propelled Earth-Mars cycler Moran, Matthew E. Lockheed Martin Response to the OSP with nuclear thermal propelled transfer Challenge – 20 vehicle, volume 2 – 144 A summary of the slush hydrogen technology program for the National Aero- Mehrotra, P. Miele, Angelo Space Plane – 140 Compute as Fast as the Engineers Can Optimal trajectories for an aerospace Moran, Patrick Think! ULTRAFAST COMPUTING TEAM plane. Part 1: Formulation, results, and FINAL REPORT – 90 analysis – 176 NASA Exhibits – 59 Mehta, Unmeel B. Optimal trajectories for an aerospace Moravie, Michel A Two-Stage-to-Orbit Spaceplane Con- plane. Part 2: Data, tables, and Earth-to-Orbit Rocket Propulsion – 17 cept With Growth Potential – 80 graphs – 176 Morgan, Carolyn Air-breathing aerospace plane develop- Miller, Jeffrey H. Statistical Methods for Rapid Aerother- ment essential: Hypersonic propulsion Hot dynamic test rig for measuring hyper- mal Analysis and Design Technol- flight tests – 138 sonic engine seal flow and durabil- ogy – 55 The aerospace plane design challenge: ity – 151 Morris, Charles Credible computational fluid dynamics Milos, Frank S. Integrated Technology Assessment Cen- 176 results – Active Wireless Temperature Sensors for ter (ITAC) Update – 42 Mei, Chuh Aerospace Thermal Protection Sys- Moses, Paul 46 Panel Flutter and Sonic Fatigue Analysis tems – Design and Analysis of a Two-Stage-to- for RLV – 77 Health Monitoring Technology for Ther- Orbit Airbreathing Hypersonic Vehicle Melko, Joseph mal Protection Systems on Reusable Hy- Concept – 129 personic Vehicles – 71 Ion engine propelled Earth-Mars cycler Mouyos, William with nuclear thermal propelled transfer Wireless Subsurface Microsensors for Flight Demonstration of X-33 Vehicle vehicle, volume 2 – 144 Health Monitoring of Thermal Protection Health Management System Compo- Systems on Hypersonic Vehicles – 69 Melton, Gregory S. nents on the F/A-18 Systems Research 15 Design, Development,and Testing of Um- Wireless Subsurface Sensors for Health Aircraft – billical System Mechanisms for the X-33 Monitoring of Thermal Protection Sys- X-33/RLV System Health Management/ Advanced Technology Demonstra- tems on Hypersonic Vehicles – 68 Vehicle Health Management – 107 tor – 90 Mital, Subodh K. Munafo, Paul M. Menter, F. Ceramic Matrix Composites (CMC) Life Demonstration of Advanced C/SiC Hypersonic flows as related to the na- Prediction Method Development – 88 Cooled Ramp – 56 tional aerospace plane – 162 Mitchell, Daniel W. Development of Aluminum Composites Hypersonic flows as related to the Na- Testing the Preliminary X-33 Navigation for a Rocket Engine’s Lightweight Thrust tional Aerospace plane – 172 System – 84 Cell – 80

B-15 Findings from the X-33 Hydrogen Tank Neal, Bradford A. Newquist, Charles W. Failure Investigation – 60 Flight Testing the Linear Aerospike Advanced High Temperature Structural Manufacturing Process Simulation of SR-71 Experiment (LASRE) – 100 Seals – 43 Large-Scale Cryotanks – 36 Propellant Feed System Leak Detection: Nguyen, Dalton National Rocket Propulsion Materials Lessons Learned From the Linear Aero- Thermal/Fluid Analysis of a Composite 93 Plan: A NASA, Department of Defense, spike SR-71 Experiment (LASRE) – Heat Exchanger for Use on the RLV and Industry Partnership – 73 Neely, M. A. Rocket Engine – 37 Munafo, Paul Integrated Technology Assessment Cen- Nguyen, Viet ter (ITAC) Update – 42 Permeability After Impact Testing of Dynamics and Stability and Control Composite Laminates – 57 Nelson, M. L. Characteristics of the X-37 – 62 Permeability Testing of Impacted Com- Compute as Fast as the Engineers Can Niedermeyer, M. posite Laminates for Use on Reusable Think! ULTRAFAST COMPUTING TEAM Launch Vehicles – 80 FINAL REPORT – 90 X-33 LH2 Tank Failure Investigation Findings – 74 X-33 LH2 Tank Failure Investigation Nelson, Michael A. Findings – 58 Rocket Engine Health Management: Niedermeyer, Melinda X-33 Tank Failure During Autoclave Fab- Early Definition of Critical Flight Mea- Findings from the X-33 Hydrogen Tank rication – 60 surements – 19 Failure Investigation – 60 Munkres, Randy Nelson, Nick X-33 LH2 Tank Failure Investigation Findings – 14 Lockheed Martin Response to the OSP Configuration development study of the Challenge – 20 OSU 1 hypersonic research ve- Nixon, Robert F. Murphy, Kelly J. hicle – 154 Advanced Low-Cost O2/H2 Engines for X-33 Computational Nelson, Stacy the SSTO Application – 123 Aeroheating/Aerodynamic Predictions New V and V Tools for Diagnostic Mod- Noll, T. E. and Comparisons With Experimental eling Environment (DME) – 58 Data – 25 NASP aeroservothermoelasticity stud- Survey of NASA V and V ies – 174 Murrow, Harold N. Processes/Methods – 58 Structures and materials technology for Noll, Thomas E. hypersonic aerospacecraft – 178 V and V of Advanced Systems at An overview of aeroelasticity studies for NASA – 58 the National Aerospace Plane – 160 Murthy, Pappu L. Ceramic Matrix Composites (CMC) Life Nemeth, Noel N. Noneman, Steven R. Prediction Method Development – 88 Design and Analysis of UHTC Leading Operations Analysis of the 2nd Genera- Edge Attachment – 44 Myers, David E. tion Reusable Launch Vehicle – 43 Parametric Weight Comparison of Ad- Nettles, A. T Norris, Rick B. vanced Metallic, Ceramic Tile, and Ce- A Damage Resistance Comparison Be- The design and evolution of the beta ramic Blanket Thermal Protection Sys- tween Candidate Polymer Matrix Com- two-stage-to-orbit horizontal takeoff and tems – 88 posite Feedline Materials – 88 landing launch system – 159 Myers, Thomas T. Nettles, A. T. Nowak, Robert J. Advanced aeroservoelastic stabilization Damage tolerance of candidate thermo- Hypersonic Boundary-Layer Transition techniques for hypersonic flight ve- set composites for use on single stage to for X-33 Phase 2 Vehicle – 111 hicles – 163 orbit vehicles – 148 X-33 Computational Assessment of flying-quality criteria for Permeability Testing of Impacted Com- Aeroheating/Aerodynamic Predictions air-breathing aerospacecraft – 167 posite Laminates for Use on Reusable and Comparisons With Experimental Nadell, Shari-Beth Launch Vehicles – 83 Data – 25 Mission and sizing analysis for the Beta 2 Nettles, Alan T. two-stage-to-orbit vehicle – 168 Ogbuji, Linus U. J. Permeability After Impact Testing of Oxidation Behavior of Copper Alloy Can- Naftel, J. Christopher Composite Laminates – 21 didates for Rocket Engine Applications Analysis of the staging maneuver and Permeability Testing of Impacted Com- (Technical Poster) – 48 booster glideback guidance for a two- posite Laminates for Use on Reusable stage, winged, fully reusable launch ve- Oka, Kris Launch Vehicles – 80 hicle – 159 The Evolution of Flexible Insulation as X-33 Tank Failure During Autoclave Fab- Natividad, Roderick Thermal Protection Systems for Reus- rication – 60 able Launch Vehicles: AFRSI (Advanced Thermal Structures Technology Develop- Flexible Reusable Surface Insulation) to ment for Reusable Launch Vehicle Cryo- Nettles, Alan CRI (Conformal Reusable Insula- genic Propellant Tanks – 107 Impact Characteristics of Candidate Ma- tion) – 67 Naughton, Jonathan W. terials for Single-Stage-to-Orbit (SSTO) 123 Olds, John R. Reduction of Base Drag on Launch Ve- Technology – A Method of Integrating Aeroheating into hicles – 50 Nettles, A.T. Conceptual Reusable Launch Vehicle Wind-Tunnel Investigations of Blunt- Permeability After Impact Testing of Design: Evaluation of Advanced Thermal Body Drag Reduction Using Forebody Composite Laminates – 57 Protection Techniques for Future Reus- Surface Roughness – 85 Neudeck, P. G. able Launch Vehicles – 77 Navard, Sharon E. A Hazardous Gas Detection System for Evaluation of Advanced Thermal Protec- An Overview of PRA with Applications to Aerospace and Commercial Applica- tion Techniques for Future Reusable Aerospace Systems – 51 tions – 98 Launch Vehicles – 62

