DOCSLIB.ORG

Staged Combustion Cycle Rocket Engine Subsystem Definition for Future Advanced Passenger Transport Martin Sippel, Tobias Schwanekamp [email protected] Tel

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Staged Combustion Cycle Rocket Engine Subsystem Definition for Future Advanced Passenger Transport Martin Sippel, Tobias Schwanekamp Martin.Sippel@Dlr.De Tel View metadata, citation and similar papers at core.ac.uk brought to you by CORE provided by Institute of Transport Research:Publications Space Propulsion 2014, Session 30 - ST - Future Liquid Stages & Engines (3) Staged Combustion Cycle Rocket Engine Subsystem Definition for Future Advanced Passenger Transport Martin Sippel, Tobias Schwanekamp [email protected] Tel. +49-421-244201145 Space Launcher Systems Analysis (SART), DLR, Bremen, Germany Markus Ortelt Institute of Construction and Design, DLR, Stuttgart, Germany DLR’s launcher systems analysis division is investigating since a couple of years a visionary, extremely fast passenger transportation concept based on rocket propulsion. Thanks to the multi-national collaboration, the technical lay-out of the SpaceLiner has now matured to Phase A conceptual design level. Full-flow staged combustion cycle rocket engines with a moderate 15 to 17 MPa range in chamber pressure have been selected as the baseline propulsion system. The expansion ratios of the engines are adapted to their respective optimums required by the stages; while the mass flow, turbo-machinery, and combustion chamber are assumed to remain identical. The paper describes the SpaceLiner 7 propulsion system: • The reference vehicle’s preliminary design, • Main propulsion system definition and architectural lay-out, • Thrust chamber geometries, • Pre-design of different turbomachinery and attached preburners, • Advanced ceramic material fuel- and oxidizer-rich pre-burners and injectors as an alternative to increase lifetime of components. The presented work is including preliminary sizing on component level and first mass estimation data. Nomenclature 1 INTRODUCTION c* characteristic velocity m / s A strategic vision of DLR which ultimately has the Isp (mass) specific Impulse s (N s / kg) potential to enable sustainable low-cost space M Mach-number - transportation to orbit is under technical evaluation T Thrust N since a couple of years. The number of launches per year m mass kg should be strongly raised and hence manufacturing and operating cost of launcher hardware should dramatically shrink. The obvious challenge of the vision is to identify ε expansion ratio - the very application creating this new, large-size market. Subscripts, Abbreviations Ultra long distance travel from one major business center of the world to another major agglomeration on CFRP Carbon Fiber Reinforced Plastics earth is a huge and mature market. Since the termination CMC Ceramic Matrix Composite of Concorde operation, intercontinental travel is FFSC Full-Flow Staged Combustion restricted to low-speed, subsonic, elongated multi-hour flight. An interesting alternative to air-breathing FRSC Fuel-Rich Staged Combustion hypersonic passenger airliners in the field of future high- FTP Fuel Turbo Pump speed intercontinental passenger transport vehicles LH2 Liquid Hydrogen might be a rocket-propelled, suborbital craft. Such a new LOX Liquid Oxygen kind of ‘space tourism’ based on a two stage RLV has MECO Main Engine Cut Off been proposed by DLR under the name SpaceLiner [1]. MR mixture ratio Ultra long-haul distances like Europe – Australia could NPSP Net Positive Suction Pressure be flown in 90 minutes. Another interesting intercontinental destination between Europe and North- OTP Oxidizer Turbo Pump West America could be reduced to flight times of SLME SpaceLiner Main Engine slightly more than one hour. SSME Space Shuttle Main Engine TET Turbine Entry Temperature Ultra-fast transportation far in excess of supersonic and TRL Technology Readiness Level even potential hypersonic airplanes is definitely a fundamental new application for launch vehicles. By no C chamber more than partially tapping the huge intercontinental vac vacuum travel and tourism market, production rates of RLVs and their rocket engines could increase hundredfold which is Copyright 2014 by DLR-SART 1 out of reach for all other known earth-orbit space subsequent configuration numbering has been transportation. The fast intercontinental travel space established for all those types investigated in sufficient tourism, not only attracting the leisure market, would, as level of detail. The genealogy of the different a byproduct, also enable to considerably reduce the cost SpaceLiner versions is shown in Figure 2. The box is of space transportation to orbit. marking the configuration trade-offs performed in FAST20XX in 2009/10. 