Propulsion for Space Transportation for the XXIst...

AAAF-02-100

Recent Developments for Future Launch VehicleLOX/HC Rocket Engines

D. Haeseler, C. Mäding, A. GötzAstrium GmbHSpace InfrastructureMunich, Germany

V. Roubinski, S. Khrissanfov, V. BerejnoyChemieautomatics Design BureauVoronezh, Russia

6th International SymposiumPropulsion for Space Transportation

of the XXIst Century13-17 May 2002, Versailles, France

For permission to copy or republish, contact Astrium GmbH, Space Infrastructure81663 Munich, Germany

AAAF-02-100

AAAF-02-100, Versailles, May 2002 - 1 -

Recent Developments For Future Launch Vehicle LOX/HC Rocket Engines

D. Haeseler1, C. Mäding2, A. Götz3, Astrium GmbH, Space Infrastructure Div., 81663 Munich, Germany

V. Roubinski, S. Khrissanfov, V. Berejnoy, Chemieautomatics Design Bureau, 394005 Voronezh, Russia

1 email: [email protected]; 2 email: [email protected] email: [email protected]

Copyright © 2002 by the authors. Published by the Association Aéronautique et Astronautique de France (AAAF) with permission.Released to AAAF to publish in all forms.

ABSTRACTRecent investigations into the use of non-toxicpropellants for propulsion of reusable first stages formedium launchers or booster stages for TSTO launchvehicles show the potential to satisfy the market’sperformance and cost requirements. The main expectedadvantages are high propellant density, reduced handlingeffort, and reduced safety precautions.

System studies to identify the best non-toxic hydrocarbonpropellants have been performed. The chamber coolingperformance was assessed regarding limitations bypropellant dissociation and coking. Advantageouspropulsion system configurations were investigated.

Engines from Astrium and from CADB have been testedalready with hydrocarbon fuels. Currently, severalinjection elements are tested in a subscale chamber withLOX-methane and LOX-kerosene in a cooperation ofAstrium, DLR, CADB, and Rosaviakosmos.

ENGINE AND THRUST CHAMBERCONCEPT STUDIES

This study builds on the discussion in a previous paper[1, 2, 3]. Non-toxic propellants are of interest assubstitutes for current storables like NTO, MMH orUDMH which are highly toxic and chemicallyaggressive. Another application would be to replace solidpropellants for boosters because of their low specificimpulse and pollutive combustion products.

Requirements for non-toxic propellants were formulated,among these: Performance characteristics and densityequal to or higher than storable propellants, storability atambient conditions with at most moderate cooling effort,easy and cheap handling, and material compatibility.Screening a variety of candidates identified hydrocarbon(HC) fuels as suitable non-toxic propellants.

Engine system studies were carried out in order togenerate reference engine parameters for the threepropellant combinations chosen, LOX-methane, LOX-propane and LOX-kerosene. Thrust chamber design andperformance evaluations, in particular with respect to thecooling design, were performed.

Performance of Non-Toxic Propellants

An ODE-analysis of the theoretical specific impulse ofseveral LOX-HC and H2O2-HC combinations revealedmethane (CH4), propane (C3H8) and kerosene as bestperforming. The results are plotted in Figure 1.

Methane and propane require a moderate cryo-coolingeffort for liquid ground storage. Kerosene can be storedand handled at ambient conditions. It is the densest of allthree fuels and can enable compact launcher design. HCfuels will potentially reduce the propellant costs [4].

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N2O4-MMH

Figure 1: Theoretical mass-specific vacuum impulse ofvarious non-toxic propellants

Main Chamber Size Consideration

Possible future applications like liquid boosters require atotal thrust of 4000 kN and more [9]. The chosen enginedesign parameters of 2000 kN thrust at a chamberpressure of 150 bar reflect a typical future booster stageengine. Figure 2 shows the thrust chamber throatdiameter for a LOX-HC liquid booster. The thrustchamber size for the chosen design parameters seems tobe similar to Vulcain 2. Thus, the thrust chambermanufacturing facilities available in Europe pose noconstraint on the envisaged size of the chamber while a4000 kN-chamber would result in a 40% larger chamberdimensions.