B-16 Highly reusable space transportation: Palac, Donald T. Patniak, Surya N. Approaches for reducing ETO launch NASA Glenn Research Center’s Hyper- Neural Network and Regression Approxi- costs to $100 - $200 per pound of pay- sonic Propulsion Program – 95 mations in High Speed Civil Transport load – 135 106 Performance Evaluation of the NASA Aircraft Design Optimization – Highly Reusable Space Transportation GTX RBCC Flowpath – 70 Patterson, B. System Concept Evaluation (The Argus Palaszewski, Bryan Ceramic Matrix Composite Cooled Launch Vehicle) – 103 Preliminary Assessment of Using Gelled Nozzle Material Development Pro- Improvements in Thermal Protection Siz- and Hybrid Propellant Propulsion for gram – 21 ing Capabilities for TCAT: Conceptual VTOL/SSTO Launch Systems – 104 Pecheur, Charles Design for Advanced Space Transporta- New V and V Tools for Diagnostic Mod- tion Systems – 61 Pallix, Joan B. Wireless Subsurface Microsensors for eling Environment (DME) – 58 Results of a Rocket-Based Combined- Health Monitoring of Thermal Protection Survey of NASA V and V Cycle SSTO Design Using Parametric Systems on Hypersonic Vehicles – 69 Processes/Methods – 58 MDO Methods – 121 Pallix, Joan V and V of Advanced Systems at Supercharged Ejector Ramjet with Ma- Evaluation of Thermal Control Coatings NASA – 58 glifter Launch Assist (for the Argus Con- for Flexible Ceramic Thermal Protection 120 Peinemann, Manfred cept) – Systems – 117 System Level Uncertainty Assessment Liquid Engine Test Facilities Assess- Materials Testing on the DC-X and DC- for Collaborative RLV Design – ment – 1 46 XA – 121 Pelaccio, Dennis G. System Sensitivity Analysis Applied to Pamadi, Bandu N. the Conceptual Design of a Dual-Fuel Preliminary Assessment of Using Gelled A simple analytical aerodynamic model Rocket SSTO – 125 and Hybrid Propellant Propulsion for of Langley Winged-Cone Aerospace VTOL/SSTO Launch Systems – 104 Olds, John Robert Plane concept – 145 Peterson, W. Multidisciplinary design techniques ap- Constrained minimization of smooth plied to conceptual aerospace vehicle Liquid flyback booster pre-phase: A study functions using a genetic algo- assessment – 146 design – 156 rithm – 141 Petko, Jeanne F. Olds, John Pandey, Awadh Characterization of C/SiC Ceramic Ma- Options for flight testing rocket-based Selection and Evaluation of An Alloy for trix Composites (CMCs) with Novel Inter- combined-cycle (RBCC) engines – 130 Nozzle Application – 22 face Fiber Coatings – 45 OLeary, Robert Pandya, Shishir Phillips, Alan Preliminary Assessment of Using Gelled Automated CFD Parameter Studies on Analysis of Parallel Burn Without Cross- and Hybrid Propellant Propulsion for Distributed Parallel Computers – 36 VTOL/SSTO Launch Systems – 104 feed TSTO RLV Architectures and Com- Pannell, Bill parison to Parallel Burn With Crossfeed Oleson, Steven An Approach to Establishing System and Series Burn Architectures – 19 Advanced Electric Propulsion for RLV Benefits for Technologies In NASA’s Phillips, Steven Launched Geosynchronous Space- Spaceliner Investment Area – 77 Manufacturing Process Simulation of craft – 90 An Approach to Establishing System Large-Scale Cryotanks – 36 Benefits for Technology in NASA’s Hy- ONeil, Daniel Phillips, W. P. personics Investment Area – 67 General Public Space Travel and Tour- Subsonic-to-Hypersonic Aerodynamic 96 ism – Pao, Jenn-Louh Characteristics for a Winged, Circular- Opila, Eliizabeth J. Design and Analysis of a Two-Stage-to- Body, Single-Stage-to-Orbit Spacecraft Tantalum Addition to Zirconium Diboride Orbit Airbreathing Hypersonic Vehicle Configuration – 124 Concept – 129 for Improved Oxidation Resistance – 18 Pichon, T. Opila, Elizabeth J. Papadopoulos, P. E. Ceramic Matrix Composite Cooled Ceramic Matrix Composites (CMC) Life A Geometry-Centered Approach For Nozzle Material Development Pro- Prediction Method Development – 88 Two-Stage-To-Orbit Vehicle Synthe- gram – 21 sis – 45 Orient, G. E. Plencner, Robert M. Papila, Nilay Multidisciplinary Approach to Aerospike Overview of the Beta II Two-Stage-To- CFD-Based Design Optimization for Nozzle Design – 120 Orbit vehicle design – 172 Single Element Rocket Injector – 25 Poladian, David Osborne, Robin Paquette, T. Dynamics and Stability and Control Evaluation and Characterization Study of Ceramic Matrix Composite Cooled Characteristics of the X-37 – 62 Dual Pulse Laser-Induced Spark (DPLIS) Nozzle Material Development Pro- for Rocket Engine Ignition System Appli- gram – 21 Porter, John cation – 14 Ceramic Matrix Composite Cooled Parker, Nelson Nozzle Material Development Pro- Overbey, Glenn Natural Atmospheric Environment Model gram – 21 Natural Atmospheric Environment Model Development for the National Aeronau- Development for the National Aeronau- tics and Space Administration’s Second Poteet, C. C. tics and Space Administration’s Second Generation Reusable Launch Ve- Advanced Metallic Thermal Protection Generation Reusable Launch Ve- hicle – 57 System Development – 64 hicle – 57 Paseur, Lila Poteet, Carl C. Padula, S. L. Design and Testing of Non-Toxic RCS Hypervelocity Impact Test Results for a Multidisciplinary Optimization Branch Ex- Thrusters for Second Generation Reus- Metallic Thermal Protection Sys- perience Using iSIGHT Software – 91 able Launch Vehicle – 14 tem – 10

