[American Institute of Aeronautics and Astronautics 9th International Space Planes and Hypersonic Systems and Technologies Conference - Norfolk,VA,U.S.A. (01 November 1999 - 05 November

(c)l999 American Institute of Aeronautics & Astronautics or published with permission of author(s) and/or author(s)’ sponsoring organization.

AlAA-99-4813. Airbreathing Hypersonic Propulsion System Integration within FESTIP FSSC-12

Sven Kopp, Sebastian Hollmeier, Hans Rick Technische Universitgt Miinchen D-85748 Garch/ng,Germany

Otfrid Herrmann DaimlerChrysler Aerospace AG D-8 1663 Mijnchen,Germany

AIAA 9th International Space Planes and Hypersonics Systems and Technologies

Conference

1-5 November 1999 Norfolk, VA

For permission to copy or to republish, contact the American lnstitute of Aeronautics and Astronautics, 1801 Alexander Bell Drive, Suite 500, Reston, VA, 20191-4344.

(c)1999 American institute of Aeronautics & Astronautics or published with permission of author(s) ardor author(s)’ sponsoring or!4anization.

Airbreathing Hypersonic Propulsion System Integration within FESTIP FSSC-12

S. Kopp, S. Hollmeier, H. Rick Institute of Flight Propulsion, Technische Universit%t Miinchen

.

0. Herrmann Military Aircraft, DaimlerChrysler Aerospace AG

Abstract

This paper presents the first stage of a two-stage-to- orbit (TSTO) space transportation system, studied within the European FESTIP-programme, as well as its highly integrated air-breathing propulsion system. Compared to previous design studies (e.g. the former German hypersonic research reference concept “Sanger”) stage separation will be at a relatively low flight Mach number of 4.0. The lower stage is powered on its entire flight by turbo engines with reheat only. Therefore, the use of ramjets together with complex turbo engine closure mechanisms can be avoided. Performance characteristics of the lower stage considerably depend on engine integration and thus on the components air inlet and exhaust nozzle. Within this study a two- dimensional ramp inlet with mixed external and internal compression, a single expansion ramp nozzle (SERN) and the performance of the installed turbo engine have been examined in detail.

Nomenclature flow area force length Mach number mass-flow rotor speed pressure dynamic pressure

Copyright 6 1999 by DaimlerChrysler Aerospace AG. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission.

1

T Fl 6 1-l P I-l.

temperature pressure parameter, p/prs~ mass-flow ratio, tic/r& = Ao/Ac pressure ratio thrust vector angle temperature parameter, T/TISA

Indices / Abbreviations

0” C corr. FESTIP

HPC HPE HPT LPC LPT SERN SST0 TSTO t th

1 Introduction

ambient condition inlet entry condition capture corrected Future European Space Transpor- tation Investigations Programme high pressure compressor high pressure engine high pressure turbine low, pressure compressor low pressure turbine Single Expansion Ramp Nozzle Single Stage To Orbit Two Stage To Orbit total throat

Within FESTIP (the Future European Space Transportation Investigations Programme) various future space transportation systems have been examined. The target is to assess the manifold systems using the same criteria and hence compare them in an unbiased manner. At the end of this process, in the presence of increasingly severe com- petition on the launcher market, one single con-

(c)1999 American Institute of Aeronautics & Astronautics or published with permission of author(s) and/or author(s)’ sponsoring organization. ___-

FSSC-01 FSSG01 HPE FSSC-02 TRIPROP FSSCS

FSSG9 FSSGI 5SOH--IT FSSGQeO-TT FSSG12T FSSGI 5SOH-FT FSSC-Qel-FT FSSC-12D FSSGI 50AE

FSSG5 AEROSPIKE FSSC-5 HPE

FSSGI 6SR FSSC-16ASR FSSC-i 6FR

Figure I: The different FESTIP system study concepts

cept shall be identified, which promises drastic cost reduction as well as technical feasibility. Along with development and operation cost, reliability, reusability, operational restrictions, and contingent system failures have to be studied in detail [2][11]. In the following the different concepts discussed will be presented briefly. This paper is dedicated to concept (FSSC) 12, which is a horizontal take-off and landing TSTO with an air breathing lower stage (turbofan engines with afterburner) and a rocket powered upper stage. The technological challenge of then lower stage propulsion system is the high temperature level,of the working cycle at high flight Mach numbers as well as the interaction of airframe and engine. Particularly the design of the air intake has a considerable impact on the performance characteristics of the installed propulsion system [I].