Manufacturing installations as well as test benches inEurope are designed for today’s thrust levels of up toapproximately 2500 kN. A larger chamber and enginesize would require costly modifications.

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ηc* = 0.98, ηcf = 0.975

Fvac = 2000 kN

Fvac = 4000 kN

Vulcain-2: 274 mm

Figure 2: Main combustion chamber throat diameter asa function of chamber pressure and thrust

Main Chamber Cooling

The regenerative cooling capability of the main chamberis limited by the available fuel flow rate and by thecoking temperature limit. Above a fluid-specifictemperature threshold, which may occur inside thecooling channels of a regeneratively cooled combustionchamber or nozzle extension, hydrocarbons tend todecompose leading to the formation of a carbon layer atthe cooling channel wall (coking). This layer has aninsulating effect reducing the heat transfer from thecooled chamber wall into the coolant.

In existing LOX-kerosene engines soot deposition fromimperfect combustion is observed on the hot-gas side ofthe thrust chamber. The formation characteristics and theinsulating effect of a soot layer are important for thedesign of thrust chambers. Experience shows that LOX-ethanol does not produce significant soot, while LOX-propane and LOX-kerosene exhibited a heat fluxreduction due to soot on the chamber wall [5, 4].

Laboratory experiments performed with liquid methaneand propane flowing through copper tubes showed thateven trace amounts of sulfur-containing impurities resultin corrosion of copper-alloys used nowadays for chamberliners [7]. Electro-deposited gold and platinum coatingsgreatly reduced the corrosion. However, reliableapplication of such coatings will increase costs formanufacturing and quality assurance. Other authorsreported that no coking was detected in methane-cooledstainless-steel and copper tubes [8].

The extent of coking and sooting depends on theoperational conditions like chamber pressure, mixtureratio, coolant pressure and temperature. The upper walltemperature limit is given by the material properties andthe required lifetime. For the study the hot-gas walltemperature was limited to 800 K. The possible effect of

soot and coking could not be considered for lack ofreliable data.

The coolant-side wall temperature is limited to thecoking limit of the coolant. Cooling analyses for largethrust chambers were performed for methane, propaneand kerosene. Figure 3 shows the coolant and the walltemperature for a predicted chamber cooling withkerosene. A ZrO2 thermal barrier coating has beenassumed for this prediction to achieve coolant-side walltemperatures below the coking limit for kerosene.

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Figure 3: Coolant analysis for LOX-kerosene

The predicted coolant bulk temperature and the marginto the coking temperature limit are compared in Figure 4for the three propellants pairs with the usual chamberliner material temperature limit.

For the Astrium thrust chamber design with a copperliner, methane as coolant creates the lowest pressure lossand the largest margin towards the coking temperaturelimit. Propane can be employed, too, without a basicdesign concept change. The maximum occurring coolant-side wall temperature of around 700–730 K for copperliners is below the coking limit of methane and very nearthe coking limit of propane. Thus, the cokingtemperature is no cooling limitation for methane andalmost none for propane.

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Figure 4: Coolant heating and margin to coking


In contrast, kerosene cooling may require a thermalbarrier coating to keep the coolant-side wall temperaturebelow the coking limit of around 560 K. Zirconium oxideZrO2 was assumed for the cooling assessment. Filmcooling or mixture ratio bias near the wall (trimming)may be also necessary. Both film cooling and trimmingcan cause additional thrust chamber performance losses.

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Figure 5: Main chamber coolant pressure loss

Engine Cycle Selection

For all three propellants, the main properties of gasgenerator (GG) cycles and staged combustion (SC)cycles were compared, each cycle in turn with fuel-rich,and oxidizer-rich preburners. Examples of these cyclesare depicted in Figure 6.

