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Deutsche Forschungsanstalt Fiir Luft- Und Raumfahrt DLR Lampoldshausen Research Centre Space Propulsion Division D-74239 Hardthausen A.K., Germany 1

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Deutsche Forschungsanstalt Fiir Luft- Und Raumfahrt DLR Lampoldshausen Research Centre Space Propulsion Division D-74239 Hardthausen A.K., Germany 1 Acta Astronautica Vol. 40, No. 2-8, PP. 151-163, 1997 Pergamon Published by Elscvier Science Ltd Printed in Grcst Britain 0094-5765197 517.00 + 0.00 PII: soo94-5765(97)00119-7 OPTIMIZATION OF DUAL-EXPANDER ROCKET ENGINES IN SINGLESTAGETO-ORBIT VEHICLES Detlef Manski, Gerald Hagemann and Hagen D. Sahick Deutsche Forschungsanstalt fiir Luft- und Raumfahrt DLR Lampoldshausen Research Centre Space Propulsion Division D-74239 Hardthausen a.K., Germany 1. Abstract Pn PI Dual-expander rocket engines offer a trajectory f adapted dual-mode operation during the ascent of a launcher, which may be of significant advantage for single-stage earth-to-orbit vehicles, when compared to conventional rocket engines with bell-type nozzles. Thii paper investigates a reusable single-stage earth- to-orbit vehicle with a constant payload capability of 16.5 Mg into low earth orbit, for the comparison of the dual-expander rocket engines with conventional rocket engines, using only hydrogen and oxygen as propellant combination in all engines. Published by Elsevier Science Ltd. 2. Nomenclature 2.1 Symbols A area Figure 1: Full-flow dual-expander cycle with F thrust oxidizer- aud fuel-rich preburuers 90 uormal earth gravitation I impulse 7 ratio mass flow s specific z Mach uumber tot total P pressure I’rK vacuu111 T radius mp vaporizatiou TOIF mass flow ratio ox/fu pra priinaq T temperature rel relative t time set secoudar> V velocit! 8 i~oiiiiiial propellaut Q angle of at tack 11 difference 2.3 Abbreviations c area ratio LEO low earth orbit ’ 2.2 Subscripts ODE one-dinlensional equilibrium C combustiou rhaiiiber SST0 Single-Stage-To-Orbit comb combustion ST System _Iualysis Rocket Launcher div divergence. untltidimeusional effect,s TSTO Two-Stage-To-Orbit e esit SSME Space Shuttle Main Engine eff effective zne iiiertial kine kinetic ft1 fuel 3. Irltroductiorr aud Lit.erature Review frzc frict iou h single-stage-to-orbit vertical-takeoff and landing grav gravitation mission was applied to find out the potential ad- hl heat loss vautage of full-flow dual-expander cycle eugiues ver- ma2 misiiig. misiiig point. sus conveutional staged combustion engines. For the oz oxidizer latter, t,wo differcut staged combustion engines cy- c:opyWlt @199G b? II,C RU~IKW~ i’~~1~1~4~~I I,? III,. .\E,o,.u tcstfil 1~ I- cles were esatniucd. .Iuy ad\;\utage of mixed-mode lUi@ or Aclolln”lKs a,,d :\\s,lo,l;mtIrr. 1111 , \I,,,, ,I( ,,,,, a.,,,,, II<.. lcwcrl lo IAF / Al:\:\ to ,vrld,.h >,I .,II r<a~,,u. 151 152 47th IAF Congress Mode 1 operation Mode 2 operation Figure 2: Dual-expander cycle mode 1 and 2 operations, following Berchel 12) bustion engine have no positive or, in some cases, even negative effects on payload delivery or dry mass re- duction on advanced shuttles. The main reason for this result is that the mass of a dual-expander engine is larger than that of a conventional engine under the constraint of using the same number of engines for the launchers. In order to get a more reliable comparison, a further analysis was initiated taking into account all the additional advantages of the dual-expander en- gine. In this analysis the optimization of the dual- expander engine for single-stage-to-orbit vehicles and the comparison with different staged combustion cycle I engines comprises the following items: common divergent nozzle extension 1 using CFD-calculations to estimate the higher Figure 3: Sketch of dual-expander thrust chamber, specific impulse losses of dual-expander nozzles, cut-away view of combustion chambers and nozzles for the comparison with conventional engine noz- zles [lo], [ll], propulsion could best be shown with single-stage-to- varying additional parameters in contrast to ear- orbit vehicles, because these vehicles are the most sen- lier investigations [G], [lo] for a fair comparison. sitive launchers with regard to performance and mass. This comprises also the number of engines. Dual-expander engines for future launcher applica- tions using one or two fuels and oxygen as oxidizer Taking the benefit of the cycle immanent thrust were suggested by Beichel [l]. The concept of this en- reduction capability of dual-expander engine for gine involves the use of a dense propellant combination the comparative analysis, and a with moderate performance during lift-off to provide complete redesign of the dual-expander engine cy- high thrust for the initial flight phase, and a lower den- cle shown in Fig. 1. sity, but better performing propellant combination in vacuum with a high specific impulse to reach t,he cle- sired orbit velocity. Figure 1 explains the principle Due to time restrictions, the last two items are not of the dual-expander cycle [3], and Fig. 