B-17 Preliminary Thermal-Mechanical Sizing Rasmussen, M. L. Riehl, John of Metallic TPS: Process Development An integrated aerodynamic/propulsion SRM-Assisted Trajectory for the GTX and Sensitivity Studies – 33 study for generic aero-space planes Reference Vehicle – 47 based on waverider concepts – 173 Pototzky, Anthony S. Rivers, H. Kevin A method of predicting quasi-steady Ray, Ronald J. Cyclic Cryogenic Thermal-Mechanical aerodynamics for flutter analysis of high Linear Aerospike SR-71 Experiment Testing of an X-33/RLV Liquid Oxygen speed vehicles using steady CFD calcu- (LASRE): Aerospace Propulsion Hazard Tank Concept – 94 lations – 157 Mitigation Systems – 102 Detection of Micro-Leaks Through Com- Aerothermoelastic analysis of a NASP Reed, R. Dale plex Geometries Under Mechanical Load demonstrator model – 155 Wingless Flight: The Lifting Body and at Cryogenic Temperature – 79 Potter, Seth D. Story – 105 Thermal Structures Technology Develop- A Cryogenic Propellant Production Depot Rehder, J. J. ment for Reusable Launch Vehicle Cryo- genic Propellant Tanks – 107 for Low Earth Orbit – 81 Compute as Fast as the Engineers Can Powell, R. W. Think! ULTRAFAST COMPUTING TEAM X-37 C-SiC CMC Control Surface Com- Comparison of Two Multidisciplinary Op- FINAL REPORT – 90 ponents Development: Status of the timization Strategies for Launch-Vehicle NASA/Boeing/USAF Orbital Vehicle and Reuther, James Related Efforts – 20 Design – 123 Multi-Disciplinary Analysis for Future Optimal technology investment strate- Launch Systems Using NASA’s Ad- Roback, V. Eric gies for a reusable launch vehicle – 134 vanced Engineering Environment Flush Airdata Sensing (FADS) System (AEE) – 17 Calibration Procedures and Results for Powell, Richard W. Blunt Forebodies – 92 Analysis of the staging maneuver and Reuther, J. booster glideback guidance for a two- A Geometry-Centered Approach For Robbie, Malcolm G. stage, winged, fully reusable launch ve- Two-Stage-To-Orbit Vehicle Synthe- Development and Capabilities of Unique hicle – 159 sis – 45 Structural Seal Test Rigs – 41 Powers, Bruce G. Reuthers, James Update on the Development and Capa- bilities of Unique Structural Seal Test Flight Stability and Control and Perfor- Thermal Protection Materials for Reentry Rigs – 3 mance Results from the Linear Aerospike Applications – 59 SR-71 Experiment (LASRE) – 101 Roberts, Barry C. Rezin, Marc Natural Atmospheric Environment Model Flight Testing the Linear Aerospike The Evolution of Flexible Insulation as SR-71 Experiment (LASRE) – 100 Development for the National Aeronau- Thermal Protection Systems for Reus- tics and Space Administration’s Second Prabhu, Ramadas K. able Launch Vehicles: AFRSI (Advanced Generation Reusable Launch Ve- An implementation of a chemical and Flexible Reusable Surface Insulation) to hicle – 57 thermal nonequilibrium flow solver on CRI (Conformal Reusable Insula- Roberts, J. K. unstructured meshes and application to tion) – 67 blunt bodies – 145 Composite,Cryogenic, Conformal, Com- Richards, W. Lance mon Bulkhead, Aerogel-Insulated Tank An Inviscid Computational Study of an Flight Demonstration of X-33 Vehicle (CBAT) – 74 X-33 Configuration at Hypersonic Health Management System Compo- Speeds – 94 nents on the F/A-18 Systems Research Robertson, Anthony Aircraft – 15 Rocket-Induced Magnetohydrodynamic Computational Study of a McDonnell Ejector: A Single-Stage-to-Orbit Ad- Douglas Single-Stage-to-Orbit Vehicle Richardson, Joseph vanced Propulsion Concept – 131 Concept for Aerodynamic Analy- The 1994 NASA/USRA/ADP Design sis – 126 Projects – 140 Robinson, James C. Analytical comparison of three stiffened Preston, M. L. Richter, G. P. panel concepts – 135 Compute as Fast as the Engineers Can Background, current status, and progno- Robinson, Philip J. Think! ULTRAFAST COMPUTING TEAM sis of the ongoing slush hydrogen tech- FINAL REPORT – 90 nology development program for the Non-Toxic Dual Thrust Reaction Control NASP – 177 Engine Development for On-Orbit APS Prosser, W. H. Applications – 11 Development of Structural Health Man- Ricketts, Rodney H. agement Technology for Aerospace Ve- Robles, Bryan An overview of aeroelasticity studies for Using CFD as Rocket Injector Design hicles – 1 160 the National Aerospace Plane – Tool: Recent Progress at Marshall Space Psiaki, Mark L. NASP aeroservothermoelasticity stud- Flight Center – 3 Real-time trajectory optimization on par- ies – 174 allel processors – 158 Roche, Joseph M. Structural dynamic and aeroelastic con- Affordable Flight Demonstration of the Pulley, John siderations for hypersonic ve- GTX Air-Breathing SSTO Vehicle Con- Thermal-Mechanical Cyclic Test of a hicles – 173 cept – 26 Composite Cryogenic Tank for Reusable Launch Vehicles – 23 Ridge, Jerry Performance Evaluation of the NASA Evaluation of Thermal Control Coatings GTX RBCC Flowpath – 70 Quisenberry, Bob for Flexible Ceramic Thermal Protection Preliminary Sizing of Vertical Take-Off Reusable Launch Vehicle Re- Systems – 117 Rocket-Based Combined-Cycle Pow- search – 119 Riehl, John P. ered Launch Vehicles – 85 Rahman, S. Affordable Flight Demonstration of the Structural Sizing of a 25,000-lb Payload, Engineering the Future of Full-Scale Pro- GTX Air-Breathing SSTO Vehicle Con- Air-Breathing Launch Vehicle for Single- pulsion Testing – 7 cept – 26 Stage-to-Orbit – 85