LOXFuel LOXFuel LOXFuel

Figure 6: Studied engine cycles: Gas generator (left),fuel-rich staged combustion (center),oxidizer-rich staged combustion (right)

The gas generator cycles were rated by the engineperformance they produced. A fuel-rich gas generatormixture ratio produces better engine specific impulsethan an oxidizer-rich mixture ratio due to higher specificheat capacity of the turbine gas, see Figure 7. Thedifference in specific vacuum impulse is about 14 s forall three fuels studied.

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Figure 7: Performance for gas generator engines

For staged combustion cycles, the pressure cascadesresulting from thrust chamber cooling and turbopumppower requirements are compared in Figure 8 takingLOX-methane as example. The pumping requirementsare lower for the cycle using a oxidizer-rich preburner,because no fuel is rerouted to the preburner after passingthrough the thrust chamber cooling channels. The smallamount of fuel required for the preburner is delivered bya low-powered kick stage.

The fuel-rich cycle results in higher fuel-pumprequirements and requires an additional LOX-kick-stage,while the ox.-rich cycle results in similar lower pumprequirements without a kick-stage. The preburnerpressure is also lower in the ox.-rich cycle. However, theoxygen-rich environment of the preburner gas may causeadditional complexity for the turbines as well as for hot-gas lines and valves.

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Figure 8: Pressures in LOX-methane stagedcombustion engine

Non-toxic Propellant Comparison

Based on above results the fuel-rich gas generator cycleand the staged combustion cycle with oxidizer-richpreburner were chosen for comparing the three non-toxicpropellant pairs. Figure 9 gives the obtained engineperformance in terms of vacuum specific impulse.


Clearly, LOX-methane yields the best performance,LOX-propane is inferior by 5-9 s, while LOX-kerosene iseven more inferior by 10-17 s.

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Figure 9: Engine performance for hydrocarbon fuels

In the light of the above considerations, the threepropellant alternatives were rated as follows:

• Methane appears to be the first choice for an enginewith Astrium thrust chamber design operating onnon-toxic propellants due to its good performance,cooling capabilities and low soot production.

• Kerosene has the advantages of ambient storability,high density, and a vast amount of practicalexpertise available for rocket engine design.However, its cooling capabilities are inferior tomethane for conventional thrust chamber design. Acountermeasure is to reduce the heat flux byapplying a thermal barrier coating (TBC) to the linerhot-gas side. Coatings are currently underdevelopment at Astrium.

• Propane has less performance and cooling capabilitythan methane. Unlike kerosene it has to be cooledfor liquid storage and has lower density. Propanedoes not seem to yield a significant advantage andwas therefore eliminated as an alternative for thepurpose of further studies.

Propellant Cost Considerations

For today's launch vehicles the costs for the tankedpropellants are of secondary importance compared to thecosts of preparing the launch vehicle structure andsystems for launch and launch operations. In case drasticreduction of the latter costs become available either forlow-cost expendable vehicles or reusable vehicles, theinfluence of the propellant costs will increase. Somepropellant costs are given in [4]. Hydrogen isapproximately 15 times more expensive than 98% H2O2,while aviation kerosene costs only one quarter of H2O2.Liquid oxygen is even cheaper.

Kerosene will get expensive when the world-wide oilresources will cease, current estimations give a timeperiod of approx. 30-40 years. However, methane asliquefied natural gas is said to be available for another100-120 years. The costs for methane are said to be threetimes less than kerosene [10].

LOX-HC ENGINE CONCEPTSSeveral engine concepts were defined based on theresults of engine system trades as given above for thepropellants LOX-methane and LOX-kerosene.

The gas generator cycle engine as illustrated in Figure 10uses fuel-rich gas for the turbine drive. Boost pumps areused to obtain better turbopump efficiencies, and hence abetter engine performance.

