DOCSLIB.ORG

Chamber Valve Characteristic

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text

Load more

Appendix A Chamber Valve Characteristic The position of a valve can be read by the MCC system from end switches; they indicate the full open or close status (Fig. A.1). In a valve switch test characteristic time delays can be identified. Open Close 250 256 371 248 Appears Appears Disappears Disappears Open Closed Open Closing Command Closed Fig. A.1 Time delay in Opening Command milliseconds for the manoeuvring of a valve Feed Back Back Feed (chamber valve hydrogen of a Back Feed Feed Back Back Feed Vulcain engine) (Photo: DLR) Back Feed W. Kitsche, Operation of a Cryogenic Rocket Engine, Springer Aerospace Technology, 95 DOI 10.1007/978-3-642-10565-4, C Springer-Verlag Berlin Heidelberg 2011 Appendix B Chamber Igniter Characteristic The chamber igniter is itself ignited by a pill of powder at its tip. A wire in a short- circuit conducting a current of 6 A for 50 ms ignites the pill. Thus the main powder charge is ignited and the internal pressure (Fig. B.1) of the pyrotechnical element increases. At a pressure of about 120 bar, three membranes which had sealed the element burst. Via three openings at an angle of 120◦ to each other the hot smoke is injected into the combustion chamber and ignites the fuel/oxidiser mixture. The time taken to reach 80% of the maximum pressure inside the igniter in this context is called the ignition delay (Fig. B.1). Figure B.2 is a series of photos showing the ignition of the fuel/oxidiser mixture and the flame evolution in the combustion chamber. The photos were taken with a high speed camera (2000 frames per second) via a mirror below the engine. 140 120 100 80 60 Pressure [bar] Pressure 40 20 0 0 100 200 300 400 500 600 700 Time [ms] Fig. B.1 Ignition delay of a chamber igniter (Photo: DLR) Fig. B.2 Ignition of a Vulcain-combustion chamber (Photo: DLR) 97 Appendix C Measurement and Correction of the Ovality of a Vulcain Nozzle The measurement and correction of the ovality of the Vulcain nozzle has been per- formed regularly since 1993. The protocol (Fig. C.1) documents such a procedure. The diameter was measured at different circumferential positions (before and after) and with a special clamping device the nozzle exit was reformed into a shape close to the circle. Fig. C.1 Inspection sheet (Inspection sheet: VOLVO) 99 Appendix D Compare of Flow Schemes Vulcain/Vulcain 2 Fig. D.1 Compare of Flow Schemes Vulcain/Vulcain 2 (Photo: SNECMA) 101 Appendix E Flow Scheme of the Ariane 5 Main Stage Fig. E.1 Flow scheme of the Ariane 5 main stage (Photo: SNECMA/EADS (formerly SEP/ Aerospatiale)) 103 Flow Scheme of Appendix F LAO + EVSCP (conditionnement GHe) FEFP (évacuation huile GAM) BMO (balayage TPO) LAO (alim.LOX) PGD (GHe commande) ECCH (conditionnement GHe) LAH (conditionnement GHe) PPH (Pressurisation RLH2) EVAE (alimentation LHe) TPCO (purge LOX) CTPCH (chasse ergols) BMH (balayage hydrogène) PPO (pressu. RLOX) EVSCP (purge SCP) ETAGE TPCH (purge LH2) LAH (alim.LH2) F F F CBPTM1 F B E P F FBSH P CBOR CBPTM2 FPOE FCE FPHE S P R FEC C H O 1 O V U L C A I N R E LEFP1 R 1 VCBPTM TP LH2 TP LOX DM CBCER VGC CBJTOS LTO Ce documentestlapropriétédeSNECMA.Ilnepeutêtrereproduitoucommuniquésanssonautorisationpréalabl LTH CBJPH CBJTO CBJTH Vulcain 2 LPRO CEVC VMRH CBGO1 GG L CBGH r F AG LEP u e P f f m u O SFAVGO u a i l SFAVGH h F é CBGO2 c h é P EVVGO R EVVGC O M CAVGO V V CAVGH EVVGH EVBO G G (BCO + BGO) VPO VPH CBCO2 O H FPHM EVVCO L L EVVP G G (VPO + VPH) O H LMRH EVVCH (VCH+VMRH) SFAVCO CBCO1 BEVC BEVB CBH LCO AC SFAVCH VCO CBCH CP LEO LCH2 LCH1 L F H VCH e etécrite LEH SYNOPTIQUE VULCAIN 2 Hydrogène liquide Gaz chaud + GHe Huile GAM Schéma fonctionnel fluides Diffusion provisoire Hydrogène gazeux GOX+GHe Volume conditionné pour information, en MOTEUR 105 présenté en phase propulsée Oxygène liquide Hélium gazeux Conditionnement attente d'une édition LEFP2 Mise à jour : 119901 du document officiel EPC A5E Gaz chaud Hélium liquide Matériels Fig. F.1 Flow scheme of Vulcain 2 (Photo: SNECMA/EADS (formerly SEP/Aerospatiale)) Appendix G Flow Scheme of Vinci (Photo: SNECMA) Vinci Flow scheme of Fig. G.1 107 Appendix H Mixtures of Oxygen and Nitrogen Close to Their Boiling Points The main component of an oxygen (O2) supply to a rocket engine is the run tank with liquid O2. If the tank is pressurised