DOCSLIB.ORG

Hydrogedoxy Gen Auxiliary Propulsion Technology

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Hydrogedoxy Gen Auxiliary Propulsion Technology NASA Technica Memorandum 05249 AIAA-91-3440 L HydrogedOxygen Auxiliary Propulsion Technology Brian D. Reed and Steven J. Schneider Lewis Research Center Cleveland, Ohio Prepared for the Conference on Advanced Space Exploration Initiative Technologies cosponsored by AIAA, NASA, and OAI Cleveland, Ohio, September 4-6, 1991 NASA HYDROGEN/OXYGEN AUXILIARY PROPULSION TECHNOLOGY Brian D. Reed and Steven J. Schneider National Aeronautics and Space Administration Lewis Research Center Cleveland, Ohio 44135 Abstract use hydrogen, oxygen, and water could lead to a simpler and more operationally efficient vehicle. This paper provides a survey of hydrogen-oxygen (H/O) auxiliary propulsion system (APS) concepts and In the past, the potential benefits of H/O propellants low-thrust H/O rocket technology. A review of H/O have made them attractive candidates for APS’s, includ- APS studies performed for the Space Shuttle, Space Tug, ing Space Shuttle and Space Station Freedom. However, Space Station Freedom, and Advanced Manned Launch they have yet to be implemented on an actual flight System programs is given. The survey also includes a system due, primarily, to concerns about the feasibility review of low-thrust H/O rocket technology programs, of developing the technology to implement the concepts. covering liquid H/O and gaseous H/O thrusters, ranging Various APS concepts using liquid, gaseous, and super- from 6600 N (1500 lbf) to 440 mN (0.1 lbf) thrust. critical H/O have been proposed over the years. Prob- Ignition concepts for H/O thrusters and high- lems, though, are usually encountered in the design of temperature, oxidation-resistant chamber materials are various concepts, such as large volume and high-pressure also reviewed. storage requirements for gaseous H/O, long-term cryo- genic storage and distribution, complexity and perfor- Introduction mance losses of propellant conditioning equipment, and uncertainties about thruster technology. Auxiliary propulsion is required on every launch vehicle, satellite, and spacecraft. Examples of auxiliary The primary propellant choice of APS’s has been propulsion include attitude control and orientation, monopropellant hydrazine or the bipropellant combina- stationkeeping, apogee insertion, rendezvous, docking, tion of monomethylhydrazine (MMH) and nitrogen separation, midcourse correction, and planetary retro. tetroxide (NTO). These Earth storable propellants have Auxiliary propulsion maneuvers can range from milli- the advantages of a long, successful flight history, long- second pulse trains to long, steady-state burns. Depend- term storability with minimal thermal control, high bulk ing on the application, thrust levels for auxiliary density, and either catalytic (monopropellant) or propulsion can range from 440 mN (0.1 lbf) to 27 kN hypergolic (bipropellant) ignition. Earth storable (6000 lbf). Thrusters in an auxiliary propulsion system propellants are also low performing compared with H/O, (APS) are usually located throughout the vehicle, are highly toxic, and have limited capability for integra- requiring a distribution network to supply propellants tion with other subsystems. The high reliability of Earth from the tanks. An APS must be able to provide storable systems has outweighed the potential benefits of frequent and rapid restarts, must have high cyclic life, H/O, because of the uncertainties associated with H/O and must have the flexibility to operate over a wide systems. However, the need for high performance in , range of environmental conditions, often with long quies- STV missions and an increased awareness of the impor- cent periods. tance of efficient servicing, checkout, and maintenance, have called attention to H/O APS. Currently, the primary candidate propellant com- bination for the APS of the space transfer vehicle (STV) This paper surveys H/O APS concepts and the H/O is hydrogen-oxygen (H/O) .