Overall Oxidizer Pump:Pin = 3.0 barTin = 91 KPout = 187.5 barTout = 97.1 KmP = 415.05 kg/sPP = 8352 KWηP = 0.798DImp = 201 mm

Gas Generator:PGG = 127.2 barO/FGG = 0.405TGG = 900 K

LOX CH4

Thrust Chamber:Pc = 150 bar(O/F)cc = 3.7Ae/At = 38.7pNE = 0.429 barmcc = 514.46 kg/sIsp,v,TC = 356.8 s∆pCR = 80.9 bar

Turbine:Pin = 114.5 barPout = 8.0 barTout = 771 KηT = 0.558mT = 34.87 kg/sN = 15500 rpmDtip = 276 mm

Overall Fuel Pump:Pin = 3.0 barTin = 111 KPout = 260.9 barTout = 123.6 KmP = 134.28 kg/sPP = 10966 KWηP = 0.734DImp = 276 mm

Engine Parameters:Ispvac,Eng = 345.2 secFvac,Eng = 1860 kN(O/F)Eng = 3.091

Turbine Exhaust:mTEG = 34.87 kg/sFv,TEG = 60 kN

Figure 10: LOX-methane fuel-rich gas generator cycleengine concept.

The staged combustion cycle as illustrated in Figure 11uses oxygen-rich turbine gas from the preburner in orderto obtain lower and more balanced pump and turbinepressures. It is assumed that the hot oxygen-richenvironment can be handled without large risk. Boostpumps are used to allow for higher main pump speeds,which is important for the chosen single-shaftconfiguration.

Overall Oxidizer Pump:Pin = 3.0 barTin = 91 KPout = 358.21 barTout = 103 KmP = 405 kg/sPP = 15896 KWηP = 0.782UImp =199 mm

Preburner:PGG = 311.5 barO/FGG = 58.3TGG = 702 KmGG = 411.95 kg/s

LO2 CH4

Thrust Chamber:Pc = 150 barAe/At = 38.7pNE = 0.429 barmcc = 514.46 kg/sIsp,v,TC = 356.78 s∆pCR = 80.85 bar

Turbine:Pin = 311 barPout = 187.5 barTout = 642 KηT = 0.724N = 21800 rpmDtip = 239 mm

Overall Fuel Pump:Pin = 3.0 barTin = 111 KPout = 260.9 barTout = 126 KmP = 109.46 kg/sPP = 9843 KWηP = 0.667UImp = 278 mm

Engine Parameters:Ispvac,Eng = 356.8 secFvac,Eng = 1800 kN(O/F)Eng = 3.7

Fuel Kick Stage:Pout = 358.21 barTout = 150.16 KmP = 6.95 kg/sPP = 651 KWηP = 0.243

Figure 11: LOX-methane oxygen-rich staged combustioncycle engine concept.


MethaneLOXOxidizer Pump:Pin = 3.0 barTin = 91 KPout = 187.5 barTout = 97 KmP = 405.0 kg/sPP = 8062 KWηP = 0.807NPSP = 1.9 bar

Preburner:PGG = 323.5 barO/FGG = 0.36TGG = 904 KmGG = 148.87 kg/s

Thrust Chamber:Pc = 150.0 barAe/At = 38.7pNE = 0.429 barmcc = 514.46 kg/s∆pcr = 80.9 bar

Turbine:Pin = 323.5 barPout = 180.0 barTout = 863 KηT = 0.71mT = 148.9 kg/sn = 21200 rpm

Fuel Pump:Pin = 3.0 barTin = 111 KPout = 452.9 barTout = 136.2 KmP = 109.46 kg/sPP = 16713 KWηP = 0.678NPSP = 2.04 bar

Engine Parameters:Ispvac,Eng = 356.8 secFvac,Eng = 1800 kNIspSL,Eng = 308.8 secFSL, Eng = 1558 kN(O/F)Eng = 3.7

Oxidizer Kick Stage:Pout = 372.0 barTout = 108.2 KmP = 39.41 kg/sPP = 1113 KWηP = 0.564

Figure 12: LOX-methane fuel-rich staged combustioncycle engine concept.

