DOCSLIB.ORG

Development of P4.1 Altitude Simulation for Vinci Engine

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Development of P4.1 Altitude Simulation for Vinci Engine DEVELOPMENT OF P4.1 ALTITUDE SIMULATION FOR VINCI® ENGINE K. Schäfer, C. Böhm, H. Kronmüller, H. Zimmermann, German Aerospace Centre (DLR), Institute of Space Propulsion, Lampoldshausen, 74172 Hardthausen, Germany 1 ABSTRACT pean launcher programs and one of its main tasks has always been high altitude testing of To test rocket propulsion systems as much as rocket engines. possible under operational conditions are one of The P4 test facility with two test positions P4.1 the main drivers for flight qualification. For up- and P4.2 was build in the mid-sixties for the per stage engines the tests in vacuum conditions ELDO program. In 1973 when the ARIANE with fully expanded nozzle defines the essential launcher program started the P4 was equipped criteria for altitude simulation test facilities. for altitude simulation tests of the VIKING en- SNECMA is developing a new upper stage en- gine with 80 t vacuum thrust. ® gine called VINCI . For the engine tests the ex- With the development of the ARIANE 5 the isting test bench P4.1 at the DLR test centre P4.2 was adapted in 1992 to the altitude simula- Lampoldshausen was adapted to high altitude tion of the AESTUS upper stage engine. conditions. To maintain the experience of the test facility ® VINCI is an Expander cycle-type engine with engineering and to improve the altitude simula- up to 180 kN thrust. An extendable nozzle will tion especially the development of rocket steam ® give VINCI a specific impulse of 465 s in vac- generators the department of engineering was uum. founded in 1996 within the institute of space The task of altitude simulation consists of creat- propulsion. ing the test condition within a vacuum cell. This In 1998 the decision was taken to modify the is primarily low ambient pressure of just few existing test position P4.1 for the VINCI® de- mbar. Special operational conditions are linked velopment tests. Since February 2005 the test to the transients during Start-Up and Shut-Down bench is operational (FIG. 1). of the engine with respect to the nozzle loads. Maintenance of the vacuum with running engine is achieved by recompression of the supersonic exhaust jet in a diffuser and the extraction of the exhaust gas by steam jet ejectors. To provide the large quantities of steam, rocket steam generators with liquid Oxygen and Alco- hol are used. The maximum flexibility is reached by adapta- tion of the test bench to the different engine configurations and the test conditions. This con- tents modular systems. FIG. 1: P4.1 ALTITUDE SIMULATION 2 INTRODUCTION P4.1 is now the larges altitude simulation test From the very beginning in the 1960s, DLR facility in Europe. Lampoldshausen has been involved in all Euro- − Ballistic phase and reignition in vacuum 3 TEST REQUIREMENTS conditions. ® The VINCI engine (FIG. 2) will be tested in 3 4 DEVELOPMENT OF P4.1 ALTITUDE configurations (Table.1). SIMULATION Configuration The development of the altitude simulation was done in different phases: I Combustion Phase 1: Basic studies and general lay out. Chamber Phase 2: Development of basic components II Combustion − Centre Body Diffuser Chamber and − Rocket Steam Generators Fix Nozzle − Ejector System − Modular system to different test configura- III Complete tions. Engine with Phase 3: Final lay out of the test bench. Extendable Nozzle 5 STUDIES AND GENERAL LAY OUT (PHASE 1) FIG. 2: VINCI ENGINE The basic studies are performed to define the VINCI Test Expansion Ratio ε principal conditions and dimensions of the Configuration Length l [m] Exit Diameter Ø [m] components, particularly the super sonic dif- fuser and the dimensioning of the suction sys- Configuration Combustion chamber I ε = 22,3, l ∼ 1,4 m, Ø∼ 0,7 m tem with adapters for the different test configu- rations. Configuration Chamber with the fix nozzle II ε =93, l ∼ 2,2 m, Ø∼ 1,4 m Centre Body Diffuser: Due to the limited high Configuration Complete engine with the extendable of the existing P4.1 building a centre body dif- III nozzle fuser was chosen with reduced length. ε = 243, l ∼ 4,2 m, Ø∼ 2,3 m ® The centre body diffuser has the same behaviour TABLE 1: VINCI TEST CONFIGURATIONS like a second throat diffuser. The supersonic flow is stable down to lower pressures ratios The P4.1 test conditions are: (Hysteresis). The length L to diameter D ratio − Vertical test position with maximum test should be 6 < L/D < 8 for reduced pressure time of 770 s. losses. The second throat is realised by a ring − Pressure at engine interface during chill channel around a centre body. The overall down p < 200 mbar. length of the diffuser is short because of the re- − Ignition at ambient pressure p < 60 mbar duced hydraulically diameter of a ring channel. The centre body has to be cooled