DOCSLIB.ORG

Reliability As an Independent Variable Applied to Liquid Rocket Engine Test Plans

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Reliability As an Independent Variable Applied to Liquid Rocket Engine Test Plans JOURNAL OF PROPULSION AND POWER Vol. 27, No. 5, September–October 2011 Reliability as an Independent Variable Applied to Liquid Rocket Engine Test Plans Richard Strunz∗ EADS Astrium, GmbH, 81663 Munich, Germany and Jeffrey W. Herrmann† University of Maryland, College Park, Maryland 20742 DOI: 10.2514/1.B34279 Manufacturers lack an adequate method to balance performance, reliability, and affordability. The reliability-as- an-independent-variable methodology is the solution proposed by expressing quantitatively the reliability trade space as ranges of a number of hardware sets and a number of hot-fire tests necessary to develop and qualify/certify a liquid rocket engine against a stated reliability requirement. Therefore, reliability-as-an-independent-variable becomes one of the key decision parameters in early tradeoff studies for liquid rocket engines because the reliability trade space directly influences the performance requirements and, as a result, the affordability. The overall solution strategy of reliability-as-an-independent-variable is based on the Bayesian statistical framework using either the planned or actual number of hot-fire tests. The planned hot-fire test results may include test failures to simulate the typical design-fail-fix-test cycles present in liquid rocket engine development programs in order to provide the schedule and cost risk impacts for early tradeoff studies. The reliability-as-an-independent-variable methodology is exemplarily applied to the actual hot-fire test history of the F-1, the space shuttle main engine, and the RS-68 liquid rocket engine, showing adequate agreement between computed results and actual flight engine reliability. Nomenclature = parameter of Poisson probability mass function Beta = Beta probability density function = probability of success of a functional node Bi = Binomial probability density function = posterior probability C = level of confidence (credibility bound) 0 = prior probability c = number of cycles = standard deviation CTD = cumulated test duration, s Subscripts FX = cumulative density function HW = number of hardware D = integration domain L = likelihood function i = functional node n = number of trials, i.e., hot-fire tests j = hot-fire test group p = probability of success med = median q = failure fraction MD = mission duration q = candidate probability density function ind MIQ = mission ignition quantity R = reliability pr = prior used in empirical Bayes estimation r = number of failures rem = remaining t = hardware life, s tot = total tp = hot-fire test proportion u = upper percentile T = median lifetime, s 50 0 = used in empirical Bayes estimation w = weighting factor x = number of successes, i.e., hot-fire tests = shape parameter I. Introduction 1, 2 = weighting factors IQUID rocket engines have always been one of the major = shape parameter affordability drivers of launch vehicles, but in the past, national fi L = median regression coef cients prestige or military requirements were dominating the decisions = integration domain for the development of a new launch vehicle. This paradigm has = parameters to be estimated changed. Performance, reliability, and affordability have equal = acceptance rate importance in the decision-making process. Europe is currently = Gamma function facing this paradigm change by defining the requirements for an expendable next generation launcher in the frame of the ESA’s Future Received 7 March 2011; revision received 25 May 2011; accepted for Launchers Preparatory Program.‡ Various launch vehicle architec- publication 9 June 2011. Copyright © 2011 by Astrium, GmbH. Published by tures were identified, ranging from a two-stage pure liquid rocket- the American Institute of Aeronautics and Astronautics, Inc., with engine-based architecture to a three-stage launch vehicle with two permission. Copies of this paper may be made for personal or internal use, on solid propellant stages and a cryogenic upper-stage engine. Although condition that the copier pay the $10.00 per-copy fee to the Copyright innovative technologies are identified in all relevant areas, the focus Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923; include § the code 0748-4658/11 and $10.00 in correspondence with the CCC. will be on affordability in order to develop a competitive launcher ∗Program Manager, Advanced Programs and Systems Technology; Ph.D. Student, Department of Mechanical Engineering, University of Maryland, ‡Data available at http://www.esa.int/esaMI/Launchers_Home/ College Park, Maryland 20742 (Corresponding Author). SEMNCI1PGQD_0.html [retrieved 30 October 2010]. †Associate Professor, Department of Mechanical Engineering and Institute §Data available at http://www.satnews.com/cgi-bin/story.cgi?number= for Systems Research, Martin Hall, Room 2181. 2083189758 [retrieved 30 October 2010]. 