2005 SpaceLiner (1) 2006 2007 SpaceLiner 2 SpaceLiner 3 2008 2009 configuration trade-offs SpaceLiner 4 SpaceLiner 5 LH2 / RP1 SpaceLiner 6 2010 Single Stage 2011 2012 SpaceLiner 7 Figure 2: Evolution of the SpaceLiner concept At the end of 2012 with conclusion of FAST20XX the SpaceLiner 7 reached a consolidated technical status. 2.2 Latest SpaceLiner 7 Configuration The general baseline design concept consists of a fully reusable booster and passenger stage arranged in parallel. All rocket engines should work from lift-off until MECO. A propellant crossfeed from the booster to the passenger stage (also called orbiter) is foreseen up to Figure 1: The SpaceLiner vision of a rocket-propelled separation to reduce the overall size of the intercontinental passenger transport could push configuration. After a rapid acceleration to its maximum spaceflight further than any other credible scenario speed the hypersonic transport is gliding for the remaining more than one hour flight to its destination. 2 TECHNICAL OVERVIEW OF THE The current arrangement of the two vehicles at lift-off is SPACELINER CONCEPT presented in Figure 3. Stage attachments are following a classical tripod design. The axial thrust of the booster is introduced through the forward attachment from booster 2.1 Configuration Evolution intertank into the nose gear connection structure of the First proposed in 2005 [1], the SpaceLiner is under orbiter. The aft attachment takes all side and constant development and descriptions of some major maneuvering loads. Major geometrical data of the updates have been published since then [2, 3, 4]. The th SpaceLiner7 stages are provided in Table 1 and Table 2. European Union’s 7 Research Framework Programme has supported several important aspects of Total dry mass of the SpaceLiner 7 launch configuration multidisciplinary and multinational cooperation in the is estimated at 310.9 Mg with a total propellant loading projects FAST20XX, CHATT [6], HIKARI and of 1520 Mg resulting in 1838.7 Mg GLOW incl. HYPMOCES. passengers & payload. Latest improvements of the SpaceLiner 7-3 configuration allow for a slight Different configurations in terms of propellant reduction in masses. combinations, staging, aerodynamic shapes, and structural architectures have been analyzed. A Table 1: Geometrical data of SpaceLiner 7-2 booster stage length [m] span [m] height [m] fuselage wing leading wing pitch wing dihedral diameter [m] edge angles angle [deg] angle [deg] [deg] 83.5 36.0 8.7 8.6 82/61/43 3.5 0 Table 2: Geometrical data of SpaceLiner 7-1 passenger stage length [m] span [m] height [m] fuselage wing leading wing pitch wing dihedral diameter [m] edge angle angle [deg] angle [deg] [deg] 65.6 33.0 12.1 6.4 70 0.4 2.65 2 Figure 3: Sketch of SpaceLiner 7 launch configuration with passenger stage on top and booster stage at bottom position showing the SLME arrangement in the lower right figure LOX orbiter feedline LOX cross-feed line LH2 orbiter feedline LH2 crossfeed line Figure 4: Arrangement of propellant tanks, feed- and pressurization system of SpaceLiner 7 preburner gas turbine driving the LH2-pump and an 3 MAIN PROPULSION SYSTEM oxidizer-rich preburner gas turbine driving the LOX- Staged combustion cycle rocket engines around a pump is a preferred design solution for the SpaceLiner. moderate 16 MPa chamber pressure have been selected This approach should allow avoiding the complexity and as the SpaceLiner main propulsion system called SLME cost of additional inert gases like Helium for sealing. (SpaceLiner Main Engine). Such engine operational data are not overly ambitious and have already been 3.2 Propellant feed