The staged combustion cycle with fuel-rich preburner isalso feasible for LOX-methane as shown in Figure 12.The decision for the type of preburner depends on thecriticality of the hot ox.-rich environment for feedlinesand turbines created by a oxidizer-rich preburner.

Such and other engine concepts have been proposed alsoby other authors [9, 10, 11].

TESTS OF LOX-HC THRUSTCHAMBERS AND ENGINES

Aestus Tests with LOX-Ethanol and LOX-Methanol

The Ariane 5 upper stage engine “Aestus” was developedby Astrium for the storable propellant pair NTO-MMH.Tests were carried out with LOX-Methanol and LOX-Ethanol, which showed good thrust chamber operationalbehavior and performance, see Figure 13. Only minorengine modifications were necessary like implementationof an igniter and adaptation of some seal materials. Thesuccessful ignition and stable operation wasdemonstrated in a cooperation of Astrium and BoeingPropulsion & Power towards a non-toxic orbitalmaneuvering engine OME for the Space Shuttle [1].

Figure 13: Aestus engine test with LOX-ethanol byAstrium and Boeing Propulsion & Power

Conversion LOX-Kerosene to LOX-Methane

In 1998 CADB carried out two hot tests of theexperimental demonstrator engine RD-0110MD. Thebasic engine RD-0110 is currently used in the third stageof the Russian Soyuz Launcher using LOX-kerosene[12]. The engine RD-0110MD run during thedemonstrator test with LOX-LNG (liquefied natural gas)at chamber pressure 54 bar and GG pressure of 48 bar,see Figure 14.

Figure 14: RD-0110MD demonstration test with LOX-LNG (liquefied natural gas) by CADB

Cooperation Astrium - CADB

In order to characterize different LOX-Hydrocarboninjection concepts, several types of injection elementsare studied theoretically and experimentally in the frameof the TEHORA-2 research cooperation in cooperationwith CADB in Russia.

The thrust chamber and injection technologies currentlyestablished in Europe stem from various enginedevelopments. Chambers with storable propellants(Aestus) as well as chambers for cryogenic propellants(HM-7B, Vulcain, Vinci) were developed. A futuredevelopment of a LOX-HC engine in Europe and theassociated thrust chamber technology will be basedprimarily on the application of existing technologies tothe new propellant combination with necessarymodifications.

First activities focused on the injection concepts andelectrical igniters. Later work needs to consider issueslike compatibility of current liner materials copper andcopper-alloys with the hydrocarbon fuel and itscombustion products, sooting of combustion products onthe hot gas side and the coking of the fuel in the coolingchannel.

Today, the prediction and layout of the chamber coolingis performed with analytical tools, which are based uponthe existing test data from cryogenic and storable

A

propellant engine chambers. The application of theseanalytical models to hydrocarbon propellants needs to beverified and checked with experimental data.

Injection

Table 1 compares the propellant injection temperaturesfor the two hydrocarbon fuels to the conditions withhydrogen. In the gas generator cycle LOX is injectedalmost at tank temperature from the pump discharge intothe main chamber (and into the gas generator), while thefuel is heated in the chamber cooling, lowering itsdensity to gaseous ranges at supercritical state.

In the staged combustion cycles the hot oxygen-richturbine exhaust is injected into the main chamber at highmixture ratio resembling a hot oxygen gas. The heatingof the fuel in the chamber cooling is comparable to thegas generator cycle.