intensively. − Start up phase with simulation of in-flight engine start up conditions The main parameters for the centre body dif- fuser design were linked to the characteristics of − Steady state phase with the operational enve- the combustion chamber pressure Pc, the ex- lope in vacuum conditions pandable nozzle with the expansion ratio ε = − Shut down phase with transient conditions 243, the heat loads and the given dimensions of considering the maximum nozzle loads P4.1. Suction system: The main parameters for the suction system were linked to the transient pres- Phase 1 Procedure sure conditions during Start-Up and Shut-Down Studies and General Layout of the engine. The trade-off was between pow- P4.1 Altitude Simulation erful ejectors with high steam consumption for fast transients and a big condenser with high Basic Studies cooling water flow. Adapters: The adapter for combustion chamber tests is designed as a short diffuser and for Gim- balling. The adapter for the fix nozzle replaces Step 1 Geometrical Data the extendable nozzle and guides the exhaust to the centre body diffuser. Main drivers were: Design of Interface Components Adjustments − Consumption of steam for the sizing of the rocket steam generator plant. − Ignition and reignition conditions. Step 2 − Consumption of cooling water for the sizing Cross Check of the cooling water plant, the underground storage and the condenser. First Lay Out Check of − The transient behaviour of the engine re- Altitude Simulation Characteristics quires the supersonic Start and Un-Start of the diffuser at low combustion chamber pres- sures to prevent the back flow of hot gases. Step 3 Check Reliability − Flexibility to adapt the altitude simulation to the test configurations by short notice. − Reliability of the altitude simulation Design Check of Altitude Simulation Sensitivity/Margins The altitude simulation has been achieved by an experienced calculation process. The general General Design way of calculation (FIG. 3) consists of three Altitude Simulation steps. FIG 3: CALCULATION PROCESS Step 1: A forward calculation, to determine geometric parameters from input parameters, Test Phase Combustion Vacuum represents the first step. pressure [bar] chamber [mbar] Step 2: The second step is like a cross-check Chill down - < 40 procedure by a backward calculation, receiving (< 200 at chill down the characteristic curve from geometric parame- Interface) ters. Ignition - 25 - 40 Step 3: Finally in a third step, the sensitivity to Start 10 – 15 < 60 parameter-variations is evaluated to check the Diffuser reliability of the prediction. Steady 25 – 80 2,5 - 7 To get the required accuracy, the steps are ad- State justed by experienced loss- and gain- UN Start 10 – 20 < 80 parameters, which depends on type and basic Diffuser configuration. The result arises iteratively by TABLE 2: PARAMETERS ε = 243 CONFIGURATION this way. LOX buffer LH2 buffer LOX runtank vacuum chamber P4.1 ejectors N2 / He panels VINCI engine P4.1 condenser center body diffuser P4.2 condenser deflector chill down ejectors sea level configuration vacuum flap first ejector stage steam lines underground 0 1 5 10 m water reservoir FIG 4: GENERAL DESIGN P4.1 The general design (FIG. 4) is: down against higher pressure of the con- − Feed System with run tanks for LH2 and denser. LOX designed for chill down and 770 s hot − A big condenser reduces the mass flows to be run. sucked by the end stage ejector to ambient − Buffer tanks for LH2 and LOX to simulate pressure conditions. The water is routed the dynamical transient feeding conditions of barometrically to an underground water res- the stage during Start up and Shut Down ervoir maintaining the vacuum pressure in- side the condenser. − The VINCI engine is vertically mounted in- side a vacuum chamber within a thrust meas- − The “chill down” ejectors maintain the vac- urement device. uum simulation during the cool down phase of the engine before ignition. − The supersonic exhaust jet will be recom- pressed inside the centre body diffuser to subsonic conditions. 6 DIFFUSER SUBSCALE TESTING − The heat loads of the hot gases will be re- (PHASE 2) duced by water injection after the diffuser. All parts up to the vacuum flap are double CFD calculations and model tests under cold wall water cooled. The exhaust gases are and hot conditions were performed to verify the guided to the first ejector stage by the deflec- heat loads and functional behaviour of the cen- tor. tre body diffuser. − The first ejector stage maintains the pressure conditions during Ignition, Start up and Shut 6.1 SUBSCALE DIFFUSER COLD GAS Different diffusers were tested for basic investi- TESTS gations (fig. 8). The results show a supersonic Start and Un-Start of the centre body diffuser at The test position P6.2 (fig. 5) has been devel- low ratios of chamber pressure to vacuum pres- oped and erected to investigate components for sure.
