(c)2001 American Institute of Aeronautics & Astronautics ... · ramjet rocket combined cycle...

(c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

AiAA A01-28119

AIAA 2001-1793Vehicle Concepts for an EjectorRamjet Combined Cycle EngineA. Lentsch , H. Taguchi , R. Shepperd and M. MaltaNational Aerospace Laboratory of Japan,7-44-1 Jindaiji-Higashi, Chofu, Tokyo 182-8522, JAPAN

10th AIAA International Space Planesand Hypersonic Systems and Technologies

ConferenceMarch 24-27, 2001 / Kyoto, Japan

For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics1801 Alexander Bell Drive, Suite 500, Reston, VA 20191-4344


Vehicle Concepts for an EjectorRamjet Combined Cycle Engine

A. Lentsch*, H. Taguchi1', R. Shepperd* and M. Malta5

National Aerospace Laboratory of Japan,7-44-1 Jindaiji-Higashi, Chofu, Tokyo 182-8522, JAPAN

The study looks at air-breathing vehicle concepts from the lower end of the (missionaveraged) specific impulse range. The selected concepts are ejector rocket and ejectorramjet rocket combined cycle propulsion systems, which are compared to rocket-onlyconcepts. The use of such engines leads to high propellant mass fractions but also offersthe potential of only moderate increase of propulsion system mass. Furthermore, thesimple and technologically proven designs of such engines require lower developmentcost and are likely to result in low production and operational cost compared to otherairbreathing vehicle concepts.

The study tries to identify if single stage vehicles can obtain the little extra bit ofspecific impulse required to make them technically feasible as compared to rocket onlyconcepts. Alternatively an identical (sometimes Siamese or biamese) two stage vehicleis assessed, because also for this concept only one vehicle needs to be developed, sodevelopment cost remain nearly the same, only operational cost roughly double.

A one-dimensional thermodynamic model has been developed for the combined cyclepropulsion system. The trajectories are obtained from a mass-point model, in a singleplane, with standard atmosphere, Earth rotation and potential gravity field. With agenetic algorithm the vehicle trajectory as well as wing size are optimized for a giventarget orbit to obtain the lowest possible propellant mass fraction.

Nomenclaturea ejector mass flow ratio m^/miAvapo apogee orbit circularisation

speed incrementAOA Angle Of AttackCC Combined CycleCL->CD lift and drag coefficientsEJ ejector rocketGTOM Gross Take-Off Mass [Mg]leff Effective Specific Impulse [sec]lsp Specific Impulse [sec]L/D lift over drag ratioMECO Main Engine Cut-OffMg Mega-gram = metric tonMPa Mega-PascalO/F Oxygen-fuel mass ratioPMF Propellant Mass Fractionq dynamic pressureRJ ramjetRBCC Rocket based Combined CycleSMF Structural Mass FractionSSTO Single Stage To OrbitTSTO Two Stage To Orbit

* Guest Researcher, Spaceplane Group^Research Scientist, Aeroengine Division* Guest Researcher, Spaceplane Group§ Former Head, Spaceplane GroupCopyright (£) 2001 by the authors. Published by the American

Institute of Aeronautics and Astronautics, Inc. with permission.

IntroductionIn the past years single stage to orbit vehicles have

been studied intensively as candidates for the nextgeneration of fully reusable launch vehicles. However,recently rocket only single stage vehicles are increas-ingly believed to be beyond today's technologicalcapabilities.

The main question of this paper is: Can the mostsimple rocket based combined cycle engines offer justenough increase in specific impulse to put single stagevehicles into a technologically realistic perspective?

The reason for looking at "most simple" air-breathing concepts are low development cost and apotentially small increase in system mass. Develop-ment cost will be low compared to other air-breathingengines, because the concepts are proven, thereforethe required technological development steps are mod-erate and so is the development risk. Also the systemmass can be low because the mechanical systems ofthe engines are simple (e.g. no large turbo-machinery,heat-exchangers, etc.), the heat-loads are low dueto comparably low flight speeds and highly efficientinlets are not a necessity.