1032 STRUNZ AND HERRMANN 1033 Table 1 Qualification/certification hot-fire test attributes Test attributes F-1 J-2 RL10 LR87 LR91 SSME RS-68a Number of tests required 20 30 20 12 12 10 12 Total test duration required, s 2250 3750 4500 1992 2532 5000 1800 Number of samples 1 2 3 1 1 2 2 Hardware changes allowed Yes Yes No Yes Yes Yes Yes Fleet leader concept used No No No No No Yes No Overstress testing No No Yes No No Yes Yes aValues are based on data given in [6] Table 2 Detailed analysis of qualification/certification hot-fire test attributes Test attributes F-1 J-2 RL10 LR87 LR91 SSME RS-68 Test per hardware 20 15 6.7 12 12 5 6 Test duration per hardware, s 2250 1875 1500 1992 2532 2500 900 Duration per test per hardware, s 112.5 125.0 225.0 166.0 211.0 500.0 150.0 Mission nominal time (max), s 165 500 700 165 225 520 250 Multiple of mission nominal time, s 13.6 3.8 2.1 12.1 11.3 4.8 3.6 but at no expense of the mission success reliability and other qualification or certification tests. In short, the confidence-building launch service competitive factors compared with the current process is dominated by an expensive and schedule-impacting hot- European launch vehicle Ariane 5, as well as other operational fire test program [5]. launch vehicles. Historically, the hot-fire test program definitions experienced an The affordability of expendable launch vehicles is largely deter- evolution from a formal reliability demonstration to an aggressive mined by the development and production costs of liquid rocket cost minimization approach. Initially, liquid rocket engine devel- engines [1,2]. The major part of the development cost is spent on opment programs included a formal reliability demonstration development test hardware subjected to hot-fire tests in order to requirement (e.g., F-1 or J-2) but they were discarded in favor of sufficiently demonstrate design maturity/robustness and to qualify/ design verification specifications (DVSs) [e.g., space shuttle main certify the liquid rocket engines for a successful flight operation [3]. engine (SSME)] due to prohibitively high hot-fire test costs [2]. The The reliability-as-an-independent-variable (RAIV) methodology most recent approach is the objective-based variable test/time philos- provides the framework for specifying qualification/certification hot- ophy executed for qualifying the RS-68 liquid rocket engine that fire test attributes in terms of the number of tests, number of hardware required the least amount of hot-fire tests and accumulated hot-fire sets, and total test duration that are allocated at the component, the test duration (HFTD) [6]. subsystem, and the engine system level. The production cost is Although these different test program philosophies were applied mainly driven by performance and reliability requirements that can for various liquid rocket engines with large performance differences, be transferred into a manufacturing complexity expressed as a one may wonder why no significant trend can be seen on the number of parts, precision of the parts, and selected materials. One of qualification/certification hot-fire test attributes as listed in Table 1 the main leverages on the development cost is the chosen verification [7]. The numbers of tests required per hardware set are higher for the strategy with the objective to minimize the number of hardware sets F-1 and J-2 compared with the SSME, which were all man rated, and by testing the mission requirements on a single hardware set multiple the SSME is even reusable but subjected to different hot-fire test times but at the expense of a higher production cost because of definitions, i.e., the formal reliability demonstration versus the DVS. increased performance requirements, the selection of special mate- The J-2 and RL10 are both cryogenic upper-stage liquid rocket rials, and the need for elevated manufacturing precisions in order to engines, but hardware changes were allowed only for the J-2 and not guarantee the longer life capability. Affordability of the launch for the RL10. Table 2 may reveal the only difference among the test vehicle would be incomplete without the consideration of vehicle attributes that is linked to the propellant combination used and the operation and support, mission assurance, and range cost, as well as resulting internal loads present during engine operation; that is, more insurance fees [4]. tests and, as a consequence, a higher accumulated test duration, Therefore, the selection of the optimal choice for the settings of the which is expressed as a number of multiple mission durations, is conflicting trade spaces for performance, reliability, and affordability placed on hardware sets for the propellant combination liquid oxygen becomes a multiple-criterion decision-making (MCDM) problem. (LOX)/liquid hydrogen (two- to fivefold) compared with the The trade spaces for affordability and performance are generated propellant combination LOX/kerosene or hypergolic storable pro- with parametric cost models and thermodynamic cycle codes. pellants (more than tenfold). This identified difference may be biased However, the main shortcoming of the current MCDM solutions is by the lack of visibility on the extent of the prior component level or the lack of an adequate modeling technique for the reliability trade the development engine test history.