and tank pressurization exceeded by existing engines like SSME or RD-0120. system However, the ambitious goal of a passenger rocket is to All main engines of the configuration should work from considerably enhance reliability and reusability of the lift-off until MECO. A propellant crossfeed from the engines beyond the current state of the art. The booster to the passenger stage is foreseen up to expansion ratios of the booster and orbiter engines are separation to reduce the latter’s overall size. No adapted to their respective optimums; while the mass crossfeed system for a configuration like the SpaceLiner flow, turbo-machinery, and combustion chamber are has ever been built and therefore investigations have assumed to remain identical in the baseline been performed to determine how such a system could configuration. be implemented and how complexity issues can be addressed. 3.1 Previous Engine Analyses Three main options of crossfeed exist: The best mixture ratio of the SpaceLiner main • Line-to-line propulsion system along its mission has been defined by • Tank-to-tank system analyses optimizing the full trajectory. Nominal • Tank-to-buffer-tank engine MR control at two engine operation points (6.5 These options are investigated in the FP7-project from lift-off until reaching 2.5 g acceleration and 5.5 CHATT with steady-state flow-simulation along the full afterwards) is found most promising [5]. powered trajectory and transient simulation of critical phases like engine cut-off or valve closing. In particular, Two types of staged combustion cycles (one full-flow the process of booster separation is a dimensioning and the other fuel-rich) have been considered for the factor for the design of the crossfeed system due to the SLME and traded by numerical cycle analyses [5, 7]. A switch of the propellant supply from the booster to the Full-Flow Staged Combustion Cycle with a fuel-rich orbiter tanks. 3 The propellant feed- and pressurization system is kN, 2000 kN respectively for the passenger stage. All preliminarily designed using the DLR-tool pmp.
Load more

Recommended publications
  • IAF Space Propulsion Symposium 2019
    IAF Space Propulsion Symposium 2019 Held at the 70th International Astronautical Congress (IAC 2019) Washington, DC, USA 21 -25 October 2019 Volume 1 of 2 ISBN: 978-1-7138-1491-7 Printed from e-media with permission by: Curran Associates, Inc. 57 Morehouse Lane Red Hook, NY 12571 Some format issues inherent in the e-media version may also appear in this print version. Copyright© (2019) by International Astronautical Federation All rights reserved. Printed with permission by Curran Associates, Inc. (2020) For permission requests, please contact International Astronautical Federation at the address below. International Astronautical Federation 100 Avenue de Suffren 75015 Paris France Phone: +33 1 45 67 42 60 Fax: +33 1 42 73 21 20 www.iafastro.org Additional copies of this publication are available from: Curran Associates, Inc. 57 Morehouse Lane Red Hook, NY 12571 USA Phone: 845-758-0400 Fax: 845-758-2633 Email: [email protected] Web: www.proceedings.com TABLE OF CONTENTS VOLUME 1 PROPULSION SYSTEM (1) BLUE WHALE 1: A NEW DESIGN APPROACH FOR TURBOPUMPS AND FEED SYSTEM ELEMENTS ON SOUTH KOREAN MICRO LAUNCHERS ............................................................................ 1 Dongyoon Shin KEYNOTE: PROMETHEUS: PRECURSOR OF LOW-COST ROCKET ENGINE ......................................... 2 Jérôme Breteau ASSESSMENT OF MON-25/MMH PROPELLANT SYSTEM FOR DEEP-SPACE ENGINES ...................... 3 Huu Trinh 60 YEARS DLR LAMPOLDSHAUSEN – THE EUROPEAN RESEARCH AND TEST SITE FOR CHEMICAL SPACE PROPULSION SYSTEMS ....................................................................................... 9 Anja Frank, Marius Wilhelm, Stefan Schlechtriem FIRING TESTS OF LE-9 DEVELOPMENT ENGINE FOR H3 LAUNCH VEHICLE ................................... 24 Takenori Maeda, Takashi Tamura, Tadaoki Onga, Teiu Kobayashi, Koichi Okita DEVELOPMENT STATUS OF BOOSTER STAGE LIQUID ROCKET ENGINE OF KSLV-II PROGRAM .......................................................................................................................................................