‘I2 These propellants are thruster technology required to enable them. Included in attractive because of their high performance and their this review are concepts and technologies generated from nontoxic, noncorrosive nature and because of their H/O development programs for the Space Shuttle, the compatiblity with other subsystems, including the main proposed Space Tug, Space Station Freedom, and propulsion, power generation, and environmental control proposed advanced manned Earth-to-orbit vehicles. A and life support systems. Integrating subsystems that matrix matching references with H/O APS concepts is 1 given in Table I. A matrix of references and subjects in H/O thruster technology is given in Table 11. TABLE I.-REFERENCES FOR H/O APS CONCEPTS ~~ ~ Refer- Liquid Liquid stor- Gasous Super-critical ence storage age and gas- storage storage and electro- and feed eous feed and feed feed 3 X Shuttle 4 X Shuttle 5 Shuttle 7 Shuttle 8 Shuttle 9 X Shuttle 10 Space Tug 12 X Space Station 13 X Space Station 15 Space Station 17 X AMLS 32 33 Satellite 34 Satellite TABLE 11.-REFERENCES FOR H/O THRUSTER TECHNOLGY fa) Thruster Refer- Liquid hydrc en/oxygen Gaseous ydrogen/c Yeen ence 5500 N 110 N 5500 N 110 N < 22 N (1250 Ibf) (25 Ibf) (1500 Ibf) (25 Ibf) (5 Ibf) 3 X 6 X 11 X 18 X 19 X 20 X 21 X 22 X 23 X 24 X 25 X 26 X 27 X 28 X 29 X 30 X 32 X 33 X 34 X 35 X 36 X 2 TABLE 11.-CONCLUDED. [b) Ignition concepts and chamber materials Hydrogen-Oxygen Auxiliary Propulsion distribution system (see Fig. 1) using expansion bellows System Concepts accumulators and vacuum jacketed lines was proposed. 5 The pumps in this system would be located near the Space Shuttle storage tanks to avoid problems with cavitation caused by pressure drop and vaporization in the lines. Vacuum There was an extensive program’’‘ in the late 1960’s jacketed insulated lines would offer dual containment of and early 1970’s to develop H/O APS for the Space propellants. The expansion bellows accumulators would Shuttle. At the time, the Shuttle was envisioned to be accommodate the expansion of propellant due to heat a larger vehicle than the current version. The main soak. Recirculation fans in the distribution manifold propulsion tankage was internal to the orbiter vehicle. would prevent localized hot spots from vaporizing the The orbiter and the booster were to be flyback and reus- propellant. It was calculated5 that the oxygen manifold able. The Shuttle APS was to be composed of a liquid could accept 67.4 kJ/kg (29 Btu/lb) of heat before hydrogen and liquid oxygen orbital maneuvering system boiling of the propellant begins, while the hydrogen (OMS) for the orbiter and an H/O attitude control manifold could accept 36.3 kJ/kg (156 Btu/lb) of heat. system (later called the reaction control system (RCS)) This would allow for a seven-day Shuttle mission with for the orbiter and booster vehicles. little or no boiloff in the distribution manifold. A cryogenic H/O RCS was considered, but there were Pulsing of cryogenic propellants was also a technol- concerns about excessive heat leaks into the cryogenic ogy concern for a liquid H/O RCS.4 Heat soakback into distribution system causing two-phase (and therefore the injector manifold and feed system could lead to two- uncontrollable) flow to the thruster^.^ A cryogenic phase propellant flows. There were also concerns about 3 reliable low-temperature ignition. However, operation of An RCS using gaseous hydrogen and liquid oxygen liquid H/O thrusters in the pulse mode was demonstra- feed was proposed as a possible compromise between the ted,6 with successful ignition of liquid hydrogen and cryogenic and the liquid storage and gaseous feed sys- oxygen (as well as gaseous and two-phase) propellants tem~.