Load more

Recommended publications
  • Rocket Propulsion Fundamentals 2
    https://ntrs.nasa.gov/search.jsp?R=20140002716 2019-08-29T14:36:45+00:00Z Liquid Propulsion Systems – Evolution & Advancements Launch Vehicle Propulsion & Systems LPTC Liquid Propulsion Technical Committee Rick Ballard Liquid Engine Systems Lead SLS Liquid Engines Office NASA / MSFC All rights reserved. No part of this publication may be reproduced, distributed, or transmitted, unless for course participation and to a paid course student, in any form or by any means, or stored in a database or retrieval system, without the prior written permission of AIAA and/or course instructor. Contact the American Institute of Aeronautics and Astronautics, Professional Development Program, Suite 500, 1801 Alexander Bell Drive, Reston, VA 20191-4344 Modules 1. Rocket Propulsion Fundamentals 2. LRE Applications 3. Liquid Propellants 4. Engine Power Cycles 5. Engine Components Module 1: Rocket Propulsion TOPICS Fundamentals • Thrust • Specific Impulse • Mixture Ratio • Isp vs. MR • Density vs. Isp • Propellant Mass vs. Volume Warning: Contents deal with math, • Area Ratio physics and thermodynamics. Be afraid…be very afraid… Terms A Area a Acceleration F Force (thrust) g Gravity constant (32.2 ft/sec2) I Impulse m Mass P Pressure Subscripts t Time a Ambient T Temperature c Chamber e Exit V Velocity o Initial state r Reaction ∆ Delta / Difference s Stagnation sp Specific ε Area Ratio t Throat or Total γ Ratio of specific heats Thrust (1/3) Rocket thrust can be explained using Newton’s 2nd and 3rd laws of motion. 2nd Law: a force applied to a body is equal to the mass of the body and its acceleration in the direction of the force.
    [Show full text]
  • Materials for Liquid Propulsion Systems
    https://ntrs.nasa.gov/search.jsp?R=20160008869 2019-08-29T17:47:59+00:00Z CHAPTER 12 Materials for Liquid Propulsion Systems John A. Halchak Consultant, Los Angeles, California James L. Cannon NASA Marshall Space Flight Center, Huntsville, Alabama Corey Brown Aerojet-Rocketdyne, West Palm Beach, Florida 12.1 Introduction Earth to orbit launch vehicles are propelled by rocket engines and motors, both liquid and solid. This chapter will discuss liquid engines. The heart of a launch vehicle is its engine. The remainder of the vehicle (with the notable exceptions of the payload and guidance system) is an aero structure to support the propellant tanks which provide the fuel and oxidizer to feed the engine or engines. The basic principle behind a rocket engine is straightforward. The engine is a means to convert potential thermochemical energy of one or more propellants into exhaust jet kinetic energy. Fuel and oxidizer are burned in a combustion chamber where they create hot gases under high pressure. These hot gases are allowed to expand through a nozzle. The molecules of hot gas are first constricted by the throat of the nozzle (de-Laval nozzle) which forces them to accelerate; then as the nozzle flares outwards, they expand and further accelerate. It is the mass of the combustion gases times their velocity, reacting against the walls of the combustion chamber and nozzle, which produce thrust according to Newton’s third law: for every action there is an equal and opposite reaction. [1] Solid rocket motors are cheaper to manufacture and offer good values for their cost.