    [Show full text]
  • Basic Analysis of a LOX/Methane Expander Bleed Engine
    DOI: 10.13009/EUCASS2017-332 7TH EUROPEAN CONFERENCE FOR AERONAUTICS AND AEROSPACE SCIENCES (EUCASS) DOI: ADD DOINUMBER HERE Basic Analysis of a LOX/Methane Expander Bleed Engine ? ? ? Marco Leonardi , Francesco Nasuti † and Marcello Onofri ?Sapienza University of Rome Via Eudossiana 18, Rome, Italy [email protected] [email protected] [email protected] · · †Corresponding author Abstract As present trends in rocket engine development recommend overall simplicity and reliability as the main design driver, while preserving high performance, expander cycle engines based on the oxygen-methane pair have been considered as a possible upper stage option. A closed expander cycle is considered for Vega Evolution upper stage, while there are no studies published in the literature on methane-based expander bleed cycles. A basic cycle analysis is presented to evaluate the performance of an oxygen/methane ex- pander bleed cycle for an engine of 100 kN thrust class. Results show the feasibility of the system and its peculiarities with respect to the better known expander bleed cycle based on hydrogen. 1. Introduction The high chamber pressure required to achieve high specific impulse in liquid propellant rocket engines (LRE), has been efficiently obtained by pump-fed systems. Different solutions have been proposed since the beginning of space age and just a few of them has found its own field of application. In these systems the pumps are driven by gas turbines whose power comes from two possible sources: combustion or cooling system. The different needs for the specific applications (booster, sustainer or upper stage of different classes of rockets) led to classify pump-fed LRE systems in open and closed cycles, which differ because of turbine discharge pressure.14, 16 Closed cycles are those providing the best performance because the whole propellant mass flow rate is exploited in the main chamber.
    [Show full text]
  • Space Shuttle Main Engine Orientation
    BC98-04 Space Transportation System Training Data Space Shuttle Main Engine Orientation June 1998 Use this data for training purposes only Rocketdyne Propulsion & Power BOEING PROPRIETARY FORWARD This manual is the supporting handout material to a lecture presentation on the Space Shuttle Main Engine called the Abbreviated SSME Orientation Course. This course is a technically oriented discussion of the SSME, designed for personnel at any level who support SSME activities directly or indirectly. This manual is updated and improved as necessary by Betty McLaughlin. To request copies, or obtain information on classes, call Lori Circle at Rocketdyne (818) 586-2213 BOEING PROPRIETARY 1684-1a.ppt i BOEING PROPRIETARY TABLE OF CONTENT Acronyms and Abbreviations............................. v Low-Pressure Fuel Turbopump............................ 56 Shuttle Propulsion System................................. 2 HPOTP Pump Section............................................ 60 SSME Introduction............................................... 4 HPOTP Turbine Section......................................... 62 SSME Highlights................................................... 6 HPOTP Shaft Seals................................................. 64 Gimbal Bearing.................................................... 10 HPFTP Pump Section............................................ 68 Flexible Joints...................................................... 14 HPFTP Turbine Section......................................... 70 Powerhead...........................................................
    [Show full text]
  • Experimental Study of the Combustion Efficiency in Multi-Element Gas
    energies Article Experimental Study of the Combustion Efficiency in Multi-Element Gas-Centered Swirl Coaxial Injectors Seongphil Woo 1,* , Jungho Lee 1,2, Yeoungmin Han 2 and Youngbin Yoon 1,3 1 Department of Aerospace Engineering, Seoul National University, Seoul 08826, Korea; [email protected] (J.L.); [email protected] (Y.Y.) 2 Korea Aerospace Research Institute, Daejeon 34133, Korea; [email protected] 3 Institute of Advanced Aerospace Technology, Seoul National University, Seoul 08826, Korea * Correspondence: [email protected] Received: 27 October 2020; Accepted: 16 November 2020; Published: 19 November 2020 Abstract: The effects of the momentum-flux ratio of propellant upon the combustion efficiency of a gas-centered-swirl-coaxial (GCSC) injector used in the combustion chamber of a full-scale 9-tonf staged-combustion-cycle engine were studied experimentally. In the combustion experiment, liquid oxygen was used as an oxidizer, and kerosene was used as fuel. The liquid oxygen and kerosene burned in the preburner drive the turbine of the turbopump under the oxidizer-rich hot-gas condition before flowing into the GCSC injector of the combustion chamber. The oxidizer-rich hot gas is mixed with liquid kerosene passed through combustion chamber’s cooling channel at the injector outlet. This mixture has a dimensionless momentum-flux ratio that depends upon the dispensing speed of the two fluids. Combustion tests were performed under varying mixture ratios and combustion pressures for different injector shapes and numbers of injectors, and the characteristic velocities and performance efficiencies of the combustion were compared. It was found that, for 61 gas-centered swirl-coaxial injectors, as the moment flux ratio increased from 9 to 23, the combustion-characteristic velocity increased linearly and the performance efficiency increased from 0.904 to 0.938.