B-18 Rock, K. E. Russell, S. S. Schmidt, David K. Hyper-X Wind Tunnel Program – 109 Thermographic Analysis of Composite Dynamics of aerospace vehicles – 174 Cobonds on the X-33 – 75 Rocker, Marvin Effect of Aeropropulsive Interactions and Using CFD as Rocket Injector Design Ryan, H. M. Design Sensitivities on Optimal Hyper- sonic Ascent Trajectories – 125 Tool: Recent Progress at Marshall Space Engineering the Future of Full-Scale Pro- Flight Center – 3 pulsion Testing – 7 Schmidt, Harold Vision for CFD-Based Combustion Insta- Magnetohydrodynamic Augmented Pro- Ryan, Robert bility Predictions – 23 pulsion Experiment – 52 NASA Stennis Space Center V and V Roe, Fred Capabilities Overview – 2 Schmidt, I. H. An Advanced Video Sensor for Auto- Advanced Metallic Thermal Protection mated Docking – 61 Sackheim, Robert L. System Development – 64 Earth-to-Orbit Rocket Propulsion – 17 Rogacki, J. Schultz, James Thermoviscoplastic response of Ti-15-3 Safie, Fayssal M. Winds of change: Expanding the frontiers under various loading conditions – 172 A Quantitative Reliability, Maintainability of flight. Langley Research Center’s 75 years of accomplishment, 1917- Rogacki, John R. and Supportability Approach for NASA’s Second Generation Reusable Launch 1992 – 165 2nd Generation Reusable Launch Ve- Vehicle – 41 hicle Potential Commercial Development Schulz, Paul Scenarios – 79 Safie, Fayssal Materials Research for High Speed Civil Transport and Generic Hypersonics- Progress and Status of the 2nd Genera- Life Cycle Systems Engineering Ap- Metals Durability – 127 tion Reusable Launch Vehicle Pro- proach to NASA’s 2nd Generation Reus- gram – 76 able Launch Vehicle – 39 Schuster, Dave Evaluation of Computational Aeroelastic- Roger, R. C. Saint Cyr, W. ity Codes for Loads and Flutter – 82 Hyper-X Wind Tunnel Program – 109 Engineering the Future of Full-Scale Pro- pulsion Testing – 7 Schwabacher, Mark Rogers, James P. The NASA Integrated Vehicle Health Propulsion module design review. Mile- Sakahara, Robert Management Technology Experiment for stone 8: Integrated Propulsion Technol- The X-33 Extended Flight Test X-37 – 53 ogy Demonstrator – 150 Range – 99 Schweikhard, Keith A. Rogers, Stuart E. Salas, A. O. Flight Demonstration of X-33 Vehicle Automated CFD Parameter Studies on Managing MDO Software Development Health Management System Compo- Distributed Parallel Computers – 36 Projects – 37 nents on the F/A-18 Systems Research Aircraft – 15 Rogers, Thomas F. Multidisciplinary Approach to Aerospike General Public Space Travel and Tour- Nozzle Design – 120 Scott, Robert C. ism – 96 A method of predicting quasi-steady Multidisciplinary Optimization Branch Ex- aerodynamics for flutter analysis of high Rogersm J. L. perience Using iSIGHT Software – 91 speed vehicles using steady CFD calcu- Compute as Fast as the Engineers Can lations – 157 Think! ULTRAFAST COMPUTING TEAM Samuels, Jeff FINAL REPORT – 90 The NASA Integrated Vehicle Health Scotti, Stephen J. Management Technology Experiment for Current Technology for Thermal Protec- Rorabaugh, Michael X-37 – 53 tion Systems – 165 Advanced High Temperature Structural Seals – 43 Sandra, Hayden Scottoline, Charles NASA IVHM Technology Experiment for Reusable Launch Vehicle Re- Rose, William C. X-vehicles (NITEX) – 67 search – 119 Numerical investigations in three- Selvaduray, Guna dimensional internal flows – 171 Sankaran, S. N. Development of Processing Techniques Rosen, Bruce S. Detection of Micro-Leaks Through Com- for Advanced Thermal Protection Materi- plex Geometries Under Mechanical Load Addition of equilibrium air to an upwind als – 112 and at Cryogenic Temperature – 79 Navier-Stokes code and other first steps Setlock, J. A. toward a more generalized flow Santoro, Robert J. Oxidation-Reduction Resistance of Ad- 174 solver – An Experimental Study of Characteristic vanced Copper Alloys – 11 Rosen, Robert Combustion-Driven Flow for CFD Valida- tion – 103 Shadoan, Mike The future challenge for aeropropul- Selection and Evaluation of An Alloy for sion – 167 Sarker, Bhaba R. Nozzle Application – 22 Roth, Axel Operations Analysis of Space Shuttle Shah, Sandeep System – 50 Estimating the Cost of NASA’s Space Selection and Evaluation of An Alloy for Launch Initiative: How SLI Cost Stack Up Scheiern, M. Nozzle Application – 22 Against the Shuttle – 51 Launch Vehicle Sizing Benefits Utilizing Shanabarger, M. R. Rudy, D. H. Main Propulsion System Crossfeed and An Investigation of the Effect of Surface Compute as Fast as the Engineers Can Project Status – 34 Impurities on the Adsorption Kinetics of Think! ULTRAFAST COMPUTING TEAM Hydrogen Chemisorbed onto Iron – 101 Schmidt, D. K. FINAL REPORT – 90 A parametric sensitivity study for single- Sharma, Ashley Rusick, Jeff stage-to-orbit hypersonic vehicles using X-33 Integrated Test Facility Extended NASA Overview – 35 trajectory optimization – 153 Range Simulation – 100