    [Show full text]
  • Interstellar Probe on Space Launch System (Sls)
    INTERSTELLAR PROBE ON SPACE LAUNCH SYSTEM (SLS) David Alan Smith SLS Spacecraft/Payload Integration & Evolution (SPIE) NASA-MSFC December 13, 2019 0497 SLS EVOLVABILITY FOUNDATION FOR A GENERATION OF DEEP SPACE EXPLORATION 322 ft. Up to 313ft. 365 ft. 325 ft. 365 ft. 355 ft. Universal Universal Launch Abort System Stage Adapter 5m Class Stage Adapter Orion 8.4m Fairing 8.4m Fairing Fairing Long (Up to 90’) (up to 63’) Short (Up to 63’) Interim Cryogenic Exploration Exploration Exploration Propulsion Stage Upper Stage Upper Stage Upper Stage Launch Vehicle Interstage Interstage Interstage Stage Adapter Core Stage Core Stage Core Stage Solid Solid Evolved Rocket Rocket Boosters Boosters Boosters RS-25 RS-25 Engines Engines SLS Block 1 SLS Block 1 Cargo SLS Block 1B Crew SLS Block 1B Cargo SLS Block 2 Crew SLS Block 2 Cargo > 26 t (57k lbs) > 26 t (57k lbs) 38–41 t (84k-90k lbs) 41-44 t (90k–97k lbs) > 45 t (99k lbs) > 45 t (99k lbs) Payload to TLI/Moon Launch in the late 2020s and early 2030s 0497 IS THIS ROCKET REAL? 0497 SLS BLOCK 1 CONFIGURATION Launch Abort System (LAS) Utah, Alabama, Florida Orion Stage Adapter, California, Alabama Orion Multi-Purpose Crew Vehicle RL10 Engine Lockheed Martin, 5 Segment Solid Rocket Aerojet Rocketdyne, Louisiana, KSC Florida Booster (2) Interim Cryogenic Northrop Grumman, Propulsion Stage (ICPS) Utah, KSC Boeing/United Launch Alliance, California, Alabama Launch Vehicle Stage Adapter Teledyne Brown Engineering, California, Alabama Core Stage & Avionics Boeing Louisiana, Alabama RS-25 Engine (4)
    [Show full text]
  • Delta IV Parker Solar Probe Mission Booklet
    A United Launch Alliance (ULA) Delta IV Heavy what is the source of high-energy solar particles. MISSION rocket will deliver NASA’s Parker Solar Probe to Parker Solar Probe will make 24 elliptical orbits an interplanetary trajectory to the sun. Liftoff of the sun and use seven flybys of Venus to will occur from Space Launch Complex-37 at shrink the orbit closer to the sun during the Cape Canaveral Air Force Station, Florida. NASA seven-year mission. selected ULA’s Delta IV Heavy for its unique MISSION ability to deliver the necessary energy to begin The probe will fly seven times closer to the the Parker Solar Probe’s journey to the sun. sun than any spacecraft before, a mere 3.9 million miles above the surface which is about 4 OVERVIEW The Parker percent the distance from the sun to the Earth. Solar Probe will At its closest approach, Parker Solar Probe will make repeated reach a top speed of 430,000 miles per hour journeys into the or 120 miles per second, making it the fastest sun’s corona and spacecraft in history. The incredible velocity trace the flow of is necessary so that the spacecraft does not energy to answer fall into the sun during the close approaches. fundamental Temperatures will climb to 2,500 degrees questions such Fahrenheit, but the science instruments will as why the solar remain at room temperature behind a 4.5-inch- atmosphere is thick carbon composite shield. dramatically Image courtesy of NASA hotter than the The mission was named in honor of Dr.