2 shows the treated in full detail in this paper. two operation modes, following Bezchel [2]. In addi- tion, Fig. 3 gives a cut-away view of both combustion chambers and the nozzle extensions. Seyeral analyt- 4. System Allalysis Programme ST ical works on SSTO- and TSTO vehicles using hy- To analyse various kinds of rocket engine cycles for drogen/propane [2], [4], [S], [G] or hydrogen/methaue future space transport,ation systems, a propulsion sys- [G], [7] as fuels revealed lowest vehicle dry ~nasses for tem analysis programme ST has been developed which dual-expander engines in comparison to other engines. contains both DLR and NASA developed met,hods, Thus, the better ranking of dual-expander engines us- see [12]. [13], [14]. [15], [lG] for further details. The ing different fuels is well known. programme consists of several routines carrying out Other dual-expander engines with hydrogen as a sin- engine performance calculations [17], [18], engine sys- gle fuel, but with dual mixture ratios [5]. [8]. [9) also rem power matching calculations, engine mass calcu- revealed some benefits over conventional engines for lations, vehicle mass calculations, vehicle performance SSTO- and TSTO applications. Despite of t,his, ear- calculations and trajectory calculations [21]. By us- lier investigations by the authors [4] have led to the re- iug ST, many vehicle paramet,er are to be determined sults that dual-expander engines with hydrogen as sin- to fulfill the given conditions, such as payload mass, gle fuel compared with the conventional staged com- Euginr type, propellant combination. 47th IAF Congress 153 5. Reference vehicle A single-stage-to-orbit vertical-takeoff and landing ve- hicle was chosen for this comparative cycle analysis. Fuel The vehicle is of BETA- or Delta-Clipper-type, follow- ing proposals by [19] and [20] and a re-examination by [S]. In contrast to the proposal [19], in this analy- sis the number of dual-expander engines integrated in the SST0 launcher is assumed to be half of the corre- sponding numbers of conventional engines for better comparison purpose, because each dual-expander en- gine has a duplication of components such as nozzles, chamber, turbopumps etc.. Thus, eight engines were used in case of the staged combustion cycles and 4 engines in case of the full-flow dual-expander cycle. Arrangement of the engines are shown in Figure 4, viewing on the scaled nozzle exit areas at the base. Figure 5: Staged combustion cycle with fuel-rich Thrust reduction by this arrangement for the staged preburner and split oxidizer pump combustion engines will be performed at a predefined value for the maximum allowable acceleration by shut- 5.1 Vehicle model ting down two engines located opposite to each other, . Using ST-vehicle mass model [15], which can occur up to three times. 6 landing gears, The dual-expander engine has a build-in acceleration . cylindrical tank, diameter 8 m, elliptical endcaps reduction capability, achieved by shutting down the with an ellipse ratio of 1.5, secondary inner flow. The total engine thrust will then be provided only by the outer or primary flow . main propellants and extra tanks for reentry, or- which uses the total nozzle exit area, leading to an bit control and reserves, increase in specific impulse. This shut-down of the . constant payload of 16.5 tons into a 200 km circu- secondary flow is determined by the stage parameter lar orbit from French Guyana, Kourou. following mass ratio m,, an ESA requirement, ms,,,, . constant payload fairing of 2.6 tons until injec- m, = (1) tion, aud an ma,.,, + ms.,. ’ . acceleration limit 3.g0. and the propulsion parameter thrust ratio F,, F P*l Additionally, the following assumptious for the F, = (2) Fpn + Fsec propulsion model were considered: 5.2 Propulsion model . Using ST-performance. cycle and eugine mass model [lG]. hydrogen-rich preburner mixture ratio 0.85:1. oxygen-rich prel,urner nlMure ratio 100.1, isentropic efficiencies for a11 turbines and pumps at 75 %) pressurants. helium for osygen, hydrogen for hy- staged combustion dual-expander drogen, Figure 4: Sketch of lauuch vehicle base showing constant 11~~efficiency of 99 Y0, the arrangement of engines nozzle efficiencies taken from CFD-calculations. In this paper. a coustaut 3.g0-limit. in the trajectorl calculations aas assumed for the vehicles with staged 6. Cycles for coillparative analysis combustion cycles. For vehicles powered by dual- For comparison wit 11the advanced dual-espander eu- expander cycles. an acceleration decrease occurs at gine cycle, the mono mode staged combustion cycle switch-over from mode 1 to mode 2 operation. will be used, which has the highest performance of The following assumptions for the vehicle model were all 1nono mode rocket engines The expansion of all taken into consideration: plopcllauts from a high (.lli\llll)cr plessule makes it 154 47th IAF Congress possible to attain high overall specific impulses. How- ever, the staged combustion cycle has a maximum at- tainable chamber pressure.
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