~This concept eliminated the propellant con- with minimal amount of soakback to the feed system. ditioning equipment on the oxygen side, while avoiding Demonstration of pulsing liquid H/O thrusters, however, the distribution of liquid hydrogen, which is a deep came too late to impact the Shuttle APS program. cryogen. Ignition requirements with gaseous hydrogen and liquid oxygen were thought to be similar to the The primary candidates for Shuttle RCS were Pratt & Whitney RL-10 engine. systems with liquid storage and gaseous feed of propel- lant~.~~~After a preliminary screening, two gaseous feed Vehicle studies concluded that a smaller orbiter RCS concepts3 (one using high-chamber-pressure engines, vehicle with external, expendable main engine tankage the other using low-chamber-pressure engines) were would provide a more cost effective launch system. 4 examined in detail by McDonnell Douglas Astronautics This resulted in a reduction in shuttle size along with a Corp. and TRW Systems Group, Inc. reduction in RCS impulse requirements. An analysis showed that the downsized vehicle did not have sufficient In the high-chamber-pressure RCS concept^,^*^*^ the volume for the H/O OMS, and development of H/O propellants would be stored as liquids in low-pressure, technology was not considered justifiable without insulated tanks. Turbopumps, driven by gas generators, application to both OMS and RCS. A program decision would be used to raise propellant pressures. Heat was then made to design the APS for MMH and NTO exchangers, also driven by gas generators, would condi- propellants. The H/O technology development contin- tion the propellant. Different configurations using the ued in a limited fashion to provide a backup for the gas generators in series and in parallel with the turbo- MMH-NTO system. pumps and heat exchangers were investigated, as was the use of multiple gas generators. The propellants would be Space Tug stored in gaseous accumulators, to be fed to the thrusters at a regulated pressure. The accumulators would also In the 19'10'~~the Space Tug vehicle was proposed to supply propellants to the gas generators. Thrusters carry payloads from the Shuttle payload bay in low would operate at 2.07 to 3.45 MPa (SO0 to 500 psia) Earth orbit to high Earth orbit.
Load more

Recommended publications
  • Microgravity and Macromolecular Crystallography Craig E
    CRYSTAL GROWTH & DESIGN 2001 VOL. 1, NO. 1 87-99 Review Microgravity and Macromolecular Crystallography Craig E. Kundrot,* Russell A. Judge, Marc L. Pusey, and Edward H. Snell Mail Code SD48 Biotechnology Science Group, NASA Marshall Space Flight Center, Huntsville, Alabama 35812 Received August 24, 2000 ABSTRACT: Macromolecular crystal growth is seen as an ideal experiment to make use of the reduced acceleration environment provided by an orbiting spacecraft. The experiments are small, are simply operated, and have a high potential scientific and economic impact. In this review we examine the theoretical reasons why microgravity is a beneficial environment for crystal growth and survey the history of experiments on the Space Shuttle Orbiter, on unmanned spacecraft, and on the Mir space station. The results of microgravity crystal growth are considerable when one realizes that the comparisons are always between few microgravity-based experiments and a large number of earth-based experiments. Finally, we outline the direction for optimizing the future use of orbiting platforms. 1. Introduction molecules, including viruses, proteins, DNA, RNA, and complexes of those molecules. In this review, the terms Macromolecular crystallography is a multidisciplinary protein or macromolecule are used to refer to this entire science involving the crystallization of a macromolecule range. or complex of macromolecules, followed by X-ray or The reduced acceleration environment of an orbiting neutron diffraction to determine the three-dimensional spacecraft has been posited as an ideal environment for structure. The structure provides a basis for under- biological crystal growth, since buoyancy-driven convec- standing function and enables the development of new tion and sedimentation are greatly reduced.