2 FESTIP

FESTIP has been initiated in 1994 by the Euro- pean Space Agency (ESA) in order to assess future cost-efficient concepts of fully or partially reusable space transportation systems. Together with the

2

concepts the required technologies to be developed in near and mid term are to be identified. In figure 1 the most important concepts can be seen [7]:

FSSC-1 represents a single-stage vertical take- off and horizontal landing winged spacecraft (FSSC-2 with a hydrogen-oxygen-kerosene mix as propellant).

FSSC-3 is a vertical take off and landing concept without wings similar to the “Delta Clip- per” DC-X studied in the US.

FSSC-4 uses the initial acceleration of a sled - running on skids - to take off and lands hori- zontally.

FSSC-5, similar to the lifting body X-33/ Ven- ture Star, takes off vertically and lands hori- zontally. It has been studied with an serospike nozzle as well as a common bell nozzle.

FSSC-9 is a TSTO transport system, where the upper stage can either be a fully reusable new concept or a modified Ariane core stage.

(c)l999 American Institute of Aeronautics 2% Astronautics or published with permission of author(s) and/or author(s)’ sponsoring organization.

0 FSSC-12 is the only remaining concept which uses an air breathing engine and will be discussed in detail.

0 FSSC-15 is derived from concept 4, either or- biting only once without sustained circulation (OAE = once around earth) or even not reach- ing orbit at all and landing at a different place than launch site (SOH = sub-orbital hopper). In this concept payload has to receive an additional acceleration by an increased kickstage to be positioned in orbit.

Figure 2: Delta wing configuration FSSC-12D

. FSSC-16 is similar to concept 9, while in the fully reusable version both stages would be al- most identical (Siamese configuration).

All concepts except FSSC-12 are powered by rocket engines with hydrogen and oxygen as propellant (FSSC-2 with a hydrogen-oxygen-kerosene mix). Either newly developed high pressure engines (HPE; FT = future technology) or “Vulcain-11”-like rocket engines (TT = today’s technology) are used. Technology has to be guaranteed by the year 2005 for a following programme decision. Figure 3: Trapezoidal wing configuration FSSC-

12T.

3 Design process of an air breathing space transportation system

In the first step of FESTIP three single-stage and three two-stage-to-orbit spacecraft with airbreathing propulsion systems have been studied [S]. The engines considered’were:

l Turbojet, Turbofan, Variable Cycle Engine (VW

a Rarnjet

l Air-Turbo-Rocket

l Mach 15 Scramjet

o Ejector-Rocket + Scramjet

a Air-Ejector-Rocket + Scramjet.

l Liquid Air Cycle Engine (LACE)

l Synergetic Air Breathing Rocket Engine (SABRE)

Following a first assessment with regard to performance, operational aspects, feasibility and development cost, all single-stage concepts had to be dropped, mainly due to technology availability. For the two-stage concepts separation Mach numbers between 3.0 and 8.0 were examined. The pure turbojet engine was identified as the most adequate propulsion system with technology already avail- able. Maximum operational flight Mach number for turbojet engines has been assumed to be Mach 4.0 determining the final FSSC-12 stage separation Mach number. The differences from the original “S%nger” configuration are two new lo.wer stage de- signs which have been developed with respect to the FESTIP acceleration mission (no cruise or minimum range requirement) and the earlier separation point. It is a fairly conventional airframe-wing design with 8 engines below the wings [5]. ‘Two different “basic configurations” were studied:

1) High degree of integration of the upper stage, into the airframe and delta wing of the lower stage (figure 2). This design has favourable aerodynamic characteristics at the cost of a complex, heavy and


Engine section of “Shger” propulsion system

Engine section of FESTIP-FSSC-12

ramp nozzle

Figure 4: Differences between “SZnger” and FESTIP engine concept

expensive structure. 2) The wing of the upper stage is not integrated

into the trapezoidal wing of the lower stage, forming a “biplane” (cf. figure 3). Overall drag of this configuration is considerably higher than that of configuration 1 so that the upper stage rocket engines have to. be operated simultaneously from flight Mach number 1.3 on. During this phase the upper stage receives hydrogen and oxygen from the lower stage (“cross-feeding”).

As the turbo engines characteristics are more important for the delta wing configuration (they ex- clusively provide thrust from take-off to stage separation), this configuration has been chosen for the TU Miinchen as the baseline concept [lo].