The "most simple" rocket based combined cycleengines are seen to be the ejector rocket and ejectorramjet. These engine types provide only a moderateincrease in mission averaged specific impulse. Because

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AMERICAN INSTITUTE OF AERONAUTICS AND ASTRONAUTICS PAPER 2001-1793


this leads to high propellant mass fractions, thevehicle layout and shapes will not change significantlyfrom current expendable and semi-reusable designs.This means rather blunt "wing-tank" designs with lowaerodynamic performance. However, also this smallchange in vehicle design will keep development costrather low, and allow bulky and light-weight solutions.

As an attractive alternative to SSTO concepts, anidentical (sia- or bimese) two stage (twin) vehicleconcept is analyzed. The two vehicles are completelyidentical and can alternately be used as booster andorbiter. Development cost will only marginally in-crease, because compared to a pure SSTO vehicle therequired modifications are small (fuel-cross-feedingand stage separation). Due to initial ejector-rocketpropulsion, the stage separation can take place inthin atmosphere during a temporarily ballistic flight,so that the separation resembles the one of currentexpendable booster stages.

Engine ModelEngine Operation Modes

The basic concept of an ejector ramjet is the inte-gration of three thermodynamic propulsion cycles:

1. Ejector Rocket/Air Augmentation Mode

2. Ramjet Mode

3. Rocket only mode

The Ejector Rocket operation spans the speedrange from take-off to about Mach=2^4 (dependingon the engine layout, size/thrust). Atmospheric air isentrained into the rocket jet only through jet-pumpingcaused by the rocket exhaust jet. Combustion usuallyoccurs during or after the mixing. The mixed flow isfinally expanded in a secondary nozzle.

The Ramjet takes over at about Mach Key parame-ters are engine mass-flow and flight dynamic pressure.A take-over speed of Mach ~2 is an optimistic assump-tion and requires a large mass-flow in order to gainenough accelerative thrust. The end of ramjet opera-tion is terminated, when the acceleration becomes toolow, and the use of the rocket engine is more efficient.(Mach 6) is set as upper speed limit for ramjet opera-tion. This is due to limitations from thermal loadingsof the engine as well as vehicle structure. Often how-ever the ramjet shutdown occurs earlier due to thelimited acceleration, which comes from the diminish-ing amount of heat, which can be released into theincreasingly hot subsonic flow.

The Rocket only operation is resumed as last op-eration mode. Due to the large speed range duringwhich the rocket engine has to accelerates the vehicle(roughly 3/4th of the total Av), its specific impulse

fuel

rI

fuel

fuel

simultaneousmixing & combustion

3 = 1QJO |:! gE I 1

LaS l55 .Ex

E i -

o

CO O

eji' GO

§ i «.-§.2 j C CO

% 11 •«"§CD -Q 11 ^P -^

2oO. ^ T3 >.

5* Sio> •==='xE "c o

•— CO

— Q.•o3~fuel

- I»^ ... mixing region-•-vSill ... combustion region- ... critical flow (Mach 1)

... shock

Fig. 1 Ejector Combined Cycle Concepts

has a very dominating impact on the mission averagespecific impulse. Therefore it is mandatory, that theintegrated rocket engine operates at highest efficiency,otherwise any gains from previous airbreathing enginecycles are easily offset.

Basic Ejector Concepts

Based on Foster's1 definition, a modified catego-rization of ejector concepts is suggested as shown infigure 1. The key distinguishing elements are:

• Supersonic vs. subsonic engine outflow

• Separated vs. combined mixing & combustion

• Fuel injection into the mixed vs .the unmixed airflow

Concept Selection

With the study target in mind (highest simplicity),concept II (in figure 1) was selected for further study.

This decision is based on the following considera-tions:

• A subsonic outflow (concept I) leads to loweroverall engine performance, therefore a supersonicoutflow is the preferred case.

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Diffusor CombinedMixer & Combustor

SecondaryNozzle

rthermal choking

Fig. 2 Ejector Ramjet

• A separation between mixing and combustion(concepts II & IV) has two key drawbacks:

- The engine length and hence mass increases- The rocket engine needs to be operated with

stoichiometric oxygen/fuel ratio, which usu-ally degrades the rocket performance.

Assumptions &; LimitationsA basic sketch of the selected engine concept is

shown in figure 2.The following assumptions have been made for the

modeling of the engine:

• Mass, momentum and energy equations are usedin explicit and integrated form, written in equa-tions for stagnation conditions and Mach numbersuch as commonly used in aero-engine thermody-namic cycle analysis,2.3

• Applying these three equations for themixer/ejector represent the ideal case. Dueto the high entropy (total pressure) losses, nofurther losses have been assumed such as forturbulence and friction.