    [Show full text]
  • Summary of Liquid Propulsion System Needs in Support of The
    · m~-W Summary ofLiquid Propulsion System Needs in Support ofthe Constellation Program Terry Lorier, Program Manager, Space Propulsion Systems Development Pratt & Whitney Rocketdyne Canoga Park, CA Phil Sumrall, AdvancedPlanning Manager, Ares Projects Office NASA Marshall Space Flight Center Huntsville. AL Michael Baine, Ph.D., Orion Propulsion System Manager NASA Johnson Space Center Houston, TX Abstract In January 2004, the President ofthe United States established the Vision for Space Exploration (VSE) to complete the International Space Station, retire the Space Shuttle and develop its replacement, and expand the human presence on the Moon as a stepping stone to human exploration ofMars and worlds beyond. In response, NASA developed the Constellation Program, consisting ofthe components shown in Figure 1. This paper will summarize the manned spaceflight liquid propulsion system needs in support ofthe Constellation Program over the next 10 years. It will address all liquid engine needs to support human exploration from low Earth orbit (LEO) to the lunar surface, including an overview ofengines currently under contract, those baselined but not yet under contract, and those propulsion needs that have yet to be initiated. There may be additional engine needs for early demonstrators, but those will not be addressed as part ofthis paper. Also, other portions ofthe VSE architecture, including the planned Orion abort test boosters and the Lunar Precursor Robotic Program, are not addressed here as they either use solid motors or are focused on unmanned elements ofreturning humans to the Moon. Our Exploration Fleet What will the vehicles look like? Orion Crew Exploration Vehicle Figure 1. Constellation Program components I.
    [Show full text]
  • Failures in Spacecraft Systems: an Analysis from The
    FAILURES IN SPACECRAFT SYSTEMS: AN ANALYSIS FROM THE PERSPECTIVE OF DECISION MAKING A Thesis Submitted to the Faculty of Purdue University by Vikranth R. Kattakuri In Partial Fulfillment of the Requirements for the Degree of Master of Science in Mechanical Engineering August 2019 Purdue University West Lafayette, Indiana ii THE PURDUE UNIVERSITY GRADUATE SCHOOL STATEMENT OF THESIS APPROVAL Dr. Jitesh H. Panchal, Chair School of Mechanical Engineering Dr. Ilias Bilionis School of Mechanical Engineering Dr. William Crossley School of Aeronautics and Astronautics Approved by: Dr. Jay P. Gore Associate Head of Graduate Studies iii ACKNOWLEDGMENTS I am extremely grateful to my advisor Prof. Jitesh Panchal for his patient guidance throughout the two years of my studies. I am indebted to him for considering me to be a part of his research group and for providing this opportunity to work in the fields of systems engineering and mechanical design for a period of 2 years. Being a research and teaching assistant under him had been a rewarding experience. Without his valuable insights, this work would not only have been possible, but also inconceivable. I would like to thank my co-advisor Prof. Ilias Bilionis for his valuable inputs, timely guidance and extremely engaging research meetings. I thank my committee member, Prof. William Crossley for his interest in my work. I had a great opportunity to attend all three courses taught by my committee members and they are the best among all the courses I had at Purdue. I would like to thank my mentors Dr. Jagannath Raju of Systemantics India Pri- vate Limited and Prof.
    [Show full text]
  • Study of a Nozzle Vector Control for a Low Cost Mini-Launcher
    Study of a nozzle vector control for a low cost mini-launcher Roberto Rodr´ıguez SUPERVISED BY Joshua Tristancho Universitat Politecnica` de Catalunya Master in Aerospace Science & Technology July 2011 Study of a nozzle vector control for a low cost mini-launcher BY Roberto Rodr´ıguez DIPLOMA THESIS FOR DEGREE Master in Aerospace Science and Technology AT Universitat Politecnica` de Catalunya SUPERVISED BY: Joshua Tristancho Applied Physics department ABSTRACT This study is addressed to the control system of a mini-launcher. A reaction control based on purge pressure from a solid propellant engine of a main stage, or using cold gas from a storage tank will be the control method. Pulse modulator system will be the actuation method studied, and simulations about its suitability will be done with Simulink. In this thesis, we try to check if this method is able to be implemented in a low cost way the Wiki-Launcher, a mini-launcher less than 100 kg. At the time we try to check if the control allowed by the reaction control can manage all the phases of the flight. Both possible con- trol configurations are going to be analyzed. Keywords: Nozzle Vector Control, Low cost mini-launcher, Solid propellant Acknowledgements I’m grateful to my parents and my friends for supporting me during this work. I want thank in a special way to Joshua Tristancho for his continuous help and support during this research work and for give me the opportunity to participate within his team in this project. I’m very grateful to all the WikiSat members who helped and supporting me with the re- alization of the thesis, specially Joshua Tristancho, Victor Kravchenco, Esteve Bardolet, Sonia Perez,´ Lara Navarro and Raquel Gonzalez.´ I want to thank the EETAC school for the facilities and tools they have provided to support this work.