4 Reference turbo engine

The reference lower stage propulsion system is a two-spool no-bypa& turbo engine followed by an afterburner. At design point (flight. Mach number 1.2) the overall pressure ratio (OPR) is 13.5 (&PC = 3; Ii&c = 4.5) requiring two low- pressure and four high pressure compressor stages (cf. figure 5). Possibly the fourth HPC-stage can be omitted. Compressor delivery temperature is liited to 12OOK. i&ximum turbine entry temperature is 2100K at burner exit. High and low pressure turbine both consist of a single cooled

stage. The afterburner is operated stoichiometri- tally. Only during the separation manoeuvre over- stoichiometrical reheat becomes necessary.

main burner afterburner

LPC HPC HbT LPT

Figure 5: Reference two-spool no-bypass-turbo engine with afterburner

5 Engine airframe integration

During the design process of engine integration various aspects have to be considered: Performance analysis, thermal analysis and mass / structure analysis. Precompression, the treatment of fore body boundary layer, air intake, thrust nozzle, as well as the fuel and cooling system have to be matched and optimized in an iterative process. Ad- ditionally, external performance, bleed air off-take and especially thrust vector control have to be


taken into account [l]. To create a compatible formulation of all forces

acting on airframe and propulsion system an overall bookkeeping system was defined. All effects caused by the engine are summarized in the propulsion data set (with a resulting net thrust vector and a correlative momentum), while, forces acting on the defined airframe are contained in the aerodynamic data set.

bleedair

d

Figure 6: Hypersonic inlet system - geometry defi- nition

5.1 Inlet

The air inlet has to deliver the required air flow for the turbo engine with minimum pressure loss, decelerating supersonic ambient flow to subsonic speeds before entering the compressor section. This deceleration is performed discontinuously through several external and internal oblique shock waves. A two-dimensional geometry with two movable external ramps (C&,&J) and a fixed inlet lip was chosen (cf. figure 6). At supersonic flight the ramp position and inlet throat area Ath define the captured air mass-flow (in the relation /-I = tie/tic = Ao/Ac). In order to minimize the risk of inlet choking and hence subcritical inlet operation the external inlet ramps have to be positioned to maintain throat Mach numbers of at least Math = 1.4.

The final deceleration of the inlet flow to subsonic speed is performed by a normal shock wave which has to be stabilized in the divergent section of the inlet just behind the throat. When the captured mass-flow is given by external geometry and the required corrected mass-flow tiCorr = rFz~&/6 by the turbo engine’s rotor speeds, air flow matching is obtained automatically, determining the position of

the final shock wave and thus the compressor entry pressure, pt2.

0 0.2 0.4 0.6 0.8 1.0 1.2

A,,, [m21

% %i z t=

(a) Influence on throat Mach number

-.- 1 1.4' 1.8 2.2 2.6

Ma, L-1

(b) Influence on pressure recovery

Figure 7: Influence of inlet ramp variation on throat ‘Mach number Math and pressure recovery l&let at a flight Mach number of Ma, = 3.6

In order to avoid an inlet unstart the Mach number in front of the final shock wave has to be main- tained at least AMa = 0.2 above throat Mach number. For a local inlet entry Mach number of Mao = 3.0 (flight Mach number of Ma, = 3.3), figure 7a shows the throat Mach number as a function of throat area for various external ramp po- sitions. When & becomes too small the inlet chokes. Throat Mach numbers between 1.0 and 1.4 are difficult to stabilize. However, a higher throat Mach number results in higher air flow speeds in

(c)l999 American Institute of Aeronautics & Astronautics or published with permiss& of author(s) and/or author(s)’ sponsoring organization.

front of the normal shock wave and hence in higher pressure losses. The resulting overall inlet pressure ratio can be seen in figure 7b.

% 1.0 5.