• The specific heat capacity cp and specific heatratio 7 are computed with polynomial approxi-mations based on data from2 depending on gasmixture and temperature.

• Because of the wide temperature range of themixer and the large change of gas composition, cpand 7 are computed individually for each temper-ature according to the context, e.g. for computingthe speed of sound (^/jRT), 7 is computed withthe static temperature at this position, in case ofa change of state relation, cp and/or 7 are cal-culated between the specific temperature range,etc.

• Further more, the mixing between two inflows isperformed in 40 sequential steps. With a definedas the ratio of secondary to primary massflow, ineach step the amount of

step= (i)air is added to the primary flow. The controlvolumes for each step are shown schematically infigure 3.

Sequential Control Volumes____\_____

Fig. 3 Sequential Mixer Control Volume

• Each step in turn performs three sub-steps:

— Constant Area Mixing— Combustion— Expansion.

This approach was chosen to avoid time intensivenumerical solving of coupled differential equationsin order to gain a fast and acceptably accuratecomputer code to be integrated with the vehicletrajectory simulation, which is outlined furtherbelow.

• The nozzle flow is calculated as frozen flow.

• For the primary rocket nozzle a fixed area ratiois assumed, the secondary nozzle is adaptable inorder to ideally expand the flow to ambient pres-sure.

• In case the static pressure at the mixer outflow islower than the ambient pressure a normal shock isassumed at the end of the mixer in order to reachambient pressure.

• No dissociation effects are currently considered.

Engine Characteristics

Thermal ChokingThe choking limit is the key operation point for an

ejector rocket. The specific impulse of the ejectorrocket increases with increasing air massflow until themixer exit chokes. This is therefore the preferred oper-ation point. Choking, more precisely thermal chokingis caused by the decrease of the critical massflow den-sity. This behavior is shown in figure 4, where themassflow density, the critical massflow density as wellas the Mach number are plotted over a. a is the ratiobetween the secondary massflow (air) and the rocketmassflow.

The decrease of the critical massflow density (pu)*is caused by the total pressure loss of the flow due tomixing. The critical massflow density is proportionalto:

Pt(2)

Figure 5 shows the change of static and stagnationpressure over a. The stagnation or total pressure of the

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AMERICAN INSTITUTE OF AERONAUTICS AND ASTRONAUTICS PAPER. 2001-1793


600

Fig. 4 Massflow density over a

M (no sec.combustion)sec.combustion)

Fig. 6 Static temperature and Mach number overa with and without combustion

12

10

PPt

0.5 1.5alpha

2.5

120

100

80

60

40

20

0

Fig. 5 Static and total pressure over a

core flow drops over nearly 2 orders of magnitude, fromthe level of the rocket chamber pressure to the finalchoking condition. The massflow density also changeswith a. It is computed with the massflow conserva-tion rh = (pu) A due to the known massflow and areas(constant area mixing).

Simultaneous Mixing & Combustion

Figure 6 shows a comparison between mixing with-and without combustion. The static temperature andMach number are shown over a. This figure indicatesthe error, which occurs, when not considering the com-bustion during mixing. Although the specific impulseis quite similar at this operation point, the maximumstatic temperature as well as the maximum secondarymassflow are strongly deviating.

The reasons for the deviation are the different tem-perature due to heat release as well as the pressure lossassociated with heat addition into a high-speed flows,which is significant.3

Trajectory SimulationVehicle Definition

In order to assess the overall vehicle performance,the engine model has been integrated with a trajectorysimulation.

The trajectory simulation is a single-plane masspoint model, which considers Earth rotation (there-fore launch site latitude and orbital inclination),potential gravity field and employs the U.S. StandardAtmosphere.4 The vehicle ascent path is defined in 5sections, the first two sections are defined over time,the remaining one's as function of flight velocity.Input variables are either the pitch angle explicitly,or angle of attack, climb angle or dynamic pressure.In all cases except the direct definition of the pitchangle, proportional-differential control loops havebeen implemented to keep the vehicle on the desiredpath.