Load more

Recommended publications
  • Qualification Over Ariane's Lifetime
    r bulletin 94 — may 1998 Qualification Over Ariane’s Lifetime A. González Blázquez Directorate of Launchers, ESA, Paris M. Eymard Groupe Programme CNES/Arianespace, Evry, France Introduction Similarly, the RL10 engine on the Centaur stage The primary objectives of the qualification of the Atlas launcher has been the subject of an activities performed during the operational ongoing improvement programme. About 5000 lifetime of a launcher are: tests were performed before the first flight, and – to verify the qualification status of the vehicle 4000 during the subsequent ten years. – to resolve any technical problems relating to subsystem operations on the ground or in On-going qualification activities of a similar flight. nature were started for the Ariane-3 and 4 launchers in 1986, and for Ariane-5 in 1996. Before focussing on the European family of They can be classified into two main launchers, it is perhaps informative to review categories: ‘regular’ and ‘one-off’. just one or two of the US efforts in the area of solid and liquid propulsion in order to put the Ariane-3/4 accompanying activities Ariane-related activities into context. Regular activities These activities are mainly devoted to In principle, the development programme for a launcher ends with the verification of the qualification status of the qualification phase, after which it enters operational service. In various launcher subsystems. They include the practice, however, the assessment of a launcher’s reliability is a following work packages: continuing process and qualification-type activities proceed, as an – Periodic sampling of engines: one HM7 and extension of the development programme (as is done in aeronautics), one Viking per year, tested to the limits of the over the course of the vehicle’s lifetime.
    [Show full text]
  • High-Thrust In-Space Liquid Propulsion Stage: Storable Propellants
    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 – ID 2968378 High-Thrust in-Space Liquid Propulsion Stage: Storable Propellants Etienne Dumont, Alexander Kopp, Carina Ludwig, Nicole Garbers DLR, Space Launcher Systems Analysis (SART), Bremen, Germany [email protected] Abstract In the frame of a project funded by ESA, a consortium led Subscripts, Abbreviations by Avio in cooperation with Snecma, Cira, and DLR is ATV Automated Transfer Vehicle performing the preliminary design of a High-Thrust in- CDF Concurrent Design Facility Space Liquid Propulsion Stage for two different types of ECSS European Cooperation on Space manned missions beyond Earth orbit. For these missions, Standardization one or two 100 ton stages are to be used to propel a Elec assy electronic assembly manned vehicle. Three different propellant combinations; EPS Etage à propergols stockables (Ariane 5’s LOx/LH2, LOx/CH4 and MON-3/MMH are being storable propellant stage) compared. GNC Guidance Navigation and Control HTS High-Thrust Stage The preliminary design of the storable variant (MON- IF Interface 3/MMH) has been performed by DLR. The Aestus II ISS International Space Station engine with a large nozzle expansion ratio has been LEO Low Earth Orbit chosen as baseline. A first iteration has demonstrated, that LEOP Launch and Early Orbit Phase it indeed provides the best performance for the storable LH2 Liquid Hydrogen propellant combination, when considering all engines LOx Liquid Oxygen available today or which may be available in a short- to MLI Multi-Layer Insulation medium term.
