AIAA-2001-1838 Design Detailed d' Azur

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(c)2001 Ame rican Institute of Aeronautics & Astronautics or Published with P ermission of Author(s) and/or Author(s)' Sponso ring Organization.

AMA A A01-28133

AIAA 2001-1838Analytical Study of Pre-CooledTurbojet Engine for TSTO SpaceplaneH. TAGUCHI, H. FUTAMURA,R. YANAGI and M. MAITANational Aerospace LaboratoryTokyo, JAPAN

AIAA / NAL-NASDA-ISAS10th International SpaceHypersonic Systems andConference24-27 April 2001 Kyoto, Japan

Planes andTechnologies

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


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( c )2 0 0 1 A m e r i c a n I n s ti tu t e o f A e r o n a u t ic s & A s t r o n a u t ic s o r P u b l is h e d w i th P e r m i s s i o n o f A u t h o r ( s ) a n d / o r A u t h o r ( s ) ' S p o n s o r i n g O r g a n i z a t io n .

Analytical Study of Pre-Cooled Turbojet Engine for TSTO SpaceplaneHideyuki TAGUCHI*, Hisao FUTAMURA*, Ryoji YANAGI* and Masataka MAITA**

National Aerospace LaboratoryChofu, Tokyo 182-8522, Japanese

Phone: +81-422-40-3481, Fax: +81-422-40-3446, E-mail: [email protected]

ABSTRACTAuthors are investigating a concept of pre-cooled turbojet engine for Two-Stage-To-Orbit (TSTO)

spaceplane. In this study, the engine system is designed with some assumptions. Technological problemsand operational limits of the engine are clarified by an engine performance analysis. Then, flightperformance of the engine is compared with some air-breathing engines such as: pre-cooled air turbo ramjet,turbo-ramjet and rocket-ramjet Sizes and mass of engines are estimated by analogical estimations with useof a database of existing engines. As a result, pre-cooled engines brought better payload performance thanother engines.

INTRODUCTIONVarious concepts of air-breathing propulsion

system have been studied for next generation spacelaunchers. Th e engines must operate from Mach 0to about Mach 6. The weight of the engine shouldbe light, because the weight is just a penalty formissions. The specific impulse of the engineshould be large, because the propellant weight usedat accelerating phase is also a penalty for missions.Recently, following accelerator engines are reportedwith realistic design features: Ejector Ramjet1^Turbo Ramjet2'3), Air Turbo Ramjet4), Deep CooledTurbojet5'6), etc.

Authors selected pre-cooled turbojet as anaccelerator engine. Figure 1 shows the conceptdrawing of proposed TSTO. First stage hasair-breathing engines. Second stage hasconventional rocket engine. Operating speedranges of each engine mode are assumed as follows:pre-cooled turbojet engine is used between Mach 0*Aero-propulsion Research Center* * P l a n n i n g a n d Management O f f i c eCopyright 2001 by the Am eric an Institute ofA e r o n a u t i c s and Astronau tics, Inc.A l l rights reserved.

and Mach 6, rocket engine is used over Mach 6.Figure 2 shows the cross section of the

pre-cooled turbojet engine. The incoming air iscooled at pre-cooler by cryogenic fuels like liquidhydrogen. The air density increase by coolingbrings larger airflow rate. The temperature dropbrings less compressor load. Then, the engineproduces larger power compared to normal turbojetengines. In this concept, liquid hydrogen fuel issupplied to the pre-cooler in order to attain low airtemperature, and the fuel is injected into bothmain-burner and after-burner.

Variable shape nozzle is assumed. Gas flowrate is controlled by throat area. The engine alsooperates with lean combustion at landing phase toobtain high specific impulse. The pre-cooledturbojet engine can achieve high thrust to weightratio and high specific impulse by controlling theequivalence ratio.

In the former study?\ general performanceanalysis of the pre-cooled turbojet was conducted.As results, it was confirmed that the pre-cooledturbojet engine with fuel rich operation produceshigh thrust to weight ratio and can be adopted toSingle-Stage -To-Orbit (SSTO) spaceplane.

In this paper, conceptual design of "the


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Pre-cooled Turbojet Engine for TSTO spaceplane ismentioned. Shape and dimension of eachcomponent in the sub-scale engine are created in theconceptual design. Off-design performanceanalysis of the pre-cooled turbojet is carried out.Finally, flight analysis of the engine with anassumed mission was carried out.