The trajectory simulation determines the requiredpropellant mass in order to reach a circular targetorbit of 250 km altitude and 29deg inclination. Part ofthe trajectory simulation is the coasting of the vehiclealong a transfer orbit and a final orbit circularisationapogee burn.

Vehicle mass modeling, hence determining themasses of the individual system components (struc-ture, engines, TPS, etc.), is not part of this paper -only the propellant mass fraction is determined. Thereference vehicle is based on the assumptions given intable 1.

Vertical/Horizontal Takeoff

Vertical as well as horizontal takeoff have been sim-ulated.

Vertical TakeoffIn case of vertical takeoff the initial thrust-to-weight

ratio was assumed to be 1.2. A higher value leads to

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VEHICLE DEFINITIONMass

GTOM *J 100 MgRocket

Chamber pressure 20.0 MPaCamber temperature 3600 KO/F ratio 6.0Area ratio, primary nozzle 30Primary & secondary nozzle 80Isp (sea level) *2 355 secIsp (vacuum) *2 448 sec

Mixer/Ejectorexit/inlet area 1.2secondary fuel injection no

Ramjetmin. operational flight speed Mach 2.0max. operational flight speed Mach 6.0Freestream capture ref. area 10-20 m2

Combustion efficiency (77) 0.8Combustor inflow Mach number 0.25Combustion target temperature 2500 K

Structural Limitsmin. wing reference area (for landing) 60 m2

max. dynamic pressure 60 kPamax. acceleration 30 m/sec2

max. total wing-load 2 x GTOM ]

max. negative wing-load -1 x GTOM 1

max. vertical (lift) accel. 30 m/sec2

_____max, negative (lift) accel.___________10 m/sec2

1 ...Gross Take-Off Mass2 ... calculated, no input data

Table 1 Vehicle definition

a lower (better) PMF in the simulation. However,the value of 1.2 was not varied because of in orderto optimize the thrust-to-weight ratio the effects ofchanging propulsions system mass as well as structuralmass need to be properly assessed with vehicle massmodels, which is not part of this paper.

Fig. 7 Vehicle concept outlines

h

CO

"cCDEO)CDO)

MFig. 8 Trajectory pitch control

requirements.

The thrust-to-weight ratio refers only to the rocket-engine thrust. For the combined cycle simulationsthe take-off thrust increases due to ejector effect to avalue of about 1.43.

The inlet free-stream reference area was varied be-tween 10-20 m2. For a vehicle with 100 Mg GTOM areference area of 20 m2 corresponds to about the fullcross-section area. Therefore a vehicle concept suchas (A) in figure 7 is seen to be most realistic for suchlarge inlet reference areas.

Horizontal TakeoffFor horizontal takeoff the initial thrust-to-weight

ratio was defined between 0.8 and 1.2. No launchassisting devices have been assumed. The vehicle firstaccelerates at zero angle-of-attack. After a dynamicpressure of 15 kPa is reached pitch-up is triggered.Lift-off occurs as soon as enough lift is generated,which is typically around a speed of Mach 0.5. Thishigh take-off speed is caused due to the fact, that thewing area is not sized to fulfil realistic take-off speed

At this point it has to be acknowledged that thetake-off speed is a critical design factor. This is inaddition to the problem of the required high mass forthe under-carriage, which has to be design to supportthe full GTOM.

A typical vehicle shape for horizontal take-off isshown in as vehicle (B) in figure 7.

Pitch Control

In order to control the vertical trajectory profilethe vehicle pitch is controlled. To allow different pitchcontrol strategies, the ascent can be split in numerouspitch control segments as indicated in figure 8. Withineach segment a profile is defined for a selected inputvariable. The possible input variables are dynamicpressure, an gle-of-attack, flight-path-angle and pitch.

Except for the direct pitch-profile input, aproportional-differential control law is used to followthe desired profile.

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AMERICAN INSTITUTE OP AERONAUTICS AND ASTRONAUTICS PAPER 2001-1793


Both vehicles are completely identical.Each vehicle creates enough thrust for itself, there-

fore there is no Any changes from the The engines ofall stages are ignited at take-off simultaneously andeither of the two vehicles can alternately be used asbooster and as orbiter. By this the overall vehicledegradation will be reduced, because during boosteroperation the vehicle experiences shorter operationof the rocket engine. Further more the aerodynamicpressure and heat loads are significantly reduced. Inaddition the vehicle bending stiffness will be increasedfor the connected vehicles, which in turn is likely toreduce the dynamic and aerodynamic bending loadsin the structure.