B-19 Sharman, R. D. Sieja, James P. Smith, Natasha L. Lee waves: Benign and malig- Experimental Performance Evaluation of System Risk Assessment and Allocation nant – 157 a Supersonic Turbine for Rocket Engine in Conceptual Design – 27 Applications – 10 Shaughnessy, John D. Smith,O.E. Hypersonic vehicle control law develop- Sikora, J. G. A Compendium of Wind Statistics and ment using H infinity and mu- Detection of Micro-Leaks Through Com- Models for the NASA Space Shuttle and synthesis – 164 plex Geometries Under Mechanical Load Other Aerospace Vehicle Pro- Hypersonic vehicle model and control and at Cryogenic Temperature – 79 grams – 99 law development using H(infinity) and Sikora, Joseph Smith, Russell micron synthesis – 144 Development of a Pressure Box to Evalu- Thermal Structures Technology Develop- Trim drag reduction concepts for horizon- ate Reusable-Launch-Vehicle ment for Reusable Launch Vehicle Cryo- tal takeoff single-stage-to-Orbit ve- Cryogenic-Tank Panels – 132 genic Propellant Tanks – 107 hicles – 170 Singh, D. J. Smith, Timothy D. Shell, Dale Thrust Enhancement in Hypervelocity Experimental Evaluation of a Subscale Reusable Launch Vehicle Re- Nozzles by Chemical Catalysis – 102 Gaseous Hydrogen/Gaseous Oxygen search – 119 Coaxial Rocket Injector – 34 Sirocky, Paul J. Status of the RBCC Direct-Connect Shelton, Duane Hot dynamic test rig for measuring hyper- Mixer Combustor Experiment – 54 JTEC panel report on space and transat- sonic engine seal flow and durabil- mospheric propulsion technology – 175 ity – 151 Smitherman, David Shen, Zuojun Sitz, Joel R. A Cryogenic Propellant Production Depot for Low Earth Orbit – 81 On-Board Entry Trajectory Planning Ex- The F-18 systems research aircraft facil- panded to Sub-orbital Flight – 12 ity – 163 Snellgrove, Lauren M. On-Board Generation of Three- Six, F. Experimental Performance Evaluation of a Supersonic Turbine for Rocket Engine Dimensional Constrained Entry Trajecto- Research Reports: 1996 NASA/ASEE Applications – 10 ries – 35 Summer Faculty Fellowship Pro- Shorey, Mark W. gram – 102 Snyder, Christopher A. Thrust augmentation options for the Beta Advanced High Temperature Structural Skelley, Stephen Seals – 87 2 two-stage-to-orbit vehicle – 154 Water Flow Performance of a Superscale Shorey, Mark Model of the Fastrac Liquid Oxygen Sobieski, J. Advanced High Temperature Structural Pump – 55 Compute as Fast as the Engineers Can Seals – 43 Think! ULTRAFAST COMPUTING TEAM Smith, C. A. FINAL REPORT – 90 Shtessel, Y. B. Operations Analysis of the 2nd Genera- Direct Fault Tolerant RLV Altitude Con- tion Reusable Launch Vehicle – 43 Solano, Wanda M. trol: A Singular Perturbation Ap- Improved Testing Capability and Adapt- proach – 40 Smith, Charles A. ability Through the Use of Wireless Sen- A Quantitative Reliability, Maintainability sors – 2 Shtessel, Yuri B. and Supportability Approach for NASA’s Improved Re-Configurable Sliding Mode Second Generation Reusable Launch Solano, W. Controller for Reusable Launch Vehicle Vehicle – 41 Engineering the Future of Full-Scale Pro- of Second Generation Addressing Aero- pulsion Testing – 7 dynamic Surface Failures and Thrust De- Smith, Charles Souchier, Alain ficiencies – 34 Life Cycle Systems Engineering Ap- Earth-to-Orbit Rocket Propulsion – 17 On-Line Computation of a Local Attain- proach to NASA’s 2nd Generation Reus- 39 able Moment Set for Reusable Launch able Launch Vehicle – Sozen, Mehmet Vehicles – 42 Systems Engineering Approach to Tech- A Novel Approach for Modeling Chemical Reaction in Generalized Fluid System Reusable Launch Vehicle Attitude Con- nology Integration for NASA’s 2nd Gen- Simulation Program – 13 trol Using a Time-Varying Sliding Mode eration Reusable Launch Vehicle – 23 Control Technique – 40 Smith, Dane Spain, Charles V. Shuart, Mark J. Evaluation of Thermal Control Coatings Aerothermoelastic analysis of a NASP demonstrator model – 155 Advanced Durability and Damage Toler- for Flexible Ceramic Thermal Protection ance Design and Analysis Methods for Systems – 117 Supersonic aeroelastic instability results Composite Structures: Lessons Learned Materials Testing on the DC-X and DC- for a NASP-like wing model – 159 from NASA Technology Development XA – 121 Sprague, Stephanie Programs – 26 Smith, Dennis E. Wind-Tunnel Investigations of Blunt- Shyy, Wei NASA 2nd Generation RLV Program In- Body Drag Reduction Using Forebody CFD-Based Design Optimization for troduction, Status and Future Surface Roughness – 85 Single Element Rocket Injector – 25 Plans – 31 Springer, A. M. Sides, Steve Space Launch Initiative: New Capabili- Transonic aerodynamic characteristics of 2nd Generation Reusable Launch Ve- ties - New Horizons – 38 a proposed wing-body reusable launch hicle (2G RLV) – 82 vehicle concept – 138 Smith, Garrett Sidwell, D. Analysis of Parallel Burn Without Cross- Srinivasan, Vijayakumar Large Composite Structures Processing feed TSTO RLV Architectures and Com- Development of Processing Techniques Technologies for Reusable Launch Ve- parison to Parallel Burn With Crossfeed for Advanced Thermal Protection Materi- hicles – 75 and Series Burn Architectures – 19 als – 112

