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Acta Astronautica

Acta Astronautica 66 (2010) 1169–1176

0094-57

doi:10.1

� Cor

E-m

journal homepage: www.elsevier.com/locate/actaastro

Development study of a precooled turbojet engine

Tetsuya Sato a,�, Hideyuki Taguchi b, Hiroaki Kobayashi b, Takayuki Kojima b,Katsuyoshi Fukiba b, Daisaku Masaki, Keiichi Okai b, Kazuhisa Fujita b, Motoyuki Hongo b,Shujiro Sawai b

a Department of Applied Mechanics and Aerospace Engineering, WASEDA University, Japanb Japan Aerospace Exploration Agency, Japan

a r t i c l e i n f o

Article history:

Received 26 January 2009

Received in revised form

27 July 2009

Accepted 5 October 2009Available online 20 November 2009

Keywords:

TBCC

Precooled turbojet

Flight experiment

65/$ - see front matter & 2009 Elsevier Ltd. A

016/j.actaastro.2009.10.006

responding author. Tel./fax: þ81 3 5286 2492.

ail address: sato.tetsuya@waseda.jp (T. Sato).

a b s t r a c t

A precooled turbojet engine has been developed by JAXA used for the hypersonic

airplane and spaceplane. The subscale engine named ‘‘S-engine’’ whose thrust and

weight are about 1.2 kN and 100 kg was designed, fabricated and tested. The

components and the system firing tests under the sea-level-static condition were

successfully conducted.In the next phase, a flight test of the S-engine is planned using a

stratospheric balloon in 2010 called balloon-based operation vehicle (BOV). The vehicle

is dropped from an altitude of 40 km by a high altitude balloon. After 40 s free-fall, the

vehicle is pulled up and the S-engine operates for 30 s at about Mach 2. High-altitude

tests of the core-engine verified the performance and healthiness of the engine under

the condition corresponding to the BOV flight trajectory.

& 2009 Elsevier Ltd. All rights reserved.

1. Introduction

Innovation of the propulsion system is required torealize the hypersonic airplane and/or the low cost, highreliability and routine access to space. We proposed theprecooled turbojet (PCTJ) engine, which is a kind ofturbine-based combined cycle (TBCC) engines, as thereusable hypersonic propulsion system. This engine worksfrom takeoff to Mach 6 using liquid hydrogen as the fueland coolant of its precooling system.

The development study of PCTJ started in 2004 [1],reflecting the key technologies of the ATREX engine(expander cycle air turbo ramjet engine) [2]. A small-scaled model of PCTJ called ‘‘S-engine’’ was designednot only for ground tests but also for flight tests.Specification of the S-engine is listed in Table 1. Theengine has a rectangular cross section of 23 cm�23 cm area, 2.4 m length. The weight is about 120 kg

ll rights reserved.

at present and will be reduced to 100 kg if carbon/carboncomposites are used. This size matches to do firingtests under Mach 4–6 conditions in the ramjet enginetest facility (RJTF) at Kakuda Research Center of JAXAand is also suitable for the flight experiment withreasonable cost. The engine thrust under the sea-level-static condition is 1.2 kN corresponding to a quarter ofthe ATREX engine with 2065 s in specific impulse. Severalrig tests and CFD analyses on the engine compo-nents were conducted to verify the performance andhealthiness [3].

In this paper, the development status of the S-engine issummarized. First, the ground system firing test seriesperformed in October 2007 and March 2008 are shown.These tests verified the engine cycle, component perfor-mances and test sequence. Second, the plan of the flightdemonstration called BOV using a stratospheric observa-tion balloon is shown. The trajectory of the BOV is muchdifferent from that of the practical hypersonic vehicle.Therefore, several preliminary tests under the high-altitude conditions have been conducted to confirm theengine performance.

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Fig. 1. Configuration of the S-engine.

Fig. 2. Work-share style.

Table 1Engine design specifications at SLS condition.