with nitrogen (N2) the condensation of N2 at the surface of the liquid O2 has to be considered. The equilibrium of both fluids close to their boiling point is shown in the phase diagram (Fig. H.1). For the calculation of the dewpoint and boilingpoint curve (H.1) and (H.2), [16] were used; the vapour pressure Pi (T ) was calculated according to [21]. − ( ) = P P2 T Boiling curve X1 (H.1) P1(T ) − P2(T ) − ( ) ( ) ( ) = P1 T P2 T 1 + P1 T Dew curve X1 (H.2) P1(T ) − P2(T ) P P1(T ) − P2(T ) The symbols are: T temperature of the mixture (here the temperature of the liquid O2), P pressure of the mixture (here the pressure in the vessel), P1(T ) vapour pressure of the first component (here N2), P2(T ) vapour pressure of the second component (here O2), X liquid mole fraction of N2, 1 X1 gaseous mole fraction of N2. Pressure = 1 bar Pressure = 2 bar Pressure = 6 bar Temperature of the Mixture [ K] Temperature of the Mixture[ K] Temperature of the Mixture [ K] 100 100 120 Mixture of Gas Mixture of Gas 95 Mixture of Gas 95 115 Gas/Liquid Mixture 110 90 K Line 90 K Linie 90 90 105 Gas/Liquid Mixture Gas/Liquid Mixture 85 85 100 95 80 80 Mixture of Liquids 90 Mixture of Liquids Mixture of Liquids 90 K Line 75 75 85 0 0,2 0,4 0,6 0,8 1 0 0,2 0,4 0,6 0,8 1 0 0,2 0,4 0,6 0,8 1 Nitrogen Fraction of the Mixture Nitrogen Fraction of the Mixture Nitrogen Fraction of the Mixture Fig. H.1 Phase diagram of O2/N2 mixtures at 1, 2 and 6 bar (Photo: DLR) 109 110 Appendix H Mixtures of Oxygen and Nitrogen Close to Their Boiling Points In fact the mixture is not homogeneous in time and space and is not at equi- librium. The pressure is defined by the pressurisation system; above the liquid it is constant, below the surface it increases with depth. The temperature around the surface is dominated by the temperature of the liquid oxygen, normally close to 90 K. If the tank pressure is held at 1 bar for a long time only very small fractions of N2 can be found in the liquid. When the pressurisation is started the condensation at the surface begins and the N2 fraction increases. We can assume that the condensed N2 remains mainly in the upper layers of the liquid (see below). After test, particularly when the pressure is reduced to 1 bar, the N2 as the more volatile component evaporates again. The typical evaporation rate of a vacuum insulated tank is 0.1% of the inner volume per day. For the oxygen tank (200 m3) on the facility P5 that means 200 L per day. Hence it takes 7 days to re-evaporate the N2. But in the normal test cycle the tank is refilled the day after the test. Due to this dilution the concentration of N2 decreases to a tenth part, which is a far more drastic decrease than that due to the evaporation. Nevertheless, for the engine test the N2 fraction in the oxygen means pollution. A sample taken during a normal test period had a fraction of 0.36% of N2 in O2. Diffusion of Liquid N2 in Liquid O2 In order to check whether a fast diffusion can distribute the N2 fraction all over the tank, we have to look for the solution of the diffusion equation (H.3)atthe applicable conditions. We also need the coefficient of N2 diffusion in liquid O2. The diffusion equation has the same mathematical structure as the heat conduction equation (H.5) for which we find the general solution (H.6)in[16] and transfer it to the diffusion equation. For the diffusion coefficient D no measured value could be found in the literature or on the internet; therefore the Stokes-Einstein-equation (H.4) was applied. ∂ c ∂2c Diffusion equation = D (H.3) ∂ t ∂ xi ∂ x j where c is the concentration, t the time, and xi the position vector (or, more cor- rectly, the coefficients of the components of the position vector). The Stokes-Einstein equation gives us K T − D = B = 97.6 × 10 9 Ns/m2 (H.4) 6 πμR0 −23 where K B = 1.38 × 10 as the Boltzmann number, T = 90 K as temperature, −6 2 μ = 6.5 × 10 Ns/m as the dynamic viscosity (of O2 at 90 K, 6 bar) and R0 = −11 10.4 × 10 m as particle radius (of N2). We obtain the general solution for the concentration c(x, t) in the one dimen- sional case: Appendix H Mixtures of Oxygen and Nitrogen Close to Their Boiling Points 111 ( , ) = + + − 2 ( ( ) + ( )) c x t c0 c1x exp Dant An cos an x Bn sin an x (H.5) n with the constants c0, c1, an, An, Bn.
Recommended publications
  • Numerical Investigation of a 7-Element GOX/GCH4 Subscale Combustion Chamber