M=0.6 (FSS-4)M=1.2(FSS-4)M=3.0 (FSS-4)

M=0.6 (Hope-X)M=1.2 (Hope-X)M=3.0 (Hope,X)

M=0.6M=1.2M=3.0

5 10 15 20 25 30 35 40AOA [deg]

Fig. 9 Vehicle lift-over-drag characteristics

Aerodynamic Characteristics

To define the aerodynamic characteristics a rel-atively blunt body-wing configuration is assumed.The maximum lift-to-drag ratio in the subsonicregion (Mach 0.6) is about 5, in the transsonic region(Mach 1.2) about 3 and at Mach 3.0 about 2. Thiscan be seen in figure 9 showing the lift-to-drag ratioplotted over the angle-of-attack for the flight Machnumbers of 0.6, 1.2 and 3.0. This aerodynamiccharacteristics are similar to those of the JapaneseHope-X spaceplane5 or the FESTIP study conceptnumber 4 (FSS-4),6 which are also shown in figure 9.

The low lifting performance at super- and hyper-sonic flight speeds is a trade-off to obtain a moresimple, compact and light-weight structural design.

The aerodynamic characteristics are definedthrough the lift- and drag coefficients CD and CL,which are described by an interpolation functionwhich depends on the flight Mach number and theangle of attack.

Siamese Twin withFuel-Cross Feeding

In the last years a strong interest for single stageto orbit vehicles could be observed. The challenge ofsuch concepts are the potential for lower developingcost (only one vehicle) as well as lower operationalcost.

The technological maturity required for such vehi-cles however is critical, at least for many space faringnations. With a Siamese concept, the benefits of asingle vehicle development can be combined with thebenefits of a two stage concept.

The Siamese twin concept is basically a parallellaunched SSTO vehicle - see vehicle (C) in figure 7.

The main modifications from a true SSTO vehicleare as follows:

• Fuel cross feeding from the booster to the orbiter

• Stage separation in an approximately ballisticflight phase

• Fuel container in the booster vehicle payload bayfor rocket powered booster return

Because of the parallel launch configuration ofthe Siamese twin concept, only vertical takeoff wasassumed in this study.

Rocket-Powered Booster Return to Launch SiteTo achieve lowest possible operational cost both

vehicles should autonomously return and land atthe launch site. For the orbiter a timely ignitionof the orbital re-entry engines and the use of theaerodynamic lifting capabilities are assumed to besufficient to return to launch site.

For the booster vehicle figure 10 shows a possiblereturn to launch-site. The flight speed, altitude anddynamic pressure are plotted over ground-distance.Only a pure rocket powered return to launch site hasbeen assumed in this study.

The return sequence in figure 10 starts at thesquared marks at an altitude of 26.6 km, at a flightspeed of Mach 4 and a flight path angle of 25deg.After the main engine shutdown the vehicle coaststo an altitude of about 45 km. At this time thespeed has reduced to about Mach 3. At a dynamicpressure of 3 kPa the vehicle turns around ISOdegwith help of the attitude control system and re-ignites the main propulsion system at 1/3 of thetake-off thrust. The burn lasts for about 60 secand reverses the flight speed to about Mach 1.3in the opposite direction (back to the launch site).The altitude remains around 45 km. Now the vehi-cle re-enters into the dense atmosphere, picks up an

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60 30

22.5

7.5- 1 -

0J

-60 -40 -20 0 20 40 60 80ground distance [km]

Fig. 10 Siamese twin rocket powered booster re-turn

800

700

7T 600

-• 500

400

300

0 km5 km

10km15km20km25km30 km

2 3 4speed [Mach]

Fig. 11 Specific Impulse of ejector rocket mode

optimum flight path and glides back to the launch site.

If a booster dry mass of 18 % of GTOM is assumed,the amount of fuel required for this return maneuveris 7 % of the GTOM. Therefore booster separationmust be triggered when the booster mass reaches25 % of its GTOM. With these assumptions, theorbiter vehicle reaches the target orbit with 20.5 %GTOM (see figure 2). With the dry mass assump-tion of 18 % the payload mass fraction would be 2.5 %.