CONCEPTUAL DESIGNEngine System

The system diagram of pre-cooled turbojetcombined engine is shown in Fig. 3. Liquidhydrogen is selected as a propellant. Liquidhydrogen is supplied to the engine through thepre-cooling heat exchanger. A part of liquidhydrogen is supplied to the after-burner wall. Thehydrogen absorbs the heat, and drives the liquidhydrogen turbo pump.Air Intake

Two-dimensional bifurcated intake is selected.Required capture area greatly changes, since theflight Mach number varies from 0 to 6. Thecapture area is adjusted by varying the angle of thecowl. Maximum capture area is decided by frontface area of the pre-cooling heat ex changer.Pre-cooling Heat Exchanger

Two-dimensional pre-cooling heat exchangeris obliquely installed in a subsonic diffuser. Theflow velocity of the air is set as 25~30m/s. Theheat exchanger is made with the shell and tubesystem with fins.Turbo-machinery

Cross section of the turbo-machinery of thepre-cooled turbojet engine is shown in Fig. 2.Six-stage axial compressor with pressure ratio of 10and two-stage turbine with pressure ratio of 3 areselected. Normal materials, such as titanium fo rcompressor and nickel-based materials for hot

section are assumed .Nominal combustion temperatures of the

main-burner and after-burner are set about 1700Kand 2000K, respectively.After-Burner

Duct part in the turbojet downstream isequipped with fuel injectors for after-burning.Variable geometry nozzle is attached to theafter-burner. Exit area of the nozzle is assumedsmaller than the intake capture area.

PERFORMANCE ANALYSIS OFPRE-COOLED TURBOJET

Analysis ConditionsPerformance of the pre-cooled turbojet in the

real flight environment is analyzed. Hydrogen isassumed as fuel. The analysis is carried out withthe system diagram of Fig. 3. Thermo physicalproperties of hydrogen and air are calculated usingapproximate function obtained by Ref. 8.Heat balance calculation is executed bydesign calculation method of Ref. 9. Operatingmap of the compressor is assumed based on existingcompressor. Th e equivalence ratio is controlled tomaintain fixed combustion temperature.Rotational speed, passage area of each componentand fuel (coolant) flow rate are varied in order tosatisfy the equation of shaft power balance, flowrate balance and heat balance in each flightcondition. The convergent calculation is carriedout.Analytical Resultsa. Flow Rate an d Equivalence Ratio

Figure 4a) shows flow rate and equivalenceratio. The equ ivalence ratio is controlled to keeptemperature of main-burner (T4) constant.Equivalence ratio (0) increases when th e Machnum ber rises. Air mass flow rate decreases under


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high Mach number condition because compressorinlet temperature is increased.b. Temperature

Figure 4b) shows temperature of eachcomponent. Intake ex it temperature (Tl) increaseswith the increase of the Mach number. Therefore,heat-resistance material is required at intake andpre-cooling heat exchanger section. However,compressor outlet temperature (T3) is below 600Kin the whole flight region. The temperature isequivalent to that of turbojet engines for Mach 3class supersonic airplane. Turbine inlettemperature (T4) is kept nearly 1700K bycontrolling equivalence ratioc. Pressure

Figure 4c) shows pressure of each component.Compressor exit pressure (P3) takes the maximumvalue (about 1.4MPa) at Mach 6. Compressorpower decreases with temperature drop bypre-cooling. This brings high turbine exit pressure(P5) such as about l.OMPa.c L Passage Area

Figure 4d) shows passage area of eachcomponent. F or taking the flow rate matching ofintake and engine, intake throat area (Ait) varies.Effective capture area (Acef) with Mach 0.3 ~4 issmaller than maximum capture area (Acmax). Thearea is decided by front face area of the pre-cooler.The whole momentum of the spillage flow isconsidered as the resistance in this region. Acef atmaximum rotation speed is larger than Acmaxbeyond Mach 4. Then, rotation speed is limited inthis region.

The flow rate mismatching of compressor andturbine occurs when the compressor inlettemperature rose and the combustion temperature iskept constant. Since this mismatching is largerthan normal turbojet, turbine stator vane is assumedto move and adjust turbine throat area (Att).Operating line on the compressor map can be set to

pass through the maximum efficiency line. Nozzlethroat area (Ant) of after-burner should be slightlyvaried for taking the flow rate matching of engineand nozzle.

Nozzle exit area (Ane) is set to be smallerthan Acmax. Ac beyond Mach 1.5 limits Ane.The exhaust gas is under expansion in this region.e * Rotation Speed and Shaft Power

Figure 4e) shows analytical result of enginerotation speed and shaft power. Mechanicalrotation speed (N) is decided by the upper limit ofcorrected rotation speed (Nc) in the Mach rangebetween 0 and 2. This is because sound velocitydrops by the pre-cooling. Compressor inlettemperature rises over the design point in the Machrange between 2 and 3.5. N is limited by amechanical limit in this region. Acmax beyondMach 3.5 limits air flow rate of intake. N islimited to match the air flow rate of intake andengine.