    Numerical Investigation of a 7-Element GOX/GCH4 Subscale Combustion Chamber

    DOI: 10.13009/EUCASS2017-173 7TH EUROPEAN CONFERENCE FOR AERONAUTICS AND AEROSPACE SCIENCES (EUCASS) Numerical Investigation of a 7-Element GOX/GCH4 Subscale Combustion Chamber ? ? ? Daniel Eiringhaus †, Daniel Rahn‡, Hendrik Riedmann , Oliver Knab and Oskar Haidn‡ ?ArianeGroup Robert-Koch-Straße 1, 82024 Taufkirchen, Germany ‡Institute of Turbomachinery and Flight Propulsion (LTF), Technische Universität München (TUM) Boltzmannstr. 15, 85748 Garching, Germany [email protected] †Corresponding author Abstract For future liquid rocket engines methane has become the focus of several studies on alternative fuels in the western hemisphere. At ArianeGroup numerical simulation tools have been established as a powerful instrument in the design process. In order to achieve the same confidence level for CH4/O2 as for H2/O2 combustion, the applied numerical models have to be adapted and validated against sufficient test data. At the Chair of Space Propulsion at the Technical University of Munich (TUM) several combustion cham- bers have been designed and tests at different operating points have been conducted. In this paper one of these subscale combustion chambers with calorimetric cooling and seven shear coaxial injection elements running on gaseous methane and oxygen is used to examine ArianeGroup’s in-house tools for combustion chamber performance analysis. 1. Introduction Current development programs in many space-faring nations focus on launchers utilizing a propellant combination of liquid oxygen (LOX) and liquid methane (CH4). In Europe, hydrocarbons have been identified as an alternative fuel in the frame of the Future Launcher Preparatory Programme (FLPP).14, 23 Major industrial development of methane / oxy- gen rocket engines is ongoing in the United States at SpaceX with the Raptor engine (staged combustion), at Blue Origin with the BE-4 engine (staged combustion) and in Europe at ArianeGroup with the Prometheus engine (gas gen- erator).
  • Qualification Over Ariane's Lifetime