    [Show full text]
  • Development of a Fast-Response Variable-Amplitude Programmable Reaction Control System
    NISTIR 5118 Development of a Fast-Response Variable-Amplitude Programmable Reaction Control System Building and Fire Research Laboratory Gaithersburg, Maryland 20899 NIST United States Department of Commerce Technology Administration QC National Institute of Standards and Technology 100 .056 //5118 1993 /' 'i \ NISTIR 5118 r. Development of a Fast-Response Variable-Amplitude Programmable Reaction Control System William C. Stone January 1993 Building and Fire Research Laboratory National Institute of Standards and Technology Gaithersburg, MD 20899 U.S. Department of Commerce Ronald H. Brown, Secretary National Institute of Standards and Technology John W. Lyons, Director Abstract This report describes a high speed loading system comprised of: a piezoelectric stack and an associated microprocessor-based programmable DC power source; a low-loss mechanical displacement amplifier; a high pressure spring-loaded axial valve; an integral high pressure valve seat; an expansion nozzle; and a high pressure gas supply. A half square wave voltage signal of varying duration and amplitude was used to drive the piezoelectric stack in such a manner as to produce changing displacements which were subsequently amplified by a monolithically milled hinged anvil. The anvil is connected to a valve core which seats upon, and is normally forced closed upon a specially hardened throat of an expansion nozzle by means of a compression spring. The interior side of the thruster nozzle communicates with a high pressure gas supply. The flow of gas through the nozzle is normally prohibited by the spring loaded valve core. As the piezoelectric stack expands under computer control the valve core is lifted off the nozzle throat permitting gas to expand through the nozzle and create thrust that is directly proportional to the control signal from the microprocessor.
    [Show full text]
  • C. Scott Hill (575) 527-4898 (Home) (575) 635-6834 (Cell)
    Experience Summary Charles Scott Hill C. Scott Hill (575) 527-4898 (home) (575) 635-6834 (cell) Experience Summary: I have many years of direct experience in hypergolic reaction control rocket Development, Qualification, Production, Failure Investigation, Field repair and Application. Was primary contributor for Apollo Service Module Rocket, Manned Orbiting Laboratory RCS, Shuttle Primary and Vernier Thrusters for RCS. Present: Instructor Mechanical Engineering University of Texas at El Paso Teach courses in Spacecraft Design and Spacecraft Propulsion Jan 2010 to present Position Causal Employee - ARES Corporation, Huntsville, Alabama Time in Position April 2006 - December 2009 Provide mentoring and review of specifications, processes and approaches to Reaction Control Integrated Project Team for the Area I vehicle on the NASA Constellation Program at MSFC. Reported directly to NASA IPT Manager. IPT is developing system for monopropellant thrusters for active roll control and attitude control of the 2nd stage. Made monthly trips and direct support using the Internet. Participate in design reviews. Retired from full time employment - November 1, 2005 Position: Staff Advisor to Program Manager – White Sands Test Facility Time in Position: 3 Years Description and Scope of Current Job: Developed and taught instruction courses for WSTF test personnel in rocket development, Shuttle thruster development and qualification testing, application of rockets to space vehicle. Assisted in failure investigation and provided technical project management consulting. Position: Propulsion Department Manager – White Sands Test Facility Time in Position: 7 Years Number of Personnel DirectlySupervised: 6 direct reports with 115-125 people in Department Description and Scope of Job: Responsible for managing Propulsion Department including overall technical performance, budget control and reporting at the department level; the primary direct customer interface with the Propulsion Office Chief to ensure customer satisfaction and department level strategic planning.