    [Show full text]
  • 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]
  • Variations of Solid Rocket Motor Preliminary Design for Small TSTO Launcher
    View metadata, citation and similar papers at core.ac.uk brought to you by CORE provided by Institute of Transport Research:Publications Space Propulsion 2012 – ID 2394102 Variations of Solid Rocket Motor Preliminary Design for Small TSTO launcher Etienne Dumont Space Launcher Systems Analysis (SART), DLR, Bremen, Germany [email protected] NGL New/Next Generation Launcher Abstract SI Structural Index (mdry / mpropellant) Several combinations of solid rocket motors and ignition SRM Solid Rocket Motor strategies have been considered for a small Two Stage to TSTO Two Stage To Orbit Orbit (TSTO) launch vehicle based on a big solid rocket US Upper Stage motor first stage and cryogenic upper stage propelled by VENUS Vega New Upper Stage the Vinci engine. In order to reach the target payload avg average during the flight performance of about 1400 kg into GTO for the clean s.l. sea level version and 2700 to 3000 kg for the boosted version, the vac vacuum influence of the selected solid rocket motors on the upper 2 + 2 P23 4 P23: two ignited on ground and two with a stage structure has been studied. Preliminary structural delayed ignition designs have been performed and the thrust histories of the solid rocket motor have been tweaked to limit the upper stage structural mass. First stage and booster 1. Introduction combinations with acceptable general loads are proposed. Solid rocket motors (SRM) are commonly used for boosters or launcher first stage. Indeed they can provide high thrust levels while being compact, light and Nomenclature relatively simple compared to a liquid rocket engine Isp specific impulse s providing the same thrust level.
    [Show full text]
  • Los Motores Aeroespaciales, A-Z
    Sponsored by L’Aeroteca - BARCELONA ISBN 978-84-608-7523-9 < aeroteca.com > Depósito Legal B 9066-2016 Título: Los Motores Aeroespaciales A-Z. © Parte/Vers: 1/12 Página: 1 Autor: Ricardo Miguel Vidal Edición 2018-V12 = Rev. 01 Los Motores Aeroespaciales, A-Z (The Aerospace En- gines, A-Z) Versión 12 2018 por Ricardo Miguel Vidal * * * -MOTOR: Máquina que transforma en movimiento la energía que recibe. (sea química, eléctrica, vapor...) Sponsored by L’Aeroteca - BARCELONA ISBN 978-84-608-7523-9 Este facsímil es < aeroteca.com > Depósito Legal B 9066-2016 ORIGINAL si la Título: Los Motores Aeroespaciales A-Z. © página anterior tiene Parte/Vers: 1/12 Página: 2 el sello con tinta Autor: Ricardo Miguel Vidal VERDE Edición: 2018-V12 = Rev. 01 Presentación de la edición 2018-V12 (Incluye todas las anteriores versiones y sus Apéndices) La edición 2003 era una publicación en partes que se archiva en Binders por el propio lector (2,3,4 anillas, etc), anchos o estrechos y del color que desease durante el acopio parcial de la edición. Se entregaba por grupos de hojas impresas a una cara (edición 2003), a incluir en los Binders (archivadores). Cada hoja era sustituíble en el futuro si aparecía una nueva misma hoja ampliada o corregida. Este sistema de anillas admitia nuevas páginas con información adicional. Una hoja con adhesivos para portada y lomo identifi caba cada volumen provisional. Las tapas defi nitivas fueron metálicas, y se entregaraban con el 4 º volumen. O con la publicación completa desde el año 2005 en adelante. -Las Publicaciones -parcial y completa- están protegidas legalmente y mediante un sello de tinta especial color VERDE se identifi can los originales.