Trajectory Optimization

For the trajectory optimization a robust combinedGenetic Algorithm Downhill Simplex hybrid optimizerhas been developed by Shepperd.7 First the GeneticAlgorithm identifies the global optimum, which is thenrefined by the Downhill Simplex method.

Results & ConclusionsEjector Performance

In figure 11 the specific impulse of the rocket ejec-tor is plotted over the flight speed range for differentaltitudes. It is seen, that for every altitude there isa speed range, at which the engine can be operatedclose the the performance optimum. Therefore the as-cent trajectory has to be optimized in order to obtainbest performance.

Vehicle Performance

Figure 12 shows a typical vehicle ascent of a Single-Stage-To-Or bit Ejector Ramjet powered vehicle. Theeffective specific Impulse Ieff (the net acceleratingthrust divided by the fuel massflow)8 indicates thetrue engine performance. The difference between thespecific Impulse and the effective specific Impulse isclearly seen at the end of the ramjet operation. Atthis point the ramjet thrust provides only little accel-eration, therefore the fuel is not used effectively (seefigure 12). If the effective specific impulse drops below

60

50

40

^ 30

20

10

Ariane 4, altitude ——booster separation ostage 1. separatipn _ A

6000

5000

4000

3000 £L

2000

1000

4 6flight speed [Mach]

10

Fig. 12 Horizontal takeoff ejector-ramjet trajec-tory

the rocket only levels, the engine switches to rocketmode.

Table 2 and figure 16 summarize the key results ofthis paper. 20 different combinations of vehicle and en-gine concepts are shown. They can be roughly brokeninto 4 groups corresponding to the following vehicleconcepts: ballistic SSTO, vertical takeoff SSTO, hori-zontal takeoff SSTO and a Siamese twin vertical takeoffconcept. For all vehicle concepts

Vehicle Comparison

Table 2 and figure 16 show the key data of allsimulated concepts.

Figure 13 compares the ballistic SSTO, verti-cal take-off SSTO, horizontal take-off SSTO andvertical take-off identical TSTO each with rocket,ejector-rocket and ejector-ramjet propulsion system.Figure 13 shows the relative change compared tothe rocket propelled, ballistic SSTO. Most remark-able is the comparably good performance of theejector-rocket. The ejector-ramjet can increase the

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I-TSTO

Fig. 13 Absolute change of dry mass

SSTOBLST

SSTOVT

SSTO I-TSTOHT


5, 9, 15, 19 and 20 in table 2) are interpreted thatthe ramjet engine does not match well with the pooraerodynamic characteristics of the vehicle.

Ejector RocketThe match between propulsion system and vehicle

for the ejector rocket seems much better. Despitethe much lower engine Isp, the overall performanceof the ejector rocket vehicles is comparably high andsometimes achieves values close and equal alues thanthose of the ejector ramjet vehicles.

Not only will the engine be lighter (because sim-pler), there is also a good potential, that the rocketengine combustion pressure and temperature canbe reduced. This would significantly enhance thedevelopment of reusable rocket engines and also makethem more reliable

Finally the installed rocket thrust can be reducedby about 15% due to the thrust augmentation in theejector. This in turn will reduce the system mass andmaybe number of engines.

Horizontal versus Vertical TakeoffNo clear performance increase could be identified

between vertical and horizontal takeoff. Horizontaltakeoff requires larger wings, which are a penaltyduring the high speed phases of the ascent.

From a structural point of view, the mass forlarger wings as well as the structural requirementsto support the higher bending loads will more thenoffset the little (or zero) gain in mass benefits.

In addition, significant problems arise from theunder-carriage. Not only will its mass be significant(is has to support the full GTOM), also the takeoffspeeds are very high because enough dynamic pres-sure needs to be built up. If the wing-size would beenlarged to meet take-off speed restrictions, the highspeed penalty would be even larger.

This means, for a vehicle with poor aerodynamicperformance such as assumed in this study, horizontaltakeoff shows no clear benefit.

Siamese twinThe Siamese twin concept increases the vehicle

performance for all investigated engine types. Fur-thermore the gains in vehicle performance for theejector-rocket as well as ejector-ramjet engines arelarger than for the single stage vehicles. This makesthe practical implementation more realistic.