    [Show full text]
  • Space Almanac 2005
    SpaceAl2005 manac Stratosphere begins 10 miles Limit for turbojet engines 20 miles Limit for ramjet engines 28 miles Astronaut wings awarded 50 miles Low Earth orbit begins 60 miles 0.95G 100 miles Medium Earth orbit begins 300 miles 44 44 AIR FORCEAIR FORCE Magazine Magazine / August / August 2005 2005 SpaceAl manacThe US military space operation in facts and figures. Compiled by Tamar A. Mehuron, Associate Editor, and the staff of Air Force Magazine Hard vacuum 1,000 miles Geosynchronous Earth orbit 22,300 miles 0.05G 60,000 miles NASA photo/staff illustration by Zaur Eylanbekov Illustration not to scale AIR FORCE Magazine / August 2005 AIR FORCE Magazine / /August August 2005 2005 4545 US Military Missions in Space Space Force Support Space Force Enhancement Space Control Space Force Application Launch of satellites and other Provide satellite communica- Assure US access to and freedom Pursue research and devel- high-value payloads into space tions, navigation, weather, mis- of operation in space and deny opment of capabilities for the and operation of those satellites sile warning, and intelligence to enemies the use of space. probable application of combat through a worldwide network of the warfighter. operations in, through, and from ground stations. space to influence the course and outcome of conflict. US Space Funding Millions of constant FY06 dollars $50,000 DOD 45,000 NASA 40,000 Other Total 35,000 30,000 25,000 20,000 15,000 10,000 5,000 0 59 62 66 70 74 78 82 86 90 94 98 02 04 Fiscal Year FY NASA DOD Other Total FY NASA DOD Other
    [Show full text]
  • An Investigation of the Performance Potential of A
    AN INVESTIGATION OF THE PERFORMANCE POTENTIAL OF A LIQUID OXYGEN EXPANDER CYCLE ROCKET ENGINE by DYLAN THOMAS STAPP RICHARD D. BRANAM, COMMITTEE CHAIR SEMIH M. OLCMEN AJAY K. AGRAWAL A THESIS Submitted in partial fulfillment of the requirements for the degree of Master of Science in the Department of Aerospace Engineering and Mechanics in the Graduate School of The University of Alabama TUSCALOOSA, ALABAMA 2016 Copyright Dylan Thomas Stapp 2016 ALL RIGHTS RESERVED ABSTRACT This research effort sought to examine the performance potential of a dual-expander cycle liquid oxygen-hydrogen engine with a conventional bell nozzle geometry. The analysis was performed using the NASA Numerical Propulsion System Simulation (NPSS) software to develop a full steady-state model of the engine concept. Validation for the theoretical engine model was completed using the same methodology to build a steady-state model of an RL10A-3- 3A single expander cycle rocket engine with corroborating data from a similar modeling project performed at the NASA Glenn Research Center. Previous research performed at NASA and the Air Force Institute of Technology (AFIT) has identified the potential of dual-expander cycle technology to specifically improve the efficiency and capability of upper-stage liquid rocket engines. Dual-expander cycles also eliminate critical failure modes and design limitations present for single-expander cycle engines. This research seeks to identify potential LOX Expander Cycle (LEC) engine designs that exceed the performance of the current state of the art RL10B-2 engine flown on Centaur upper-stages. Results of this research found that the LEC engine concept achieved a 21.2% increase in engine thrust with a decrease in engine length and diameter of 52.0% and 15.8% respectively compared to the RL10B-2 engine.
    [Show full text]
  • Additive Manufacturing Development Methodology for Liquid Rocket Engines
    Additive Manufacturing Development Methodology for Liquid Rocket Engines Quality in the Space and Defense Industry Forum Cape Canaveral, FL March 8, 2016 Jeff Haynes Aerojet Rocketdyne Distribution Statement A: Approved for public release, distribution is unlimited Presentation Outline •The additive manufacturing “opportunity” •Specific additive manufacturing process considerations •Scale up challenges •Aerojet Rocketdyne development and production approach •Considerations and gaps to be closed for production Distribution Statement A: Approved for public release, distribution is unlimited Demonstrated Benefits of Additive Manufacturing Liquid Rocket Engine Attributes Additive Manufacturing Low production volumes Print parts when needed High complexity, compact designs Complexity adds no cost High value = high quality levels Bulk material like wrought not cast Additive Manufacturing • Part count reduction • No plating / no braze • No tooling • 60% lead time savings • 70% cost savings • Reduced weight 9-lbs Heritage Saturn V F1 Gas Generator Injector Printed F1 GG Injector (2009) Early Demonstrated Realization of Potential Distribution Statement A: Approved for public release, distribution is unlimited Demonstrated Benefits of Additive Manufacturing Transforming heritage engines and … enabling new ones 14-inch • Complex injector assemblies diameter Reduce part count Eliminates long lead forgings Eliminates high touch labor machining RS-25 Ball Shaft (left) Eliminates hundreds of braze joints RL10 Main Injector (right) • Sheet
    [Show full text]
  • Methodology for Identifying Key Parameters Affecting Reusability for Liquid Rocket Engines