0.8

0.6

0.4

0.2

0.0 0.2 0.4 0.6 0.8

%let [-I

(a) Inlet entry Mach number Mao = 3.0

z 1.0 3

0.8

0.6

0.5 0.6 0.7 0.8 0.9 ITinlet [-I

(b) Inlet entry Mach number Mao = 2.25

Figure 8: Capture area ratio /I as function of inlet pressure recovery lIiIinlet for two different inlet entry Mach numbers

At constant turbo engine speed the corrected engine mass-flow riz,,,, is approximately constant. At supersonic ‘%tarted” inlet operation the captured mass-flow is generally given by the external inlet geometry. When the captured mass-flow exceeds the required mass-flow the internal shock system collapses and mass-flow adaption occurs through a subsonic flow field behind a detached shock wave in front of the inlet lip. fiepending on this shock wave position very high surface forces (lift and drag) can

be the result as the static pressure rises considerably. In figure 8 the ratio of captured mass-flow to design mass-flow p= rits/rizc = Ao/Ac is plotted as a function of the pressure recovery &let for two different inlet entry Mach numbers (Mao = 3.0 and Mao = 2.25). Two distinct areas exist: normal inlet operation (inlet start) with higher mass-flow and choked (sub-critical) operation.

0.3 0.4 0.5 0.6 0.7 0.8 0.9 4linlet [-I

(a) Inlet entry Mach number Mao = 3.0

100

0 0.6 0.7 0.8 0.9

%let [-I

(b) Inlet entry Mach numb& Mao & 2.25

Figure 9: Corrected mass-flow .ti,,,, = tis@/6 as function of inlet pressure recovery &let for two different inlet entry Mach numbers

The effects of geometry variation on the corrected engine mass-flow. are shown in figure 9. It can be seen that in started inlet condition the captured mass-flow cannot be reduced below a minimum vahmwithout forcing the inlet to unstart.


inlet ramp failure situations. The captured air flow is assumed to pass a normal shock wave, whereas the spillage flow describes a flow past a blunt body forming a detached shock wave which may be approximated by a hyperbola (figure 10) [S].

Figure 10: One-dimensional real time model of subcritical inlet

Accordingly turbo engine thrust may not be fully reduced, especially after the separation manoeuvre, as inlet unstarts have to be avoided at high flight Mach numbers.

0 0.2 0.4' 0.6 0.8 1.0 lilhil,,,

(a) Position of detached shock

2.3 2.0 1.0 0.0 Mach number

Figure 11: Comparison of Id-realtime model and CFD flow field (Mach number distribution) at M~A, = 2.25, ti/7jlmax = 0.629

However, unstart conditions will occur at transonic and low supersonic flight Mach numbers. In order to estimate the effects on aircraft performance at transonic speed and in the case of engine failures inlet choking were studied. A one- dimensional model was developed to determine the detached shock wave position and static pressure distribution on the inlet ramps and across the capture area to calculate inlet forces reliably even in

73’ 160

g 140

Q 120

100

80

60

40 20

- - - - - L - - - -L _ - - _ c - - _I---- 1 I I I I

( r--k- t / I ----

----

1.2 1.0 0.8 0.6 0.4 0.2 IPi

(b) Static pressure distribution on 2nd inlet ramp

Figure 12: Comparison of Id model with CFD Cal- culation for a choked inlet, riz/riz,,, = 0.629

Two-dimensional Euler calculations were carried out to analyse the flow field of an imstarted inlet. At ‘a constant Mach number of 2.25 the inlet throat area was reduced and the captured mass- flow decreased accordingly. The Mach number distribution of the flow field can be seen in figure 11. Compared to the CFD-calculations, the po-

(c)l999 American Institute of Aeronautics & Astronautics or published with permission of autho;(s) and/or author(s)’ sponsoring organization.

sition of the detached shock wave is represented by the one-dimensional model extremely well. Fig- ure 12a shows that the shock position can be pre- dicted reliably for various mass-flow rates, even for a completely blocked inlet. The pressure distribution along the second inlet ramp is plotted in figure 12b with good agreement of Id and 2d calculations.

5.2 Single Exparision Ramp Nozzle

The exhaust nozzle for hypersonic aircraft has to work over a wide range of different operating conditions. From take-off to separation at IL!fa, = 4.0 the nozzle pressure ratio varies from 3 to about 100. For a complete expansion to ambient pressure large nozzle area ratios have to be realized. These large area ratios can only be achieved by using parts of the airframe structure as part of the expansion nozzle. A two-dimensional single expansion ramp nozzle is considered most suitable for hypersonic applications. A shear layer divides the hot nozzle flow from the ambient flow (Cf. figure 13).

single expansion

Figure 13: Two-dimensional CFD flow field of the FESTIP single expansion ramp nozzle, AIuoo = 1.2