    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.
  • Rocket Propulsion Fundamentals 2

    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.
  • Materials for Liquid Propulsion Systems

    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.
  • Variations of Solid Rocket Motor Preliminary Design for Small TSTO Launcher

    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.
  • Cryogenic Technology & Rocket Engines

    Cryogenic Technology & Rocket Engines

    ISSN (O): 2393-8609 International Journal of Aerospace and Mechanical Engineering Volume 2 – No.5, August 2015 Cryogenic Technology & Rocket Engines AKHIL GARG KARTIK JAKHU KISHAN SINGH ABHINAV B.Tech – Aerospace B.Tech – Aerospace B.Tech – Aerospace MAURYA Engg. Engg. Engg. B.Tech – Aerospace PUNJAB PUNJAB PUNJAB Engg. TECHNICAL TECHNICAL TECHNICAL PUNJAB UNIVERSITY, UNIVERSITY, UNIVERSITY, TECHNICAL JALANDHAR JALANDHAR JALANDHAR UNIVERSITY, akhilgarg.313@g kartik.lphawk@g kishansngh1996 JALANDHAR mail.com mail.com @gmail.com abhinavguru123 @gmail.com ABSTRACT 3.2 What is Cryogenic Rocket Engine? This paper is all about the rocket engine involving the use of A cryogenic rocket engine is a rocket engine that cryogenic technology at a cryogenic temperature (123K). This uses a cryogenic fuel or oxidizer, that is, its fuel or basically uses the liquid oxygen and liquid hydrogen as an oxidizer (or both) is gases liquefied and stored at oxidizer and fuel, which are very clean and non-pollutant very low temperatures. Notably, these engines were fuels compared to other hydrocarbon fuels like petrol, diesel, one of the main factors of the ultimate success in gasoline, LPG, CNG, etc., sometimes, liquid nitrogen is also reaching the Moon by the Saturn V rocket. used as an fuel. During World War II, when powerful rocket engines were first considered by the German, American and Keywords Soviet engineers independently, all discovered that Rocket engine, Cryogenic technology, Cryogenic temperature, rocket engines need high mass flow rate of both Liquid hydrogen and Oxygen. oxidizer and fuel to generate a sufficient thrust. At that time oxygen and low molecular weight 1.
  • Los Motores Aeroespaciales, A-Z

    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.
  • Aerospace Science and Technology 86 (2019) 444–454

    Aerospace Science and Technology 86 (2019) 444–454

    Aerospace Science and Technology 86 (2019) 444–454 Contents lists available at ScienceDirect Aerospace Science and Technology www.elsevier.com/locate/aescte Full-length visualisation of liquid oxygen disintegration in a single injector sub-scale rocket combustor ∗ Dmitry I. Suslov , Justin S. Hardi, Michael Oschwald Institute of Space Propulsion, German Aerospace Center (DLR), Langer Grund, 74239 Hardthausen, Germany a r t i c l e i n f o a b s t r a c t Article history: This work presents results of an effort to create an extended experimental database for the validation of Received 27 July 2018 numerical tools for high pressure oxygen-hydrogen rocket combustion. A sub-scale thrust chamber has Received in revised form 22 November 2018 been operated at nine load points covering both sub- and supercritical chamber pressures with respect Accepted 23 December 2018 to the thermodynamic critical pressure of oxygen. Liquid oxygen and gaseous hydrogen were injected Available online 11 January 2019 through a single, shear coaxial injector element at temperatures of around 120 K and 130 K, respectively. Keywords: High-speed optical diagnostics were implemented to visualise the flow field along the full length of the Co-axial injector combustion chamber. This work presents the analysis of shadowgraph imaging for characterising the Rocket combustion chamber disintegration of the liquid oxygen jet. The large imaging data sets are reduced to polynomial profiles of Combustion visualisation shadowgraph intensity which are intended to provide a more direct means of comparison with similarly reduced numerical results. Comparing half-lengths of these profiles across operating conditions show clear groupings of load points by combustion chamber pressure and mixture ratio.
  • Basic Analysis of a LOX/Methane Expander Bleed Engine