Fig. 14 Relative change of dry mass

25 T—

n• EJ Mach4• EJ Mach6

SSTOFig. 15 Comparison ejector shutdown speed,Mach 4.0 vs. 6.0

performance only in case of the TSTO vehicle.

Figure 15 shows the difference between the ejectorrocket operation until Mach 4 versus Mach 6, forthe SSTO and TSTO vertical takeoff concepts. Therelative increase of performance is about twice ashigh, if the ejector ramjet is operated until Mach 6.

DiscussionEjector Ramjet

The ejector ramjet combined cycle shows an increaseof performance (mass in orbit). But the small increasein propellant mass fraction make it appear difficultto accommodate the additional mass for inlet, ramjetand nozzle. The early ramjet shutdowns (concepts 4. The Siamese twin concept allows higher structural

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mass fractions, which will be achievable with lowertechnological development effort and cost.

There is also more margin to optimization thedesign for operational aspects as opposed to theperformance driven design of single stage vehicles.This will lower the operation cost.

Neither fuel cross feeding nor stage separation,which can be performed during a temporary ballistic(non-lifting) flight phase, which resembles the separa-tion of today's rocket boosters, should be fundamentalobstacles.

The launch abort capabilities for the Siameseconcept do also look favorable. In case of earlyengine failure, both vehicles could perform the rocketpowered return maneuver (joint together), performthe stage separation before the re-entry into denseatmosphere an make simultaneous unpowered land-ings on parallel or nearby runways. If geographicallypossible both vehicles can also perform unpowereddown-range landings. In case of a late engine failurean escape into a low orbit once around Earth wouldbe an option. Finally, for manned missions thecrew-container carried inside of the cargo bay couldbe ejected by rocket thrusters and rescued withparachutes.

ConclusionsSingle stage vehicle concepts using ejector rocket

and ejector ramjet engines show only a small increasein overall vehicle performance. The use of a ramjetdoes show no clear benefit. The ejector rocket has rea-sonably good performance and appears to be worth tobe looked at in more detail. Especially a clearer esti-mation of the increase of system mass will be required.

A additional, important benefit of the ejector rocketis, that the losses of a rocket engine due to the atmo-spheric back-pressure at low altitude can partially berecovered. This makes the system performance lessdependent on the rocket engine performance and al-lows to reduce the rocket engine combustion pressureand temperature in order to optimize the rocket en-gine designs for robustness, reliability and re-usability.

The Siamese twin concept increases the vehicleperformance significantly. If a rocket powered verticaltakeoff vehicle which is capable to fly SSTO missions,would be launched in Siamese twin configuration,the payload mass would increase by a factor of~8 and by far offsets the only double operationalcost. This increase in performance is of coursemainly due to the two-stage design. However, theejector-ramjet and ejector-rocket can both increase

the system performance to an extent, which makestheir practical implementation more realistic. For thereason of relaxing the high rocket system requirementsas mentioned in the previous paragraph, the imple-mentation of an ejector-rocket appears very attractive.

AcknowledgementsThe author greatly thanks the Japanese Ministry of

Education, Culture, Sports, Science and Technology(MEXT, formerly STA) for his fellowship, as wellas Dr. Malta for his generous support and helpfuldiscussions.

ReferencesFoster, R., Escher, W., and Robinson, J., "Studies of an Ex-

tensively Axisymmetric Rocket Based Combined Cycle EnginePowered SSTO Vehicle," AIAA, , No. AIAA 89-2294, July 1989.

2Hill, P. and Peterson, C., Mechanics and Thermodynamicsof Propulsion, Addison-Wesley, 2nd ed., 1992.

3Kerrebrock, J., Aircraft Engines and Gas Turbines, MITPress, 2nd ed., 1992.

4Reiser, W. and Pratt, D., Hypersonic Airbreathing Propul-sion, AIAA Education Series, AIAA, 1994.

5NAL,NASDA, NAL/NASDA, Hope-X Aerodynamic DataBook, FY09 1/3, March 2000.

6ESA, ESTEC, FESTIP Yearly Workshop Summary Hand-out, FESTIP Yearly Workshop, Oct. 1996.