    [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]
  • Aestus Engine
    ©ArianeGroup AESTUS ENGINE PROPULSION SOLUTIONS FOR LAUNCHERS POWERS THE ARIANE 5 ES UPPER STAGE * LEO: LOW-EARTH ORBIT SSO: SUN-SYNCHRONOUS ORBIT ORBITAL INSERTION OF HEAVY PAYLOADS INTO LEO, SSO AND GTO* GTO: GEOSTATIONARY TRANSFER ORBIT PRESSURE-FED ENGINE CONSUMING UP TO 10 METRIC TONS ** MMH/N204: OF BIPROPELLANT MMH/N204** MONOMETHYLHYDRAZINE/ DINITROGEN TETROXIDE PROVEN DESIGN AND FLEXIBILITY WITH MULTIPLE RE-IGNITION CAPABILITY TO PLACE 21-METRIC TON ATV INTO LEO OPERATIONAL AS OF 1997 (ARIANE FLIGHT 502) TO JULY 2018 (ARIANE FLIGHT 244) AESTUS ENGINE SPACE PROPULSION Aestus MAIN CHARACTERISTICS Propellants N2O4\ development MMH history Specific impulse vacuum 324 s Thrust vacuum 29.6 kN Propellant mass flow rate 9.3 kg/s 1988–1995: Development at the Mixture ratio (TC) 1.9 Ottobrunn Space Propulsion Centre in Germany Engine feed pressure 17.7 bar Combustion chamber 11 bar 1999–2002: Performance improvement pressure program involving propellant mixture ratio adjustment Nozzle area ratio 84 Nozzle exit diameter 1.31 m 2003–2007: Re-ignition qualification program demonstrated with the first Overall engine length 2.2 m ATV launch (9 March 2008) Thrust chamber mass 111 kg 2009–2015: Specific delta qualification Nominal single firing 1100 s for ES Galileo missions including Power 43,700 kW production restart 59,400 hp Re-ignition capability Multiple MAJOR SUB-ASSEMBLIES Injector with coaxial injection elements for mixing propellants Combustion chamber regeneratively cooled by MMH fuel Nozzle extension, radiatively cooled Propellant valves for fuel and oxidiser, pneumatically operated by pilot valves Gimbal joint mounted at the top of the injector dome allowing for pitch and yaw control The Aestus engine powers the Ariane 5 ES version bipropellant upper stage for insertion of payloads into LEO, SSO and GTO.
    [Show full text]
  • Adaptation of the Espss/Ecosimpro Platform for the Design and Analysis of Liquid Propellant Rocket Engines J
    ADAPTATION OF THE ESPSS/ECOSIMPRO PLATFORM FOR THE DESIGN AND ANALYSIS OF LIQUID PROPELLANT ROCKET ENGINES J. Amer(1), J. Moral(2), J.J. Salvá(3) (1) Final thesis. Departamento de Motopropulsion y Termofluidodinámica, E.T.S. Ingenieros Aeronáuticos, Universidad Politécnica de Madrid (email: [email protected]) (2) Empresarios Agrupados Internacional. S.A. Magallanes, 3. 28015 Madrid. Spain. (email: [email protected]) (3) Departamento de Motopropulsion y Termofluidodinámica, E.T.S. Ingenieros Aeronáuticos, Universidad Politécnica de Madrid. Pza. del Cardenal Cisneros, 3. 28040 Madrid. Spain (email: [email protected]) Abstract • Design parameters: chamber pressure (pc), mixture The preliminary design of rocket engines has been ratio (MR), and expansion area ratio (εe) historically a semi-manual process, where the • Propulsion characteristics (Isp, c*, cF) specialists work out a start point for the next design • Sizing and mass (throat diameter, thrust, propellants steps. Thanks to the capabilities of the EcosimPro mass, etc.) language, such as non causal programming, a natural pre-design model has been physically discomposed in three modules: propulsion performance, sizing and mass, and mission requirements. Most of the new stationary capabilities are based on the ESPSS libraries, which are able to simulate in great detail transient phenomena of fluid systems (e.g. tanks, turbomachinery, nozzles and combustion chambers). Once the base of the three pre-design modules is detailed, an effort has been made to determine accurately the mission requirements. For example, in a geostationary transfer orbit insertion maneuver, the maximum change of speed is obtained considering a finite combustion, instead of the ideal Hohmann transfer orbit. Finally, the model validation and two application examples are presented: the first one compares the model results with the real performance of the Aestus pressurized rocket engine.