    Methodology for Identifying Key Parameters Affecting Reusability for Liquid Rocket Engines Presented by: Bruce Askins Rhonda Childress-Thompson Marshall Space Flight Center & Dr. Dale Thomas The University of Alabama in Huntsville Outline • Objective • Reusability Defined • Benefits of Reusability • Previous Research • Approach • Challenges • Available Data • Approach • Results • Next Steps • Summary Objective • To use statistical techniques to identify which parameters are tightly correlated with increasing the reusability of liquid rocket engine hardware. Reusability/Reusable Defined Ideally • “The ability to use a system for multiple missions without the need for replacement of systems or subsystems. Ideally, only replenishment of consumable commodities (propellant and gas products, for instance)occurs between missions.” (Jefferies, et al., 2015) For purposes of this research • Any space flight hardware that is not only designed to perform multiple flights, but actually accomplishes multiple flights. Benefits of Reusability • Forces inspections of returning hardware • Offers insight into how a part actually performs • Allows the development of databases for future development • Validates ground tests and analyses • Allows the expensive hardware to be used multiple times Previous Research – Identification of Features Conducive for Reuse Implementation 1. Reusability Requirement Implemented at the Conceptual Stage 2. Continuous Test Program 3. Minimize Post-Flight Inspections & Servicing to Enhance Turnaround Time 4. Easy Access 5. Longer Service Life 6. Minimize Impact of Recovery 7. Evolutionary vs. Revolutionary Changes MSFC Legacy of Propulsion Excellence (Evolutionary Changes) Saturn V J-2S 1970 1980 1990 2000 2010 2015 Challenges • Engine data difficult to locate • Limited data for reusable engines (i.e. Space Shuttle, Space Shuttle Main Engine (SSME) and Solid Rocket Booster (SRB) • Benefits of reusability have not been validated • Industry standards do not exist Engine Data Available Engine Time from No.
    [Show full text]
  • A Methodology for Preliminary Sizing of a Thermal and Energy Management System for a Hypersonic Vehicle
    A methodology for preliminary sizing of a Thermal and Energy Management System for a hypersonic vehicle. Roberta Fusaro1, Davide Ferretto 1, Valeria Vercella1, Nicole Viola1, Victor Fernandez Villace2 and Johan Steelant2 Abstract This paper addresses a methodology to parametrically size thermal control subsystems for high-speed transportation systems. This methodology should be sufficiently general to be exploited for the derivation of Estimation Relationships (ERs) for geometrically sizing characteristics as well as mass, volume and power budgets both for active (turbopumps, turbines and compressors) and passive components (heat exchangers, tanks and pipes). Following this approach, ad-hoc semi-empirical models relating the geometrical sizing, mass, volume and power features of each component to operating conditions have been derived. As a specific case, a semi-empirical parametric model for turbopumps sizing is derived. In addition, the Thermal and Energy Management Subsystem (TEMS) for the LAPCAT MR2 vehicle is used as an example of a highly integrated multifunctional subsystem. The TEMS is based on the exploitation of liquid hydrogen boil-off in the cryogenic tanks generated by the heat load penetrating the aeroshell, all along the point-to-point hypersonic mission. Eventually, specific comments about the results will be provided together with suggestions for future improvements. Keywords: Thermal and Energy Management System, sizing models, turbopumps, LAPCAT MR2 1. Introduction The high operating temperature characterizing the hypersonic flight regime is a long-term issue. Vehicles able to reach hypersonic speed are currently considered for both aeronautical (e.g. high speed antipodal transportation systems) and space transportation purposes (e.g. reusable launcher stages and re-entry systems).
    [Show full text]
  • Heavy Lift Launch Vehicles with Existing Propulsion Systems
    SpaceOps 2010 Conference<br><b><i>Delivering on the Dream</b></i><br><i>Hosted by NASA Mars AIAA 2010-2370 25 - 30 April 2010, Huntsville, Alabama Heavy Lift Launch Vehicles with Existing Propulsion Systems Benjamin Donahue1 Lee Brady2 Mike Farkas3 Shelley LeRoy4 Neal Graham5 Boeing Phantom Works, Huntsville, AL 35824 Doug Blue6 Boeing Space Exploration, Huntington Beach, CA 92605 This paper describes Heavy Lift Launch Vehicle concepts that are based on existing propulsion systems. Both In-Line and Sidemount configurations for Crew, Crew plus cargo and Cargo only missions are illustrated. Payload data includes launches to due East LEO, ISS, Trans-Lunar Injection (TLI) and the Earth-Sun L2 point. Engine options include SSME and RS-68 for the Core stage and J-2X and RL-10 engines for Upper stages. Heavy Lift would provide the large volumes and heavy masses required to enable high science return missions, while utilizing proven propulsion elements. I. Introduction Both In-line and Sidemount Heavy Lift Launch Vehicle (HLLV) concepts, utilizing Solid Rocket Booster (SRB) and Space Shuttle Main Engine (SSME) elements, would enable exploration missions1-6 that might otherwise be impractical with current launch vehicles. Potential missions and payloads (Fig. 1) include space telescopes, fuel depots, Mars, Venus, Europa, and Titan sample return vehicles, Crewed Lunar and Near Earth Object (NEO) vehicles, power beaming platforms and others. The use of existing main propulsion systems (SSME, RS-68 engines, SRBs) would minimize the upfront cost and shorten the time to initial operational capability (IOC) of any new HLLV as compared to a similar program with new propulsion elements.