    [Show full text]
  • Materials for Liquid Propulsion Systems
    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. Liquid propellant engines offer higher performance, that is, they deliver greater thrust per unit weight of propellant burned.
    [Show full text]
  • Lox/Lch4 Upper Stage Development Strategies for Future Launchers
    TURBO, vol. VI (2019), no. 2 LOX/LCH4 UPPER STAGE DEVELOPMENT STRATEGIES FOR FUTURE LAUNCHERS Theodora ANDREESCU1, Andreea MANGRA1, Valeriu VILAG1, Ion MALAEL1, Alexandru CANCESC1, Jeni VILAG1, Dan IFRIM1, Simona DANESCU1 ABSTRACT: The reduction of Earth-to-orbit launch costs in conjunction with an increase in Launcher reliability and operational efficiency are the key requirements of future space transportation systems. This paper underlines the progress in LOX/CH4 upper stage engine development carried out by COMOTI and also being provided the prediction of the rocket engine performances at the conceptual and preliminary stages of design. This paper focuses on the trade-off studies for the engine architecture definition, considering both open and closed thermodynamic cycles. Various subsystems configurations have been taken into account, analyzing the optimum configuration in terms of performance. The main operating and geometrical parameters were discussed: combustion pressure, optimum mixture ratio, turbine pressure ratio, thrust chamber geometry, and the turbopump size is addressed. KEYWORDS: upper stage, liquid rocket engine, turbopump, LOX/CH4 cryogenic propellant, thrust chamber NOMENCLATURE A – turbine characteristic area 풎̇ – gas flow rate 푨풆 - nozzle exit diameter 푷풑풐풙 – liquid oxygen pump power 푨풕 – nozzle throat diameter 푷풑풇풖풆풍 – liquid methane pump power 푨풄 – combustion chamber area 푷푻 – turbine power 푻풄 – combustion temperature 풏풔 – pump specific rotational speed 휸 - specific heat ratio 휶 – turbine flow angle ∗ 푪풑- specific heat capacity 풑 - turbine pressure losses 풑풄 – combustion pressure R – gas constant 푪푭 – thrust coefficient – density 풄∗ - characteristic velocity M – Mach number 품ퟎ – gravitational constant 풉풑풇, 풉풑풎 – turbine blades hight 풙풑풇, 풙풑풎, - axial width 1. INTRODUCTION As the present trend in rocket engine development recommends a high versatility and low launch service cost, while preserving high performance, expander cycle upper stage based on LOX/LCH4 being a key competitiveness factor recognized by the market.
    [Show full text]
  • IAF Committee Briefs
    IAF Committee Briefs July 2021 IAF SPACE PROPULSION TECHNICAL COMMITTEE 1. Introduction/Summary The Space Propulsion Committee addresses sub-orbital, earth to orbit, and in-space propulsion. The general areas considered include both chemical and non-chemical rocket propulsion, and air-breathing propulsion. Typical specific propulsion categories of interest are liquid, solid and hybrid rocket systems, electric, nuclear, solar and other advanced rocket systems, ramjet, scramjet, and various combinations of air-breathing and rocket propulsion. (Copyright SpaceX) The Committee deals with component technologies, propulsion system aspects, the implementation and In China, a project of heavy-lift launch vehicle, for future application of overall propulsion systems and dedicated important missions such as deep-space exploration and test facilities. The Committee is also examining the manned landing on the Moon, has been carried out. A feasibility of new missions made possible by new 500 ton- thrust-class LOX/kerosene rocket engine will propulsion systems and how combinations of pro- be used for the first stage and boosters of the launch pulsion technologies, such as chemical and electrical vehicle, which is also the next generation of large thrust technologies, can be optimized for this purpose. rocket engine in China. The engine employs oxidizer-rich staged-combustion cycle system and after-pump gimbal 2. Latest Developments configuration. The first engine has been manufactured and the sub-systems have been successfully hot-fire Right now, all over the world, a lot of new launchers tested. are appearing with new engines, with first flights just completed or planned for this year or in the next few On the coming first flights, we can underline the arrival years.