MISSION ANALYSISMission performance of air-breathing

propulsion systems are analyzed by simplifiedmethod in order to make a fair comparison.Discussion of absolute values of payload andcomponents make little sense, because the valueschange sensitively with assumptions of material,aerodynamic performance, etc. Objective of thissection is the qualitative comparison of propulsionsystems with a same assumption.Engine Systems for Comparison

Table 1 shows analysis condition of enginesystems. Calculation is conducted from take off toMach 6 with dynamic pressure of 34kPa.

Figure 5a) is the engine system diagram ofpre-cooled air turbo ramjet (PCATR). Institute ofSpace and Astronom ical Science carry out firing testof pre-cooled air turbo ramjet (ATREX)4). Th e


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engine cycle under sea level static condition hasbeen verified. The fan of the engine is driven byhot hydrogen gas, which is generated by heatexchanger installed in the main combustor.Figure 5b) is the engine system diagram ofturbo ramjet (TJ+RAM). Turbo ramjet with designoperating range of Mach 0 to 5 was developed inresearch and development project ofsuper/hyper-sonic transport propulsion systems(HYPR)3) by Ministry of International Trade andIndustry. The engine operation under conditionbetween Mach 0 and 3 has been verified by highaltitude test.Engine Performances

Figure 6 shows analytical results of engineperformance analysis of the engines.

Specific impulse (Isp) of TJ+RAM is higherthan that of PC ATR and PCTJ. This is because thecompressor exit temperature is high and theequivalence ratio is relatively low with samecombustion temperature limit. Thrust / capturearea of PCATR and PCTJ is higher than that ofTJ+RAM. This is because pre-cooling increasesthe mass flow rate of air.

Differences between PCATR and PCTJ areprimary brought by pressure ratio. Thetemperature of hot hydrogen at the turbine inletlimits the pressure ratio of PCATR.Flight Analysis Condition

Figure 7 shows the force diagram of the flightanalysis. Calculation procedure and aerodynamicdata is the same as Ref. 10. The vehicle is treatedas a mass point. The motion is fixed at theequatorial plane, and the vehicle moved in thedirection of east. Thrust, gravity force at thealtitude, centrifugal force, lift and drag is involvedin the equations. Angle of attack and thrust arecontrolled to attain the designated flight pass.

Table 2 shows the flight analysis conditions.Th e flight analysis is carried out using engineperformance m entioned above, Liquid hydrogen is

used as fuel. Initial mass of 350Mg and dynamicpressure of 34 kPa are assumed. Front area of theengine is decided by the results of the performanceanalysis at sea level. Maximum thrust at eachflight speed is calculated using approximatefunction of the performance analysis results.Mass Estimation

Table 3 shows the mass estimation conditions.Mass of airframe parts is estimated using Ref. 11.Wing size is determined by total take off mass.Wing loading is assumed as 500kg/m2. Mass oftanks is estimated by design calculation.Composite material with safety coefficient of 4.5 isassumed as the materials for Liquid hydrogen tank.Aluminum alloy with safety coefficient of 1.5 isassumed as the material of liquid oxygen tank.Mass of pre-cooler is estimated by a design analysismethod, which is evaluated by experiments usingscaled models. Mass of the turbojet engine isestimated by actual engine trends, using function ofpressure ratio and volume flow rate. Mass oframjet is estimated by a duct design analysis withCarbon/Carbon panels. Mass of rocket engine isestimated using data of LE-7A engine12 ).Mass Ratio

Figure 8 shows the comparison of calculatedpayload. First stage engine mass with turbo-basedengines is very large. However, payload mass withturbo-based engines is larger than that of rocket-based engines. This is because the first stagepropellant mass of those is very small.

First stage engine mass of PCATR and PCTJare lower than that of TJ+RAM. Then, it isestimated that the pre-cooled engines bringscomparatively large payload than other enginesconsidering fly back operation. However, furtherstudy with precise mass estimation and missionanalysis is necessary for selection withinturbo-based engines. Figure 9 show the airframeconfiguration with PCTJ, obtained by abovementioned mission analysis.


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CONCLUDING REMARKSPre-cooled turbojet engine for TSTO

spaceplane is designed and its features are clarifiedby analyses. Followings are confirmed as results.1) Pre-cooled turbojet engine cycle give enough

performance for TSTO spaceplane.2) Exiting compressor structure can be used to the

pre-cooled turbojet engine.3) Payload mass of TSTO spaceplane with

turbo-based engines is larger than that ofrocket-based engines.4) F urther study with precise mass estimation and

mission analysis is necessary for selectionwithin turbo-based engines.