B-20 St. John, Clinton Third Generation RLV Structural Seal Theisen, John Propellant Feed System Leak Detection: Development Programs at NASA Flight Demonstration of X-33 Vehicle Lessons Learned From the Linear Aero- GRC – 3 Health Management System Compo- spike SR-71 Experiment (LASRE) – 93 Update on the Development and Capa- nents on the F/A-18 Systems Research Aircraft – 15 Stackpoole, M. M. bilities of Unique Structural Seal Test Fractographic Analysis of HfB2-SiC and Rigs – 3 Thomas, Dale ZrB2-SiC Composites – 57 Steinetz, Bruce Life Cycle Systems Engineering Ap- Stackpoole, Mairead Advanced High Temperature Structural proach to NASA’s 2nd Generation Reus- Thermal Protection Materials for Reentry Seals – 87 able Launch Vehicle – 39 Applications – 59 Systems Engineering Approach to Tech- Stewart, David A. nology Integration for NASA’s 2nd Gen- Stallmer, Eric W. Surface Catalysis and Characterization eration Reusable Launch Vehicle – 23 General Public Space Travel and Tour- of Proposed Candidate TPS for Access- ism – 96 to-Space Vehicles – 118 Thomas, David J. Stanic, Vesna Surface Catalytic Efficiency of Advanced Ceramic Matrix Composites (CMC) Life 88 Durability of Membrane Electrode As- Carbon Carbon Candidate Thermal Pro- Prediction Method Development – semblies (MEAs) in PEM Fuel Cells Op- tection Materials for SSTO Ve- Design and Analysis of UHTC Leading erated on Pure Hydrogen and Oxy- hicles – 133 Edge Attachment – 44 gen – 6 Thermal Protection System Evaluation Thomas, Leann Using Arc-jet Flows: Flight Simulation or Stanley, D. O. Systems Engineering Approach to Tech- Research Tool? – 31 Comparison of Two Multidisciplinary Op- nology Integration for NASA’s 2nd Gen- timization Strategies for Launch-Vehicle Stockwell, A. E. eration Reusable Launch Vehicle – 23 Design – 123 A study of facilities and fixtures for testing Thomas, Scott R. Stanley, Douglas O. of a high speed civil transport wing com- Performance Evaluation of the NASA ponent – 127 Dual-Fuel Propulsion in Single-Stage Ad- GTX RBCC Flowpath – 70 vanced Manned Launch System Ve- Stone, J. E. hicle – 125 Status of the RBCC Direct-Connect Heat pipes for wing leading edges of Mixer Combustor Experiment – 54 Stanley, Thomas Troy hypersonic vehicles – 178 A Collaborative Analysis Tool for Thermal Thomas-Ogbuji, L. Protection Systems for Single Stage to Storaasli, O. O. Oxidation-Reduction Resistance of Ad- Orbit Launch Vehicles – 56 Compute as Fast as the Engineers Can vanced Copper Alloys – 11 Think! ULTRAFAST COMPUTING TEAM Starnes, James H., Jr. FINAL REPORT – 90 Thomas-Ogbuji, Linus U. Advanced Durability and Damage Toler- Ceramic Matrix Composites (CMC) Life ance Design and Analysis Methods for Strauss, B. Prediction Method Development – 88 Composite Structures: Lessons Learned Thermal Management Design for the from NASA Technology Development X-33 Lifting Body – 104 Thomason, Lee B. Programs – 26 Supersonic aeroelastic instability results Sukanek, Peter C. for a NASP-like wing model – 159 Staszak, Paul NASA EPSCoR Preparation Grant – 49 Task 4 supporting technology. Part 1: Thompson, Richard A. Detailed test plan for leading edge tile Sullivan, Greg Hypersonic Boundary-Layer Transition development. Leading edge material de- Reusable Launch Vehicle Re- for X-33 Phase 2 Vehicle – 111 velopment and testing – 139 search – 119 X-33 Computational Steffen, C. J., Jr. Sullivan, Robert T. Aeroheating/Aerodynamic Predictions Inlet Development for a Rocket Based Lockheed Martin Response to the OSP and Comparisons With Experimental Combined Cycle, Single Stage to Orbit Challenge – 20 Data – 25 Vehicle Using Computational Fluid Dy- Sutherlin, Steve Thompson, Richard namics – 95 A Cryogenic Propellant Production Depot Statistical Methods for Rapid Aerother- Stein, Matthew D. for Low Earth Orbit – 81 mal Analysis and Design Technol- Configuration development study of the ogy – 55 Sutton, Robert OSU 1 hypersonic research ve- Throckmorton, David A. hicle – 154 Experimental Evaluation of a Subscale Gaseous Hydrogen/Gaseous Oxygen Second Generation RLV Space Vehicle Steinetz, Bruce M. Coaxial Rocket Injector – 34 Concept – 65 2000 NASA Seal/Secondary Air System Talay, Theodore A. Tietjen, Alan Workshop – 62 Reusable Launch Vehicle Technology Infrared Sensing Aeroheating Flight Ex- Advanced Control Surface Seal Develop- Program – 127 periement: STS-96 Flight Results – 84 ment at NASA GRC for 3rd Generation RLV – 29 Single-stage-to-orbit: Meeting the chal- Infrared Sensing Aeroheating Flight Ex- lenge – 136 periment: STS-96 Flight Results – 86 Development and Capabilities of Unique Structural Seal Test Rigs – 41 Taylor, J. L. Tiller, Bruce K. High temperature NASP engine seals: A Integrated Technology Assessment Cen- Propellant Densification for Shuttle: The 42 32 technology review – 173 ter (ITAC) Update – SSME Perspective – Hot dynamic test rig for measuring hyper- Tejnil, Edward Tollefson, J. C. sonic engine seal flow and durabil- Automated CFD Parameter Studies on Personnel launch system autoland de- ity – 151 Distributed Parallel Computers – 36 velopment study – 175