    [Show full text]
  • Up of Main Liquid Propellant Engine Under Hot Test Conditions
    DOI: 10.13009/EUCASS2019-236 8TH EUROPEAN CONFERENCE FOR AERONAUTICS AND SPACE SCIENCES (EUCASS) Mathematical modeling of transient processes during start- up of main liquid propellant engine under hot test conditions Dmytro Koptilyy, Roman Marchan, Sergey Dolgopolov and Olexiy Nikolayev FireFly Aerospace Ukraine, 103A, Gagarina Avenue, Dnipro 49050, Ukraine [email protected], [email protected], [email protected], aleksey.nikolaev@fireflyaerospace. Abstract Predicting parameters of transient processes before static fire tests of liquid rocket engines (LRE), issuing recommendations to ensure the engine performance, stability and dynamic strength of its structure, monitoring and analyzing technical condition of the engine during tests, are important to be addressed before conducting static fire tests of main LRE. An advanced approach to the mathematical modeling of transient processes in aggregates and systems of the main rocket engine (similar to 2000 kN RD 191, liquid oxygen and kerosene) with oxidizer staged combustion is proposed. The development of a mathematical model of nonlinear dynamics of rocket engines is based on the results of long-term research of low-frequency dynamics of cavitating pumps, injection of oxidizer gas products into the flow of liquid oxygen, the Turbo Pump Unit (TPU) dynamic interaction with oscillatory processes in gas manifolds, and dynamic processes in flow regulators and others. Applied to the LRE bench tests conditions the analysis of transient processes
    [Show full text]
  • RWTH Aachen University Institute of Jet Propulsion and Turbomachinery
    RWTH Aachen University Institute of Jet Propulsion and Turbomachinery DLR - German Aerospace Center Institute of Space Propulsion Optimization of a Reusable Launch Vehicle using Genetic Algorithms A thesis submitted for the degree of M.Sc. Aerospace Engineering by Simon Jentzsch June 20, 2020 Advisor: Felix Wrede, M.Sc. External Advisors: Kai Dresia, M.Sc. Dr. Günther Waxenegger-Wilfing Examiner: Prof. Dr. Michael Oschwald Statutory Declaration in Lieu of an Oath I hereby declare in lieu of an oath that I have completed the present Master the- sis entitled ‘Optimization of a Reusable Launch Vehicle using Genetic Algorithms’ independently and without illegitimate assistance from third parties. I have used no other than the specified sources and aids. In case that the thesis is additionally submitted in an electronic format, I declare that the written and electronic versions are fully identical. The thesis has not been submitted to any examination body in this, or similar, form. City, Date, Signature Abstract SpaceX has demonstrated that reusing the first stage of a rocket implies a significant cost reduction potential. In order to maximize cost savings, the identification of optimum rocket configurations is of paramount importance. Yet, the complexity of launch systems, which is further increased by the requirement of a vertical landing reusable first stage, impedes the prediction of launch vehicle characteristics. Therefore, in this thesis, a multidisciplinary system design optimization approach is applied to develop an optimization platform which is able to model a reusable launch vehicle with a large variety of variables and to optimize it according to a predefined launch mission and optimization objective.
    [Show full text]
  • Implications of Cycle Variants, Propellant Combinations and Operating Regimes on Fatigue Life Expectancies of Liquid Rocket Engines
    DOI: 10.13009/EUCASS2017-69 7TH EUROPEAN CONFERENCE FOR AERONAUTICS AND AEROSPACE SCIENCES (EUCASS) DOI: 069 Implications of Cycle Variants, Propellant Combinations and Operating Regimes on Fatigue Life Expectancies of Liquid Rocket Engines Günther Waxenegger?, Jörg Riccius?, Evgeny Zametaev?, Jan Deeken? and Julia Sand? ?German Aerospace Center (DLR), Germany Im Langen Grund, 74239 Hardthausen am Kocher, Germany [email protected] Abstract The paper in hand thematises a fatigue life analysis methodology for liquid rocket engines, with maiden fo- cus on the combustion chamber. The method of analysis combines the engine cycle modelling tool Ecosim- Pro/ESPSS with 2-d Finite Element analysis methods developed by DLR’s Institute of Space Propulsion. From the starting point of a generic LOx/LH2 engine three objectives are pursued, changing the power cycle, the propellant combination to LOx/LCH4 and lowering the chamber pressure. The results of the trade-off study focusing on the fatigue life expectancy are presented in detail. 1. Introduction The (partial) reusability of launch systems provides a possibility of significantly reducing the costs associated with rocket launches. Currently technically feasible concepts of reusability focus on reusable booster stages. With the Falcon 9 launch vehicles, SpaceX demonstrated its ability to successfully return the first stages of orbital launchers. Furthermore, SpaceX conducted a re-flight of a recovered first stage in March 2017 and once again landed the booster. In order to bring down the costs, it is absolutely necessary to reduce the required refurbishment to a minimum. This is particularly true of the engines which are responsible for a large fraction of the total stage cost.4 To make things worse the propulsion system belongs to the most critical subsystems of a launch system.7 This is driven by the extreme operational load conditions to which the system is exposed during operation.