    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.
  • DEVELOPMENT STATUS of the VULCAIN THRUST Chambert

    DEVELOPMENT STATUS of the VULCAIN THRUST Chambert

    Acta Astronautica Vol. 29, No. 4, pp. 271-282, 1993 0094-5765/93 $6.00 + 0.00 Printed in Great Britain. All rights reserved Copyright ~) 1993 Pergamon Press Ltd DEVELOPMENT STATUS OF THE VULCAIN THRUST CHAMBERt E. IOgNFat, D. THELEMANN and D. WOLF Deutsche Aerospace, MBB GmbH, Space Communications and Propulsion Systems Division, Postfach 801169, D8000 Munich 80, Germany (Received 6 March 1991; receivedfor publication 14 October 1992) Abstract--The Vulcain engine planned to power the cryogenic main stage of the future Ariane 5 launcher is presently under development. MBB is responsible for the thrust chamber of this engine. After 4 years predevelopment and 2 years development, numerous successful tests have been performed on thrust chamber level and the engine development tests have just started with the first ignition tests. The thrust chamber is scheduled to be qualification tested in 1993 and the first technological flight is planned for 1995. The main development results for the thrust chamber are given in this paper as well as an outlook of the further development activities. 1. INTRODUCTION 2. THRUST CHAMBER MAIN CHARACTERISTICS The Ariane 5 represents the next member of the The main data of the Vulcain TC are summarized successful Ariane launch vehicles family. The flight below: performances will be Total TC thrust (vacuum) 1007.7 (kN) * with an upper stage (Fig. 1) Chamber pressure 100 (bar) in GTO max 6800 kg payload Mixture ratio 5.6 (--) in LEO 18,000 kg payload Specific impulse (vacuum) >439 (s) • with the Hermes space plane 23,000 kg. Mass < 620 (kg) Life 20 cycles and In both cases the lower stage is powered by two 6000 s cumulated solid boosters (P230) and one single cryogenic stage lifetime (H155).
  • 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

    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).
  • Extensions to the Time Lag Models for Practical Application to Rocket

    Extensions to the Time Lag Models for Practical Application to Rocket

    The Pennsylvania State University The Graduate School College of Engineering EXTENSIONS TO THE TIME LAG MODELS FOR PRACTICAL APPLICATION TO ROCKET ENGINE STABILITY DESIGN A Dissertation in Mechanical Engineering by Matthew J. Casiano © 2010 Matthew J. Casiano Submitted in Partial Fulfillment of the Requirements for the Degree of Doctor of Philosophy August 2010 The dissertation of Matthew J. Casiano was reviewed and approved* by the following: Domenic A. Santavicca Professor of Mechanical Engineering Co-chair of Committee Vigor Yang Adjunct Professor of Mechanical Engineering Dissertation Advisor Co-chair of Committee Richard A. Yetter Professor of Mechanical Engineering André L. Boehman Professor of Fuel Science and Materials Science and Engineering Tomas E. Nesman Aerospace Engineer at NASA Marshall Space Flight Center Special Member Karen A. Thole Professor of Aerospace Engineering Head of the Department of Mechanical and Nuclear Engineering *Signatures are on file in the Graduate School iii ABSTRACT The combustion instability problem in liquid-propellant rocket engines (LREs) has remained a tremendous challenge since their discovery in the 1930s. Improvements are usually made in solving the combustion instability problem primarily using computational fluid dynamics (CFD) and also by testing demonstrator engines. Another approach is to use analytical models. Analytical models can be used such that design, redesign, or improvement of an engine system is feasible in a relatively short period of time. Improvements to the analytical models can greatly aid in design efforts. A thorough literature review is first conducted on liquid-propellant rocket engine (LRE) throttling. Throttling is usually studied in terms of vehicle descent or ballistic missile control however there are many other cases where throttling is important.