High requirements are set on nozzle performance at low and high Mach numbers. At high Bight Mach numbers even small changes in nozzle performance result in large deviations of net installed thrust. In the transonic flight range asymmetric thrust noz- zles designed for high tight Mach number operation are known to produce large changes in gross thrust vector angle u (of up to 20”) [9][12]. As !Js~m > 3 throughout the mission, only critical nozzle flow occurs. However, at subsonic and transonic flight subsonic flow velocities will be found

in the nozzle region in the external flow around the lower engine cowl. Due to the relatively low mass-flow and static pressure in the main nozzle flow (compared to the nozzle design point at separation Mach number) the afterbody volume cannot be filled sufficiently. This effect results in considerable negative hft and a nose-up pitching momentum. This behaviour notably influences the air- worthiness of the aircraft. However, the injection of the diverted forebody boundary. layer in the divergent section of the nozzle improves the nozzle ‘performance.

First of all a reference thrust is calculated that would occur with complete expansion of hot nozzle flow to the ambient pressure in the afterbody section. This will be transformed using pm-calculated thrust and momentum coefficients, which are read -from nozzle performance maps. NozzLpressure ratio, diverter flow and (if variable) nozzle throat area were used as input. These nozzle maps had been calculated with. the method of characteristics as- suming supersonic flow -throughout the afterbody region and had been verified with three-dimensional Euler- and Navier-Stokes calculations [3].

Initial geometry for

\:

Figure 14: Geometry optimization of the FESTIP single expansion ramp nozzle

To reduce system complexity, weight and cost a nozzle design with-fixed geometry (i.e. without variable nozzle throat area) is intended. Accordingly, mission, turbo engine and afterburner control have to be adapted. Especially during return-flight operational restrictions apply which will-be discussed in the following chapter.

The external nozzle geometry and the nozzle throat section were optimized with Navier-Stokes

(c)l999 American Institute of Aeronautics & Astronautics or published with permission of authoi(s) and/or author(s)’ sponsoring organization.

calculations for maximum gross thrust at flight Mach numbers of 1.2 and 4.0 [4]. Figure 14 shows the reference geometry as well as the two optimized nozzle geometries. Especially at Mach 4.0 the nozzle geometry has little influence on the gross thrust. For further performance calculations the optimum geometry for Ma=1.2 was chosen.

6 Engine Performance The reference acceleration mission of FSSC-12 does not contain cruise sections so that the lower stage will constantly be operated at full thrust along the ascent trajectory. Various limiters apply: HPC delivery temperature and HPT entry temperature may not exceed Tt3 = 1200K and Tt4 = 2100K re- spectively, mechanical rotor speeds are limited to 105% of the reference values at iUa, = 1.2, and aerodynamic rotor speeds n/a to 100%. The afterburner is operated stoichiometrically along the ascent trajectory.

200 - ~"-~'-T~---:--~---~--- t t,

01 1 I I I I I I 1 I I I I

g 3ml I I , I I I -- I--J---l---L-2-l-W-

engine station

Figure 15: Temperature and pressure distribution throughout the turbo engine, Mac0 = 4.0

The matching of inlet and turbo engine mass-flow has considerable influence on the installed performance. Within a narrow margin an equilibrium is obtained automatically through the position of the final inlet shock wave as described in section 5.1.

However, the inlet ramps have to be positioned to adapt the inlet mass-flow. Concessions have to be made to reach an ‘optimum for low additive inlet drag (high ,u) and high inlet pressure recovery (adequate inlet throat area that results in low, but still supersonic M&h). As a consequence, in cer- tain flight sections. extremely low inlet capture area ratios occur (e.g. ,u = 0.63 at Ma, = 2.4) [S][lO].

In figure 15 total pressure and temperature distribution throughout the turbo engine is plotted for flight Mach number Ma, = 4.0. It is remarkable that the pressure ratio between nozzle and compressor entry is below one (II = 0.97). At this point, the turbo engine is solely used for heat ad- dition with low pressure losses. LPT exit temperature is Tt5 = 1630K, while in the afterburner the exhaust gas temperature reaches Tt7 = 2500K.