    [Show full text]
  • Incorporation of RD-180 Failure Response 2011
    Incorporation of RD-180 Failure Response Features in the Atlas V Booster Emergency Detection System Felix Y. Chelkis NPO Energomash Khimki, Moscow Region, Russia Thaddeus D. Chwiedor United Launch Alliance, Denver, Colorado, USA Paul T. Connolly and John P. Hansen Pratt & Whitney Rocketdyne, Jupiter, Florida, USA Abstract Approximately 70% of the parts on the RD-180 engine as used on Atlas V launch vehicles are common with the RD-170 engine. The RD-170 engine was developed by NPO Energomash and certified in 1985 to power the Energia Buran rocket launch system, which was rated for human spaceflight as defined by the Russian standards. With over 45,000 seconds of operational time, the RD-180 has an extensive history of demonstrated performance, robustness, and reliability as does the Atlas vehicle that it powers. This paper describes why the RD-180/Atlas combination provides an optimal solution for human space flight directly applicable to the near-term commercial crew transportation program. The paper will examine how the RD-180/Atlas booster combination could meet Commercial Crew space flight requirements. An overview of the past NASA human space flight requirements to the Russian approach for manned flight will also be made. The paper will discuss the work that has been done to understand not only the potential failure modes to which the RD-180 propulsion system could be subjected to but also how these failure modes could be quickly identified, mitigated or managed such that a safe crew abort sequence could be initiated. The RD-180/Atlas Emergency Detection System (EDS) will be described and shown to be a viable approach to meeting Commercial Crew space flight requirements.
    [Show full text]
  • Propulsion System Choices and Their Implications
    461h AIAAJAS M EJSAEJASEE Joint Propuls ion Conference AIAA 20 I O-XXXX 25 -28 July 20 10, Nashvi lle, Tc n n~scc Propulsion System Choices and Their Implications Claude R. Joyner II ' Prall & Whi tney Rocketdyne, West Palm Beach, Florida, 334 10 Daniel J. H. Levack' Prall & Whitney Rocke/dyne, Canoga Park, Californ ia, 91309 Russel E. Rhodes' NASA , Kennedy Space Center, Florida, 32899 John W. Robinson' Th e Boeing Company, Huntington Beach, California, 92647 Abstract In defining a space vehicle architecture, the propulsion system and related subsystem choices will have a major influence on achieving the goals and objectives desired. There are many alternatives and the choices made must produce a system that meets the performance requirements, but at the same time also provide the greatest opportunity of reaching all of the required objectives. Recognizing the above, the SPST Functional Requirements subteam has drawn on the knowledge, expertise, and experience of its members, to develop insight that wiIJ effectively aid the architectural concept developer in making the appropriate choices consistent with the architecture goals. This data not only identifies many selected choices, but also, more importantly, presents the collective assessment of this subteam on the "pros" and the "cons" of these choices. The propulsion system choices with their pros and cons are presented in five major groups. A. System Integration Approach. Focused on Ibe requirement for safety, reliability, dependability, maintainability, and lo\\" cost. B. Non-Chemical Propulsion. Focused on choice of propulsion type. C. Chemical Propulsion. Focused on propeUant choice implications. D. Functional Integration. Focused on the degree of integration of the many propulsive and closely associated functions, and on the choice of the engine combustion power cycle.
    [Show full text]