B-21 Tomsik, T. M. Trinh, Huu Vaidyanathan, Rajkumar Background, current status, and progno- Evaluation and Characterization Study of CFD-Based Design Optimization for sis of the ongoing slush hydrogen tech- Dual Pulse Laser-Induced Spark (DPLIS) Single Element Rocket Injector – 25 nology development program for the for Rocket Engine Ignition System Appli- Valentine, Peter G. NASP – 177 cation – 14 X-37 C-SiC CMC Control Surface Com- Tomsik, Thomas M. Tryon, Robert ponents Development: Status of the A summary of the slush hydrogen tech- The 1994 NASA/USRA/ADP Design NASA/Boeing/USAF Orbital Vehicle and nology program for the National Aero- Projects – 140 Related Efforts – 20 Space Plane – 140 Tseng, Kevin Vause, R. F. Performance Tests of a Liquid Hydrogen Impedance-Based Structural Health A study of facilities and fixtures for testing Propellant Densification Ground System Monitoring for Composite Laminates at of a high speed civil transport wing com- 118 for the X33/RLV – Cryogenic Environments – 7 ponent – 127 Prediction of the ullage gas thermal Veazie, David R. Tu, Tina stratification in a NASP vehicle propellant Characterization of Polyimide Foams for Test Plans. Lightweight Durable TPS: tank experimental simulation using Ultra-Lightweight Space Struc- Tasks 1,2,4,5, and 6 – 146 FLOW-3D – 177 tures – 22 Recent Advances and Applications in Tucker, Kevin Veith, Eric M. Cryogenic Propellant Densification Tech- CFD-Based Design Optimization for Non-Toxic Dual Thrust Reaction Control nology – 89 Single Element Rocket Injector – 25 Engine Development for On-Orbit APS Testing of Densified Liquid Hydrogen Using CFD as Rocket Injector Design Applications – 11 Stratification in a Scale Model Propellant Tool: Recent Progress at Marshall Space Venkatapathy, Ethiras Tank – 67 Flight Center – 3 Thermal Protection System Evaluation Tong, Mike Turner, James E. Using Arc-jet Flows: Flight Simulation or High temperature NASP engine seals: A Analysis of X-33 Linear Aerospike Plume Research Tool? – 31 technology review – 173 Induced Base-Heating Environment Dur- Verilli, Mike Townsend, J. C. ing Power-Pack Out – 29 Characterization of C/SiC Ceramic Ma- Managing MDO Software Development Lightweight Chambers for Thrust Assem- trix Composites (CMCs) with Novel Inter- Projects – 37 blies – 82 face Fiber Coatings – 48 Traci, Richard M. Unic Unstructured CFD Methodology De- Verrilli, Michael J. A Thermal Management Systems Model velopment – 82 Ceramic Matrix Composites (CMC) Life Prediction Method Development – 88 for the NASA GTX RBCC Concept – 53 Unshrouded Impeller Technology Task Tramel, Terri Status – 81 Verzemnieks, Juris Advanced High Temperature Structural Design and Testing of Non-Toxic RCS Turner, Jim E. Thrusters for Second Generation Reus- Seals – 43 X-33 XRS-2200 Linear Aerospike Engine able Launch Vehicle – 14 Sea Level Plume Radiation – 79 Vesperman, Curtis L. Tran, Duoc Propulsion module design review. Mile- Evaluation of Thermal Control Coatings Turner, Jim stone 8: Integrated Propulsion Technol- for Flexible Ceramic Thermal Protection Lightweight Chambers for Thrust Assem- ogy Demonstrator – 150 blies – 66 Systems – 117 Vestal, L. Tran, Hoi X-33 XRS-2200 Linear Aerospike Engine Tether Transportation System Dynamics and Stability and Control Sea Level Plume Radiation – 66 Study – 109 Characteristics of the X-37 – 62 Turner, Larry D. Vickers, J. H. Trefny, C. J. Thermal/Fluid Analysis of a Composite Large Composite Structures Processing Analysis of a New Rocket-Based Heat Exchanger for Use on the RLV Technologies for Reusable Launch Ve- Combined-Cycle Engine Concept at Low Rocket Engine – 63 hicles – 75 Speed – 94 Turner, Susan G. Vilja, John Inlet Development for a Rocket Based Dynamics and Stability and Control Main Engine Prototype Development for 43 Combined Cycle, Single Stage to Orbit Characteristics of the X-37 – 62 2nd Generation RLV RS-83 – Vehicle Using Computational Fluid Dy- Voelker, Leonard S. namics – 95 Orbital Space Plane Program Flight Demonstrators Status – 6 Flight Testing the Linear Aerospike Trefny, Charles J. SR-71 Experiment (LASRE) – 100 Computational analysis in support of the Turner, Susan Voland, R. T. SSTO flowpath test – 143 Flight Demonstrations of Orbital Space Hyper-X Wind Tunnel Program – 109 Plane (OSP) Technologies – 16 Performance Evaluation of the NASA Volk, John GTX RBCC Flowpath – 70 Test Results for an Off-the-Shelf Evaluation of Computational Aeroelastic- Trefny, Charles GPS/INS During Approach and Landing ity Codes for Loads and Flutter – 82 Testing of the X-40A – 63 SRM-Assisted Trajectory for the GTX Vu, B. Reference Vehicle – 47 Tuttle, M. E. Computational Aerodynamics Analysis of Trent, Donnie Thermoviscoplastic response of Ti-15-3 Future Launch Vehicle Configura- Operationally efficient propulsion system under various loading conditions – 172 tions – 122 study (OEPSS) data book – 110 Unal, Resit Wadel, Mary F. Operationally Efficient Propulsion Sys- Dual-Fuel Propulsion in Single-Stage Ad- Comparison of High Aspect Ratio Cool- tem Study (OEPSS): OEPSS Video vanced Manned Launch System Ve- ing Channel Designs for a Rocket Com- Script – 110 hicle – 125 bustion Chamber – 118

B-22 Wagner, Alain Watters, D. G. Whalen, M. V. Systems Challenges for Hypersonic Ve- Health Monitoring Technology for Ther- Background, current status, and progno- hicles – 117 mal Protection Systems on Reusable Hy- sis of the ongoing slush hydrogen tech- personic Vehicles – 71 nology development program for the Wagner, John A. NASP – 177 Fatigue Crack Growth Rate Test Results Watters, David G. for Al-Li 2195 Parent Metal, Variable Po- Wireless Subsurface Microsensors for Whalen, Margaret V. larity Plasma Arc Welds and Friction Stir Health Monitoring of Thermal Protection A summary of the slush hydrogen tech- Welds – 89 Systems on Hypersonic Vehicles – 69 nology program for the National Aero- Space Plane – 140 Walberg, Gerald D. Watts, Eric T. Slush hydrogen propellant production, Multidisciplinary design techniques ap- Results of Two-Stage Light-Gas Gun De- transfer, and expulsion studies at the plied to conceptual aerospace vehicle velopment Efforts and Hypervelocity Im- NASA K-Site Facility – 173 design – 156 pact Tests of Advanced Thermal Protec- tion Materials – 100 Slush hydrogen transfer studies at the Waldrop, Glen S. NASA K-Site Test Facility – 167 Operationally Efficient Propulsion Sys- Webber, Derek Technology issues associated with using tem Study (OEPSS) data book. Volume Analysis of space concepts enabled by densified hydrogen for space ve- 2: Ground operations problems – 170 new transportation (ASCENT) hicles – 167 study – 28 Operationally Efficient Propulsion Sys- Whitmore, Stephen A. tem Study (OEPSS): OEPSS Video Weber, T. F., Jr. A Base Drag Reduction Experiment on Script – 110 Reusable launch vehicle reliability, main- the X-33 Linear Aerospike SR-71 Experi- Walker,J.L. tainability and operability assessment (LASRE) Flight Program – 97 150 Thermographic Analysis of Composite ment – Design and Calibration of the X-33 Flush Cobonds on the X-33 – 75 Weeks, D. P. Airdata Sensing (FADS) System – 113 Walker, James F. Task 4 supporting technology. Densifica- Flush Airdata Sensing (FADS) System Status of the RBCC Direct-Connect tion requirements definition and test ob- Calibration Procedures and Results for Mixer Combustor Experiment – 54 jectives. Propellant densification require- Blunt Forebodies – 92 ments definition – 139 Wind-Tunnel Investigations of Blunt- Walker, James Wehrmeyer, Joseph Body Drag Reduction Using Forebody A Thermal Management Systems Model Surface Roughness – 85 for the NASA GTX RBCC Concept – 53 Evaluation and Characterization Study of Dual Pulse Laser-Induced Spark (DPLIS) Wilkerson, C. Study Acoustic Emissions from Compos- for Rocket Engine Ignition System Appli- Acoustic Emission Monitoring of the ites – 111 cation – 14 DC-XA Composite Liquid Hydrogen Tank During Structural Testing – 126 Wang, T. S. Weilmuenster, K. James Unic Unstructured CFD Methodology De- Computational/Experimental Aeroheat- Williams, Peggy S. velopment – 82 ing Predictions for X-33 Phase 2 Ve- A Monte Carlo Dispersion Analysis of the Wang, Tee-See hicle – 110 X-33 Simulation Software – 72 Base-Bleed Effect on X-33 Aerospike Computational/Experimental Aeroheat- Williams, Robert W. Plume Induced Base-Heating Environ- ing Predictions for X-33 – 112 Unshrouded Impeller Technology Task ment During Power-Pack Out – 25 Weiser, Erik S. Status – 81 Wang, Ten-See Cryopumping in Cryogenic Insulations Williams, Robert Analysis of Aerospike Plume Induced for a Reusable Launch Vehicle – 2 Analysis of X-33 Linear Aerospike Plume Base-Heating Environment – 109 Induced Base-Heating Environment Dur- Welch, Sharon S. ing Power-Pack Out – 29 Analysis of X-33 Linear Aerospike Plume Technology Roadmap for Dual-Mode Induced Base-Heating Environment Dur- Scramjet Propulsion to Support Space- Base-Bleed Effect on X-33 Aerospike ing Power-Pack Out – 29 Access Vision Vehicle Develop- Plume Induced Base-Heating Environ- Overview of Fluid Dynamic Activities at ment – 38 ment During Power-Pack Out – 25 the Marshall Space Flight Center – 56 Wendt, Jodean Overview of MSFC’s Applied Fluid Dy- namics Analysis Group Activities – 24 Stage Separation CFD Tool Develop- Infrared Sensing Aeroheating Flight Ex- ment and Evaluation – 38 periement: STS-96 Flight Results – 84 Progress in Unsteady Turbopump Flow Simulations – 51 X-33 XRS-2200 Linear Aerospike Engine Infrared Sensing Aeroheating Flight Ex- Sea Level Plume Radiation – 66 periment: STS-96 Flight Results – 86 Time Dependent Simulation of Tur- bopump Flows – 55 Ware, George M. Wercinski, Paul Subsonic Aerodynamic Characteristics of Time-Dependent Simulations of Tur- Thermal Protection Materials for Reentry bopump Flows – 38 a Circular Body Earth-to-Orbit Ve- Applications – 59 hicle – 128 Using CFD as Rocket Injector Design Subsonic aerodynamic characteristics of West, Jeff Tool: Recent Progress at Marshall Space a proposed advanced manned launch Using CFD as Rocket Injector Design Flight Center – 3 system orbiter configuration – 160 Tool: Recent Progress at Marshall Space Wilmoth, Richard G. Flight Center – 3 Waters, W. A., Jr. Infrared Sensing Aeroheating Flight Ex- periement: STS-96 Flight Results – 84 A study of facilities and fixtures for testing West, Jeffrey S. of a high speed civil transport wing com- Vision for CFD-Based Combustion Insta- Infrared Sensing Aeroheating Flight Ex- ponent – 127 bility Predictions – 23 periment: STS-96 Flight Results – 86