Figure 16: Net installed thrust along ascent trajectory

Figure 16 shows the resulting installed net thrust along the ascent trajectory (5OWa). The relatively low thrust between Mach 2 and Mach 3.5 is a result of mechanical rotor speed limitation (HP rotor). Currently a modified configuration is being studied. After stage separation, the lower stage re- turns to the launch site cruising at subsonic speed. For deceleration the turbo engines must not be switched off as an inflight restart always represents an additional risk. To reduce structural loads and fuel consumption the return flight is performed at

= 10kPa. If the afterburner is turned off the ef- gtive nozzle throat area has to be reduced by 20% in order to maintain stable compressor operation. Since variable nozzle area is avoided for simplicity reasons this has to be obtained through spoilers or

9

(c)l999 American Institute of Aeronautics & Astronautics or published with permission of author(s) and/or author(s)’ sponsori?g organization.

lateral bleed air injection into the throat section. On the other hand, if the afterburner is not turned off at return flight, total fuel consumption may be significantly higher. Which alternative is chosen has to be identified by further investigations in con- junction with the flight mission.

7 Summary The objective of the presented study was to analyse the performance behaviour of an air breathing propulsion system for a TSTO space transportation system. Flight regimes up to Mach 4.0 can be covered by turbo engines, feasible within the next decade. This statement was confirmed by all the contacted engine specialists and is validated by this study. To predict the installed propulsion system performance engine airframe interaction has to be considered particularly. An integrated design process for the turbo engine, the air inlet and the complete mission is necessary. Subcritical inlet operation at high flight Mach numbers has to be avoided under all circumstances. A relatively low system complexity (e.g. fixed nozzle geometry, restriction to two movable inlet ramps) can be obtained with optimized flight mission and engine operation. Considering reliability, re-usability and operational flexibility, a horizonal take-off and landing air breathing TSTO is an attractive candidate for the future European space transportation system.

References

PI

PI

PI

Bauer A., Betriebsverhalten luftatmender Kombinationstriebwerke fiir den Hyper- &haL!fIug unter besonderer Beriicksichtigung der Triebwerksintegration, Dissertation, Lehrstuhl fiir Flugantriebe, TU Miinchen, 1994.

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Dujarric C., Caporicci M., Kuczera H., Sacher P.W., Conceptual Studies and Technology Re- [12] quirements for a New Generation of European Launchers, IAF-97-V.3.03,’ 48th International Astronautical Congress, Turino, Italy, 1997.

Esch Th., Zur .Integration des Antriebes in ein HyperschallAugzeug untex besonderer

Beriicksicl.@ung der Krtiebilarizierung, Dis- sertation, Lehrstuhl fir Flugantriebe, TU Miinchen, 1997.

Esch Th., Optimierung der FESTIF-Diise mit- tels Navier-Stokes CFD-Verfahren, int. report, Lehrstuhl fiir Flugantriebe, TU Miinchen, 1998.

Herrmann O., Airbreathing Propulsion Activ- ity Report, FESTIP B2 / SCT Discipline Re- port WP4010, 1996.

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Hirschel E-H., Kuczera H., The FESTIP Tech- nology Development and Verification Plan, 8th International Space and Hypersonic Systems and Technologies Conference, AIAA 98-1567, Norfolk, Virginia, April 27-30,1998.

Hollmeier S., Simulation des Betriebsverhal- tens von Antrieben fiir Ramntransporter/ Hy- perschallffugzeuge, Dissertation, Lehrstuhl filr Flugantriebe, TU Miinchen, 1997.

Hollmeier S., Kopp S., Herrmann O., Hick H., Propulsion -/ Aircraft Integration of Hyper- sonic Vehicles With Active Thrust-Vectoring and Real-Time Simulation of Dynamic E@ine Behaviour, Journal of Turbo & Jet Engines, accepted for publishing, 1998.

Hollmeier S., Kopp S:, Hick H., l%‘SC-12 In- stalled ‘Turbojet Engine Performance, WP.- No. 4010-C FESTIP Interim Report, 1998.

Kuczera H., FESTIP System Study - An Overview, sth International Space and Hyper- sonic Systems and- Technologies Conference, Norfolk, Virginia, April 27-30, 1998.

Rick H., Kopp S., Holhneier S., Esch Th., Sonderforschungsbereich 255, Teilprojekt Bl, Leistungsberechnung und Betriebsver- haltensanalyse luftatmender Antriebssysteme fiir Raumtransp&ter / Hyperschallflugzeuge, Technische Universitat Miinchen, Ergebnis: und, Arbeitsbericht, 1998.

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[American Institute of Aeronautics and Astronautics 9th International Space Planes and Hypersonic Systems and Technologies Conference - Norfolk,VA,U.S.A. (01 November 1999 - 05 November