Gas Generator Cycle

Hydrogen Methane Kerosene Hydrogen Methane Methane KeroseneLOX 95 K 95 K 95 K 95 K 95 K — —Fuel 100 K 250 K 390 K — — 250 K 390 KTEG — — —

Full Scale Main Combustion Chamber Injection ConditionsStaged Combustion Cycle

~600 K oxygen-rich

fuel-rich gas generator fuel-rich preburner ox.-rich preburner

~650 K fuel-rich

Table 1: Typical main chamber injection conditions

Coaxial injectors are used today in the engines HM-7B,Vulcain, Aestus. The liquid-liquid injection for LOX-kerosene gas generator engines is similar to the liquid-liquid injection of storable propellants, while theinjection of gaseous methane from the chamber coolingis similar to gaseous hydrogen injection. In contrast, thegas-gas injection of oxygen-rich gas together withgaseous fuel represents new conditions, which mayrequire an extension of the current coaxial injectionelement technology.

Testing injection elements for hydrocarbon fuels atsubscale level requires a good simulation of the realinjection conditions. Subscale tests may be performed at

lower chamber pressures compared to the laterapplication.

The main chamber injection conditions of methane arerepresented well if the fuel is injected not liquid butgaseous at nearly ambient temperature. It is furthermorepossible to use natural gas in place of pure methane,which needs not to be liquefied by cryo-cooling.

For representing the gas generator cycle chamber, LOXis usually present at research test benches. The hotoxygen-rich turbine exhaust for the staged combustioncycle chamber can be replaced by oxygen gas at ambienttemperature, which also is easily available (e.g. inpressure bottles). The industrial availability of methane-rich natural gas makes subscale tests for LOX-methanemain chambers easy and attractive.

Ignition

The current European engines all use pyrotechnicignition, while an electric spark igniter is foreseen for theVinci engine. Test benches nowadays already use suchspark igniters. The main task for a development of areliable ignition system is the verification of the requiredpower level, which depends also on the propellantinjection.

Subscale Injector and Chamber Testing

The test specimen consists of a subscale chamber withseveral interchangeable injection heads. The chamber,which had already been used during a previouscooperation, consists of several water-cooled sections toallow the evaluation of the heat fluxes, see Figure 15.The chamber pressure is in the range of 35-70 bar. At 70bar chamber pressure, the propellant mass flow rate isabout 7 kg/s. Ignition is performed by an electric igniterusing the same propellants.

The two propellant combinations LOX-methane andLOX-kerosene were chosen for experimentalinvestigation from the system studies. Liquid oxygen isinjected around 105 K, while both methane-rich naturalgas with approximately 98% methane content as well as

F
igure 15: Modular experimental subscale combustion chamber
AAF-02-100, Versailles, May 2002 - 6 -


kerosene are injected at room temperature. The injectionheads consist of body, LOX-posts, and fuel section withthe faceplate, see Figure 16.

Figure 16: Subscale injection head parts

Tests with LOX-methane have been performed inAugust-September 2001 on CADB’s test bench inVoronezh. A total of about a dozen tests were performedin the chamber pressure range 35 – 70 bar at mixtureratio in the range 3.1 – 3.8. Three different basicinjection types designed by CADB and Astrium weretested. The tests demonstrated successful and reliableignition and operation of all three injector types. Figure17 shows a typical test with its characteristic blue flamecolour, which is a pronounced contrast to usual tests withhydrogen.

Figure 17: Subscale chamber test with LOX-methane byAstrium and CADB

All three tested injection elements were of the coaxialtype building upon the similarity to the experience

acquired in past LOX-H2 development programs.Impinging elements were not considered. During the testprogram some geometrical parameters like injectionelement geometry and injection velocity could be variedthanks to the modular design of the test hardware. Figure18 shows the progress in optimization of the combustionefficiency achieved by this variation. It could be shownthat combustion efficiencies comparable to theexperience from LOX-H2 thrust chambers can beobtained with LOX-methane as well.