B-23 Windhorst, Robert Wu,Q.H. Zeiler, Thomas A. Minimum Heating Re-Entry Trajectories A Hazardous Gas Detection System for Aerothermoelastic analysis of a NASP for Advanced Hypersonic Launch Ve- Aerospace and Commercial Applica- demonstrator model – 155 98 hicles – 119 tions – Zhang, S. Winn, Peter M. Wu, Zhenglu Unic Unstructured CFD Methodology De- Reconfigurable Flight Control Designs Development of a Pressure Box to Evalu- velopment – 82 With Application to the X-33 Ve- ate Reusable-Launch-Vehicle Zhu, J. J. hicle – 95 Cryogenic-Tank Panels – 132 Autocommander: A Supervisory Control- Wu, Zheng-Lu ler for Integrated Guidance and Control Wong, George S. Two Reconfigurable Flight-Control De- for the 2nd Generation Reusable Launch Operationally Efficient Propulsion Sys- sign Methods: Robust Servomechanism Vehicle – 40 tem Study (OEPSS): OEPSS Video and Control Allocation – 76 Script – 110 Direct Fault Tolerant RLV Altitude Con- Wurtele, M. G. trol: A Singular Perturbation Ap- Woodcock, Gordon Lee waves: Benign and malig- proach – 40 Highly Reusable Space Transportation nant – 157 Zhu, J. Jim (HRST) Baseline Concepts and Analysis: Yarrington, Phillip Integrated G and C Implementation Rocket/RBCC Options – 114 Design and Analysis of a Two-Stage-to- within IDOS: A Simulink Based Reusable Orbit Airbreathing Hypersonic Vehicle Launch Vehicle Simulation – 11 Woods, W. C. Concept – 129 Aerodynamic Characteristics of a Vertical Reusable Launch Vehicle Attitude Con- Takeoff Vertical Landing (VTVL) Single Yoder, Dennis trol Using a Time-Varying Sliding Mode Stage to Orbit Vehicle from M(sub infin- Computations of Internal and External Control Technique – 40 ity) = 0.1 to 10 – 122 Axisymmetric Nozzle Aerodynamics at Zimmerman, Curtis Transonic Speeds – 4 An Automated Method to Compute Or- Workman,Gary Yoon, Seokkwan bital Re-entry Trajectories with Heating Study Acoustic Emissions from Compos- Calculation of Supersonic Combustion Constraints – 39 ites – 111 Using Implicit Schemes – 30 An Automated Method to Compute Or- Wright, Richard J. Youssef, Hussein bital Re-Entry Trajectories with Heating Composite Development and Applica- Predictor-Corrector Entry Guidance for Constraints – 53 tions for RLV Tankage – 16 Reusable Launch Vehicles – 74 Predictor-Corrector Entry Guidance for Yungster, S. Reusable Launch Vehicles – 74 Wu,G.D. Analysis of a New Rocket-Based Zimmerman, F. Optimal trajectories for an aerospace Combined-Cycle Engine Concept at Low Lightweight Chambers for Thrust Assem- plane. Part 1: Formulation, results, and Speed – 94 blies – 66 analysis – 176 Zang, Thomas A. Zimmerman, Frank Optimal trajectories for an aerospace System Risk Assessment and Allocation Lightweight Chambers for Thrust Assem- plane. Part 2: Data, tables, and in Conceptual Design – 27 blies – 82 graphs – 176 Zee, Warner Zoladz, Thomas Wu, K. Chauncey Configuration development study of the Water Flow Performance of a Superscale Analytical comparison of three stiffened OSU 1 hypersonic research ve- Model of the Fastrac Liquid Oxygen panel concepts – 135 hicle – 154 Pump – 55

B-24 Report Documentation Page

1. Report No. 2. Government Accession No. 3. Recipient’s Catalog No.

4. Title and Subtitle 5. Report Date January 2004 Next Generation Spacecraft, Crew Exploration Vehicle 6. Performing Organization Code

7. Author(s) 8. Performing Organization Report No.

10. Work Unit No. 9. Performing Organization Name and Address NASA Scientific and Technical Information Program Office 11. Contract or Grant No.

12. Sponsoring Agency Name and Address 13. Type of Report and Period Covered National Aeronautics and Space Administration Langley Research Center 14. Sponsoring Agency Code Hampton, VA 23681 15. Supplementary Notes

16. Abstract

17. Key Words (Suggested by Author(s)) 18. Distribution Statement Aerospace Planes Bibliographies Unclassified – Unlimited Reusable Launch Vehicles Hypersonic Vehicles Subject Category – 15 Space Shuttles Launch Vehicle Configurations 19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of Pages 22. Price Unclassified Unclassified

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