    [Show full text]
  • Jarvis Heavy Launch Vehicle
    JARVIS HEAVY LAUNCH VEHICLE By Forum Orbiter Italia Version 2.62 – October 2012 USER MANUAL Disclaimer and credits This add-on is provided “as is”, without any kind of warranty; it is compatible with Orbiter 2006-P1 (build 060929) and with Orbiter 2010-P1 (build 100830). Many thanks to Dr. Martin Schweiger, for the Orbiter Space Simulator. For the others developers: You are free to use parts of our work, eg sound and texture, but you must credit us as the original source of your work. FOI Credits - Andrew: add-on conception; rocket textures, meshes and configuration; documentation editing. - Fausto: new launch pad textures, meshes and configuration. - Pete Conrad: engine meshes and textures; Shuttle SRB meshes and textures; “dummy” payload meshes and textures. - FedeX: beta testing. - Dany: “Forum Orbiter Italia” logo. - Ripley: D3D9/D3D11 documentation. Forum Orbiter Italia: http://orbiteritalia.forumotion.com/ Introduction In the mid-eighties, the "Jarvis" project was the last serious attempt to revive the glorious Saturn V rocket, and at the same time, one of the first ideas of an alternative use for the Space Shuttle hardware, many years before the current "Ares", "Direct" and “SLS” projects. The Jarvis rocket combines the powerful Apollo-era F-1 and J-2 engines with Space Shuttle electronics and 8.4 m stages (the same size of the Shuttle External Tank). Later versions, with Space Shuttle Main Engines (SSME) and/or Solid Rocket Boosters (SRB), were proposed, but never realized. Forum Orbiter Italia has developed a complete and versatile family of heavy launchers around these original ideas and projects.
    [Show full text]
  • A Survey of Automatic Control Methods for Liquid-Propellant Rocket Engines
    A survey of automatic control methods for liquid-propellant rocket engines Sergio Pérez-Roca, Julien Marzat, Hélène Piet-Lahanier, Nicolas Langlois, Francois Farago, Marco Galeotta, Serge Le Gonidec To cite this version: Sergio Pérez-Roca, Julien Marzat, Hélène Piet-Lahanier, Nicolas Langlois, Francois Farago, et al.. A survey of automatic control methods for liquid-propellant rocket engines. Progress in Aerospace Sciences, Elsevier, 2019, pp.1-22. 10.1016/j.paerosci.2019.03.002. hal-02097829 HAL Id: hal-02097829 https://hal.archives-ouvertes.fr/hal-02097829 Submitted on 12 Apr 2019 HAL is a multi-disciplinary open access L’archive ouverte pluridisciplinaire HAL, est archive for the deposit and dissemination of sci- destinée au dépôt et à la diffusion de documents entific research documents, whether they are pub- scientifiques de niveau recherche, publiés ou non, lished or not. The documents may come from émanant des établissements d’enseignement et de teaching and research institutions in France or recherche français ou étrangers, des laboratoires abroad, or from public or private research centers. publics ou privés. A survey of automatic control methods for liquid-propellant rocket engines Sergio Perez-Roca´ a,c,∗, Julien Marzata,Hel´ ene` Piet-Lahaniera, Nicolas Langloisb, Franc¸ois Faragoc, Marco Galeottac, Serge Le Gonidecd aDTIS, ONERA, Universit´eParis-Saclay, Chemin de la Huniere, 91123 Palaiseau, France bNormandie Universit´e,UNIROUEN, ESIGELEC, IRSEEM, Rouen, France cCNES - Direction des Lanceurs, 52 Rue Jacques Hillairet, 75612 Paris, France dArianeGroup SAS, Forˆetde Vernon, 27208 Vernon, France Abstract The main purpose of this survey paper is to review the field of convergence between the liquid-propellant rocket- propulsion and automatic-control disciplines.
    [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]