REFERENCES

1. Escher, W. J. D. et. al., "Synerjet forEarth/Orbit Propulsion: Revisiting the 1966NASA/Marquardt Composite(Airbreathing/Rocket) Propulsion SystemStudy (NAS7-377)," AIAA-96-3040.

2. Lanshin, A. I. et. al., "Russian AerospaceCombined Propulsion Systems Research andDevelopment Program ("ORYOL-2-1"):Progress Review," AIAA-96-4494.

3. Mitsuoka, T. et. al., "Research andDevelopment of Combined Cycle EngineDemonstrator," Third International Symposiumon National project for Super/Hyper-sonicTransport propulsion System, pp. 229-236,1999

5. Balepin, V. V. et al., "KLIN Cycle : CombinedPropulsion for Vertical Take Off Launcher,"AIAA 97-2854.

6. Taguchi, H. et. al., "Conceptual Study ofPre-Cooled Air Turbojet / Rocket Engine withScramjet (PATRES)," ISAB E 99-7024

7. Taguchi, H. and Yanagi, R., "A Study onPre-Cooled Turbojet - Scramjet - RocketCombined Engines," AIAA-98-3777.

8. JSME Data Book: "Thermo physical Propertiesof Fluids (Japanese)," The Japan Society ofMechanical Engineers, 1983.

9. JSME Data Book: "Heat Transfer 4th Edition(Japanese)," pp.253-256, The Japan Society ofMechanical Engineers, 1994.

10. Kanda, T. and Kudo, K , "Payload to Low EarthOrbit by Aerospace Plane with ScramjetEngine," AIAA J. Propulsion, Vol. 13, No. 1,pp. 164-166

11. Harloff, G. J. and Berkowitz, B. M,"HASA-Hypersonic Aerospace SizingAnalysis for the Preliminary Design ofAerospace Vehicles," NASA CR-182226,1988.

12. Fukushima, Y. et. al., "Development Status ofLE-7A and LE-5B Engine," Proceedings of the38 th Conference on Aerospace Propulsion(Japanese), pp.27 6-281 , 1998

4. Tanatsugu, N. et. al., "Development Study onATREX Engine," IAF-94.S.4.428.


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Fig. 1 TSTO Spaceplane Concept

Turbo-pump

Air Intake Pre-Cooler Core E ngine Nozzle

Fig. 2 Cross Section of Pre-Cooled Turbojet Engine (PCTJ)LH2Pump

Fig. 3 System Diagram of Pre-Cooled Turbojet Engine (PCTJ)6


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Table 1 Engine Analysis Conditions

PCTJ PCATR TJ+RAMDynamic Pressure [kPa]:Compressor Pressure Ratio:M ain Burner Temperature [K]:After/Ram Burner Temperature [K]:

341016732273

343.3-2400

341020002400

LH2 PumpLH2 Pump

a) Pre-Cooled Air Turbo Ramjet (PCATR) b) Turbo Ramjet (TJ+RAM)Fig. 5 Engine System Diagrams

130

"o0}Q.CO

-PCTJ

7000600050004000300020001000

-PCATR -O-TJ+RAM

2 3 4Flight Mach Number

E-*

+Q.COO

-PCTJ -PCATR -o- TJ+RAM

2 3 4Flight Mach Number

Fig. 6 Comparison of Engine Performances


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Table 2 F light Analysis Condition

Fig. 7 F orce Diagram of F light Analysis

Table 3 M ass Estimation ConditionBase: HASAWing Size: Estimated by Take-off MassWing Loading: 50 0 kgf/m 2Airframe, LH 2 Tank:Composite (Safety coefficient = 4.5)LO X Tank: Alminum (Safety coefficient = 1.5)Turbojet Engine: Estimated by Trend Curve (n, Q)Ramjet Engine: C/C Panels and BeamsRocket Engine: Estimated by LE-7A

Dynamic Pressure:Fuel-Initial Mass:Initial Acceleration:Mach Number :

350

300

250

200 ~150

100

34kPaLH 2350 ton0.5G0-6

DPayload

E 3 2nd Wing,Gear2ndStructure2ndEngneM 2nd Propellantn 1 st Wing, Gear

1 stStructureM 1 st Engne

E31st Propellant

PCTJ PCATR TJ+RAM RE+RAM RE

Fig.8 Comparison of Mass Ratio

70m

22m

A\ _ ^ c x _ .OO~Dv YCOOFig. 9 Airframe configuration

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AIAA-2001-1838 Design Detailed d' Azur