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Figure 18: Results of injection element testing with LOX-methane (preliminary evaluation)

The pressure drop oscillations measured in thecombustion chamber and upstream of the injectionelements did not show significant higher values as thoseexperienced in similar tests with LOX-H2. The measuredcoolant water heat up did coincide well with theprediction based on theoretical heat transfer relations.

The chamber wall was in good condition after 12 tests.Some discoloration could be observed from the injectorto downstream of the throat, see Figure 19. A slight sootlayer was observed, which did not grow with increasingtest duration and which could be cleaned easily.

Figure 19: Chamber wall section after 12 tests showingdiscoloration due to heat load from injectionelements

A torch igniter operating with GOX and methane wasused in the tests, see Figure 20. Functional tests wereperformed with variation of the igniter energy. Reliableignitions could be achieved in all tests in the subscalechamber.


Figure 20: GOX-methane torch igniter

Similar tests with LOX-kerosene are scheduled to beperformed until July 2002. Three basic injectionelements designed for the injection of liquid oxygen andliquid kerosene will be tested. A continuation of thissuccessful cooperation is in preparation with furtherexperimental and design work.

SUMMARY AND CONCLUSIONSNon-toxic hydrocarbon propellants have becomefashionable recently aiming at more friendlyenvironmental conditions for easier handling and thusreduced space transportation costs. Liquid methaneprovides the best performance and chamber coolingcapability, while kerosene requires thermal protectioncoating or film-cooling due to the low cokingtemperature limit but is dense and easily storable. Thedanger of corrosion due to small impurities in methanewas observed differently by various authors and it needsto be checked for the envisioned methane compositionand liner material. It should be recommended to keep thesize and thrust level of future engine developmentswithin the range experienced today for Europe.

The engine system trade-off showed that a fuel-richturbine gas results in higher engine performance of gasgenerator cycle engines. The sooting associated withfuel-rich hot gases needs to be studied in further detail.Oxidizer-rich turbine gas seems to be more attractive forstaged combustion cycle engines in view of pumppressures, thus turbopump complexity. Sooting isavoided in oxidizer-rich gases. However, the oxidizer-rich environment may require specific material selectionor coatings to reliably prevent hazards. A fuel-rich stagedcombustion cycle shows quasi the same performance andacceptable pump pressures compared to the ox.-richcycle in case of LOX-methane, thus a fuel-rich cyclewould avoid problems associated with the hot ox.-richenvironment.

Further analyses of the engine configuration showed thatLOX-methane is suited for single-shaft turbomachinery

configuration like LOX-kerosene. For high performanceand compact turbomachinery, boost pumps should beemployed.

To gain experimental experience for injection, ignition,and combustion of LOX-Hydrocarbon, subscale chambertests with different types of injection elements for LOX-methane were performed in cooperation with CADB.These tests with LOX-methane demonstrated successfulignition and operation of three injector types. Thecombustion efficiency could be increased by variation ofinjector parameters to values as high as the pastexperience with LOX-H2. A GOX-methane torch igniterwas used successfully. These injector trade-off testingwill be complemented in summer 2002 with subscaletests for the propellants LOX-kerosene using the samesubscale chamber with new injection elements designedfor liquid-liquid injection of LOX-kerosene.

A new phase of the cooperation is currently underdiscussion aiming at continuing the testing of injectionelements and also including some cooling testing withmethane and kerosene.

ACKNOWLEDGEMENTSPart of the system studies were performed in the frame ofa project partially funded by the German AerospaceCenter (DLR) under project number 50JR9911. Thissupport is gratefully acknowledged here. The LOX-HCinjector element tests were contracted by CADB withcoordination by Rosaviakosmos. The fruitful cooperationwith these two institutions is acknowledged here as well.

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Gorokhov, S. Khrissanfov: Combustion Chambersand Engine Concepts Using Hydrocarbon Fuels forFuture Launch Vehicles, IAF-01-S.3.06, 52ndInternational Astronautical Congress, October 2001,Toulouse, France.

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