7Shepperd, R., "System Optimization of an Ejector RamjetRocket Based Combined Cycle Model," Nal report, NAL, Na-tional Aerospace Laboratory, May 2000.

8Escher, W., Hyde, E., and Anderson, D., "A User's Primerfor Comparative Assessments of All-Rocket Based Combined-Cycle Propulsion systems," AIAA Paper AIAA 95-2474, 1995.

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ORBIT: 250 km, 29°inclination, launch site at 10°latitude, GTOM = 100 Mg

Q)oE Engine Typez

ballis

tic

SSTO

g ae

rody

nam

ic lift

c'c/3

IP =

IP

Ver

tical

Tak

eoff

Hor

izon

tal

Take

off

1-

1 rocket2 ejector rocket, Mach4 *3 ejector rocket4 ejector ramjet5 ejector ramjet, Ar20*3

e rocket7 ejector rocket, Mach4 *s ejector rocket9 ejector ramjet10 ejector ramjet, Ar20*3

1 1 ejector ramjet, q90*4

12 rocket13 ejector rocket, Mach4 *14 ejector rocket15 ejector ramjet16 rocket17 ejector rocket, Mach4 *1 s ejector rocket19 ejector ramjet, Ar20*3

20 ejector ramjet, q90*4

A T-K, W 'sp 'sP EJorEJR Av A PMF Dry Massmlet /GTOW /GTOW (SL) (max) shutdown Mma* 3P° ^ *5 (1-PMF)

[rrf] - [-] [-] [sec] [sec] [Mach] [kPa] [m/sec] [m=] [%] [%]1.20 - 355 448 - 37 50 60 87.2 12.8

10 1.20 1.43 423 660 4.0 '2 60 55 60 85.1 14.910 1.20 1.43 423 710 6.0 60 49 60 84.1 15.910 1.20 1.43 423 4570 4.5 *1 60 89 60 84.4 15.620 1.20 1.43 423 4570 2.8 *1 46 74 60 85.0 15.0

1.20 - 355 448 - 60 84 60 86.8 13.210 1.20 1.43 423 650 4.0 '2 60 71 64 85.0 15.010 1.20 1.43 423 710 6.0 60 113 64 84.0 16.010 1.20 1.43 423 4570 4.7 *1 60 105 110 84.2 15.820 1.20 1.43 423 4570 6.0 60 66 64 81.9 18.110 1.20 1.43 423 4570 6.0 90 74 88 82.8 17.2

0.9 - 355 448 - 50 49 115 87.0 13.010 0.8 0.95 423 790 4.0 '2 56 48 142 85.2 14.810 0.8 0.95 423 790 6.0 41 51 150 84.1 15.910 1.2 1.43 423 4570 4.8 60 46 199 82.9 17.1

1.2 - 355 448 - 60 59 62 79.5 20.510 0.9 1.07 423 700 4.0'2 60 51 77 77.7 22.310 0.9 1.07 423 710 6.0 60 44 63 75.6 24.420 0.9 1.07 423 4570 4.7 *1 60 63 115 74.5 25.510 0.9 1.07 423 4570 4.7 *1 90 58 98 75.4 24.6

early ramjet shutdown due to low accelerationejector-rocket limited to Mach 4.0Ramjet reference area = 20 m2

dynamic pressure = 90 kPaPMF ... Propellant Mass Fraction

Table 2 PMF for different vehicle and engine concepts

1 - PMF [%]

c/)C/)

CD

V)

oto03

JD

O

03>^IDO

O)C

"c/313

OCD

03O

CD

03 H_C *6O CD

1 rocket2 ejector rocket3 ejector rocket4 ejector ramjet CC5 ejector ramjet CC, Ar20

16.0

6 rocket7 ejector rocket, Mach48 ejector rocket9 ejector ramjet CC10 ejector ramjet CC, Ar2011 ejector ramjet CC, q90

18.117.

12 rocket13 ejector rocket, Mach414 ejector rocket15 ejector ramjet CC16 rocket17 ejector rocket, Mach418 ejector rocket19 ejector ramjet CC, Ar2020 ejector ramjet CC, q90

0.0 2.5 5.0 7.510.012.515.017.520.022.525.027.5Fig. 16 1-PMF for different vehicle and engine concepts

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