    [Show full text]
  • The Design and Analysis of a Kerosene Turbopump for a South African Commercial Launch Vehicle
    The Design and Analysis of a Kerosene Turbopump for a South African Commercial Launch Vehicle. by Jonathan Smyth In fulfilment of the academic requirements for the degree of Master of Science in Mechanical Engineering, College of Agriculture, Engineering and Science, University of KwaZulu-Natal. Supervisor: Mr. Michael Brooks Co-supervisors: Dr. Graham Smith Prof. Jeff Bindon Dr. Glen Snedden (CSIR) DECLARATION 1 - PLAGIARISM I, Jonathan Smyth, declare that 1. The research reported in this thesis, except where otherwise indicated, is my original research. 2. This thesis has not been submitted for any degree or examination at any other university. 3. This thesis does not contain other persons’ data, pictures, graphs or other information, unless specifically acknowledged as being sourced from other persons. 4. This thesis does not contain other persons' writing, unless specifically acknowledged as being sourced from other researchers. Where other written sources have been quoted, then: a. Their words have been re-written but the general information attributed to them has been referenced b. Where their exact words have been used, then their writing has been placed in italics and inside quotation marks, and referenced. 5. This thesis does not contain text, graphics or tables copied and pasted from the Internet, unless specifically acknowledged, and the source being detailed in the thesis and in the References sections. Signed: ……………………………………………………………………………… Mr. Jonathan Smyth As the candidate's supervisor I have approved this dissertation for submission. Signed: ……………………………………………………………………………… Mr. Michael Brooks i DECLARATION 2 - PUBLICATIONS 1. Smyth J.*, Bindon J., Brooks M., Smith G. and Snedden G., "The Design of a Kerosene Turbopump for a South African Commercial Launch Vehicle", 48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit July, Atlanta, Georgia, 2012.
    [Show full text]
  • High Temperature Thruster Technology for Spacecraft Propulsion
    NASA Technical Memorandum 105348 IAF-91-254 High Temperature Thruster Technology for Spacecraft Propulsion Steven J. Schneider Lewis Research Center Cleveland, Ohio Prepared for the 42nd Congress of the International Astronautical Federation Montreal, Canada, October 5-11, 1991 NASA High Temperature Thruster Technology for Spacecraft Propulsion Steven J. Schneider NASA Lewis Research Center Cleveland, OH 44135 Abstract A technology program has been underway since 1985 to develop high temperature oxidation-resistant thrusters for spacecraft applications. The successful development of this technology will provide the basis for the design of higher performance satellite engines with reduced plume contamination. Alternatively, this technology program will provide a material with high thermal margin to operate at conventional temperatures and provide increased life for refuelable or reusable spacecraft. The new chamber material consists of a rhenium substrate coated with iridium for oxidation protection. This material increases the operating temperature of thrusters to 2200° C, a significant increase over the 1400° C of the silicide-coated niobium chambers currently used. Stationkeeping class 22 N engines fabricated from iridium-coated rhenium have demonstrated steady state specific impulses 20 to 25 seconds higher than niobium chambers. Ir-Re apogee class 440 N engines are expected to deliver an additional 10 to 15 seconds. These improved performances are obtained by reducing or eliminating the fuel film cooling requirements in the combustion chamber while operating at the same overall mixture ratio as conventional engines. The program is attempting to envelope flight qualification requirements to reduce the potential risks and costs of flight qualification programs. Introduction The development of rockets suitable for spacecraft propulsion traditionally was accomplished with project support.