    [Show full text]
  • SLS Case Study
    CASE STUDY Supporting Space Flight History with the Space Launch System (SLS) Technetics Group is a proud sup- Boeing representatives held a sup- plier to NASA, Boeing and Aerojet plier recognition presentation for Rocketdyne, and proved instru- the Technetics team members, mental in working with these orga- citing outstanding performance nizations on the development of in providing hydraulic accumula- NASA’s Space Launch System (SLS). tors and reservoirs for the Thrust The SLS is an advanced, heavy-lift Vector Control hydraulic system, BELFAB Edge-Welded launch vehicle that will send astro- located within the Core Stage of Metal Bellows nauts into deep space and open the rocket. Technetics was one up the possibility for missions to of the first suppliers on the pro- Qualiseal Mechanical neighboring planets. The program gram to have provided all Flight 1 Seals is enabling humans to travel fur- requirements and subsystem test ther into space than ever before units. and paving the way for new sci- Aerojet Rocketdyne also rec- entific discoveries and knowledge ognized Technetics and stated that was once out of reach. that Technetics “has gone above Technetics has worked closely and beyond to produce quality with program design teams for hardware and support aggressive critical applications for the Core schedules,” and that “Technetics Stage and Upper Stage on the SLS. efforts and those of Technetics NAFLEX Seals Specifically, a number of preci- employees have not gone unno- sion sealing solutions and fluid ticed.” To express their appre- management components were ciation, Aerojet Rocketdyne needed for Aerojet Rocketdyne’s representatives visited the RS-25 and RL10 engines to ensure Technetics Deland, FL facility to the integrity of the overall system.
    [Show full text]
  • Advanced Tube-Bundle Rocket Thrust Chamber
    https://ntrs.nasa.gov/search.jsp?R=19900015869 2020-03-19T22:32:33+00:00Z NASA Technical Memorandum 103139 AIAA-90-2726 Advanced Tube-Bundle Rocket Thrust Chamber John M. Kazaroff National Aeronautics and Space Administration Lewis Research Center Cleveland, Ohio and Albert J. Pavli Sverdrup Technology, Inc. Lewis Research Center Group Brook Park, Ohio Prepared for the 26th Joint Propulsion Conference cosponsored by the AIAA, SAE, ASME, and ASEE Orlando, Florida, July 16-18, 1990 NASA ADVANCED TUBE-BUNDLE ROCKET THRUST CHAMBER John M. Kazaroff National Aeronautics and Space Administration Lewis Research Center Cleveland, Ohio 44135 and Albert J. Pavli Sverdrup Technology, Inc. Lewis Research Center Group Brook Park, Ohio 44142 o SUMMARY a- w An advanced rocket thrust chamber for future space application is described along with an improved method of fabrication. Potential benefits of the concept are improved cyclic life, reusability, reliability, and perform- ance. Performance improvements are anticipated because of the enhanced heat transfer into the coolant which will enable higher chamber pressure in expander cycle engines. Cyclic life, reusability and reliability improvements are anti- cipated because of the enhanced structural compliance inherent in the construc- tion. The method of construction involves the forming of the combustion chamber with a tube-bundle of high conductivity copper or copper alloy tubes, and the bonding of these tubes by an electroforming operation. Further, the method of fabrication reduces chamber complexity by incorporating manifolds, jackets, and structural stiffeners while having the potential for thrust cham- ber cost and weight reduction. BACKGROUND Through the years many forms of rocket combustion chamber construction have been successfully employed.
    [Show full text]