    [Show full text]
  • Space Shuttle- Mission Report
    SPACE SHUTTLE- MISSION REPORT (NASA-C9-195740) STS-56 SPACE SHUTTLE t4TSSIDN REPORT (Lockheed Engineering dnd Sciences Co. 52 p Unc 1 as JULY 1993 --- i IT.IIG!H RESEARCH CEIITER LlB!?ll!?Y I!AM IlhT.;?TO:!, VIRGll!lA NASA National Aeronautics and Space Administration Lyndon B. Johnson Space Center Houston, Texas STS-56 SPACE SHUTTLE MISSION REPORT LESC/Flight valuation and Engineering Office Approved by STS-56 Lead Mission Evaluation Room Manager n Manager, Flight ~ngineeringOff ice Manager, 0rbi terwk~Projects / (1 Brewster H: Shaw ~irect';;r, Space Shuttle Operations Prepared by Lockheed Engineering and Sciences Company for Flight Engineering Office NATIONAL AERONAUTICS AND SPACE ADMINISTRATION LYNDON B. JOHNSON SPACE CENTER HOUSTON, TEXAS 77058 July 1993 STS-56 Table of Contents Title Page INTRODUCTION ....................... 1 MISSION SUMMARY ..................... 2 PAYLOADS ......................... ATLAS-2 ........................ Atmos~hericScience ................ Solar Science ............... Shuttle Pointed Autonomous Research Tool for Astronomy-201 ............ Solar Ultraviolet Ex~eriment ........ Hand-~eld. ~arth-0rik fed. Real.Time. Coopera User.Friendly. Location-Targeting and Environmental System Radiation Monitoring Equipment-I11 ........ Cosmic.... Radiation Effects and Activation............ Monitor ... Air..... Force Maui O~tical~ Site ~ ~- .............. Shuttle Amateur Radio Equipment-I1 ........ Commercial Materials Dispersion Apparatus ITA Experiments ................ S~aceTissue Loss Ex~eriment ........... ~h~siolo~icaland
    [Show full text]
  • Spacecraft Propulsion.Pdf
    SPACECRAFT PROPULSION Spacecraft propulsion .................................................................................................................................... 1 Rocket technologies .................................................................................................................................. 4 Cold gas rockets .................................................................................................................................... 7 Solid propellant rockets ........................................................................................................................ 7 Liquid propellant rockets .................................................................................................................... 11 Electrothermal rockets ........................................................................................................................ 16 Electrodynamic rockets ....................................................................................................................... 16 Rocket propulsion modelling .................................................................................................................. 19 Tsiolkovsky rocket equation ............................................................................................................... 19 Delta-v budget ..................................................................................................................................... 20 Rocket exit speed ...............................................................................................................................
    [Show full text]
  • Next Generation Spacecraft, Crew Exploration Vehicle
    Next Generation Spacecraft, Crew Exploration Vehicle 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.
    [Show full text]
  • Columbia Accident Investigation Board (Caib)/ National Aeronautics and Space Administration (Nasa) Accident Investigation Team (Nait)
    COLUMBIA ACCIDENT INVESTIGATION BOARD (CAIB)/ NATIONAL AERONAUTICS AND SPACE ADMINISTRATION (NASA) ACCIDENT INVESTIGATION TEAM (NAIT) WORKING SCENARIO FINAL VERSION: JULY 8, 2003 PREFACE This Working Scenario report was written to document the collection of known facts, events, timelines, and historical information of particular interest to the final flight of Columbia. The report was written with the understanding that it could be published, either in part or in its entirety, as part of the official Columbia Accident Investigation Board (CAIB) report. The report includes information and results from numerous analyses, tests, and simulations related to the Columbia investigation that have been completed, or were ongoing at the time that this report was completed. It is anticipated that additional analytical and test results will emerge from ongoing work, as well as from future activities associated with the Columbia investigation and efforts related to the Return-To-Flight work. This Working Scenario includes information and results as they existed up to and including July 8, 2003. CONTENTS Section Page 1.0 INTRODUCTION ........................................................................................ 1-1 1.1 SCOPE....................................................................................... 1-1 1.2 MISSION BACKGROUND.......................................................... 1-2 2.0 LAUNCH COUNTDOWN............................................................................ 2-1 3.0 LAUNCH....................................................................................................
    [Show full text]