Engine system

Engine length m 2.40

Engine width, height m 0.23

Air flow rate kg/s 1.1

Fuel flow rate kg/s 0.06

Thrust kN 1.2

Specific impulse s 2065

Compressor

Type – Single-stage, diagonal

Tip diameter m 0.10

Rotational speed rpm 80,000

Pressure ratio – 6.00

Efficiency % 81

Material – Ti-alloy

Turbine

Type – Single-stage, reaction

Driving gas – Combustion gas

Pressure ratio – 2.4

Efficiency % 83

Turbine inlet temperature K 1223

Material – Inco

Precooler

Heat exchange area m2 2.64

Number of tubes – 1296

Tube diameter mm 2

Heat exchange kW 120

Material – Stainless steel

T. Sato et al. / Acta Astronautica 66 (2010) 1169–11761170

2. System firing test of S-engine

2.1. Description of the S-engine

The S-engine consists of an air-inlet, precooler, core-engine (compressor, combustor and turbine), afterburnerand nozzle as shown in Fig. 1. The air-inlet hasthree ramps whose second and third ramps aremovable to make appropriate shock patterns due to avariety of the flight Mach number. The prototype modelmade of Al-alloy has no movable mechanism becauseit is used below Mach 2. The precooler is a shell-and-tubetype heat exchanger using fuel liquid hydrogen as thecoolant. The core-engine has severe requirements on itslimited dimensions and weight, as well as on itsperformance. The diagonal-flow compressor, reverse-flow annular-type combustor and single-stage axialflow turbine are selected. The variable geometrysingle-ramp rectangular plug nozzle is connected afterthe afterburner. The throat area of the nozzle is adjustedby moving the pear-shape part. As nozzle and after-burner are exposed to the combustion gas exceeding2000 K in temperature, the wall is regeneratively cooledby the fuel hydrogen and the nozzle tip is made ofcarbon/carbon composites. The hydrogen used forcooling is finally burnt in the afterburner mixed withthe air-rich combustion gas which is generated by thecore-engine.

Details of description and each component testing aregiven in Ref. [3].

2.2. Ground system firing tests

Two series of the system firing tests of the S-enginewere conducted at Noshiro Testing Center (NTC) of JAXA.The first test series named PCTJ-1 was conducted in March2007. In this test, newly manufactured main enginecomponents were checked. The most important objectiveis to make sure of the power balance between thecompressor and turbine. Hence, the ‘‘work-share’’style was selected as shown in Fig. 2. The compressorand turbine are connected to each other mechanically bythe shaft. However, both are aerodynamicallyindependent of each other; that is, the airflow passedthrough the compressor is ejected outside and the drivingair of the turbine is supplied from the external tank.Rotational speed of 70,000 rpm was attained in this series,

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T. Sato et al. / Acta Astronautica 66 (2010) 1169–1176 1171

which is 85% of the design value and the power balance ofthe compressor and turbine was verified.

The second test series PCTJ-2 was conducted inOctober 2007. The overall engine system including theaerodynamic fairing, gas supply system and controlsystem, which are the same as the systems of the flightmodel was built and tested. Carbon/carbon composites

Fig. 3. Flow diagram of th

0

1 104

2 104

3 104

4 104

5 104

6 104

-40 -20 0 20 4

Test_ID: bG

rotational speed

turbine inlet temp.afterburner temp.

Rot

atio

nal s

peed

[rpm

]

star

ter o

n(5

%)

star

ter p

ower

-up

(20

%)

mai

n-bu

rner

on

(GH

2 su

pply

)

Fig. 4. System firin

were applied to a part of the nozzle cowl. Fig. 3 shows theflow diagram on the hydrogen supply system. Fuel liquidhydrogen (LH2) is supplied to two combustion chambers:the core-engine combustor and afterburner.

As for the core-engine combustor system, the hydro-gen fuel weighing about 8 g/s is regulated by the maincontrol valve, vaporizes in the evaporator located down-

e system firing test.

0

500

1000

1500

0 60 80 100 120

P50-20L-02

Tem

pera

ture

[K]

Time [sec]

afte

r-bu

rner

on

(LH

2 su

pply

)sh

utdo

wn

g test result.

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T. Sato et al. / Acta Astronautica 66 (2010) 1169–11761172

stream of the turbine and burns in the combustor with thefuel-lean condition. The fuel in this system is mainly usedfor driving the turbine. Meanwhile, as for the afterburnersystem, liquid hydrogen weighing about 52 g/s is regu-lated by the other control valve, is heated regenerativelythrough the precooler and nozzle wall and burns in theafterburner with fuel-rich condition. The fuel is usedfor cooling down the airflow and hot components as wellas for increase of the engine thrust. The fuel managementof the core-engine combustor system is difficult especiallyat the engine start-up. Fuel must be supplied to thecore-engine to achieve the target rotational speed asquickly as possible without exceeding the turbine tem-perature limit. At the engine start-up, LH2 is vaporizedby the heat capacity of the pipes and valves. As a result,the phase of hydrogen near the control valve isdrastically changed from gaseous phase to liquid phase,which causes the change of the density. It is very difficultto control this drastic change of the density, and the start-up sequence has not been constructed yet in this testseries.

Hence, the gaseous hydrogen (GH2) was supplied tothe core-engine instead of LH2. One of the test results isshown in Fig. 4. First, the compressor is driven up to16,000 rpm by the electric starter, which corresponds to20% of the design rotational speed. GH2 is supplied to thecore-engine combustor to accelerate the turbine whilekeeping the turbine inlet temperature (TIT) below 1000 K(start-up mode: 0–40 s). Then, the engine is accelerated toabout 50% of the design rotational speed by increasingGH2 flow rate while keeping TIT=1100 K (accelerationmode: 40–90 s). After the compressor rotational speedreaches the steady state, LH2 is supplied to the precoolerand afterburner at 90 s. The accuracy of the afterburnertemperature is quantitatively not good. At this time, thecombustion flame of the afterburner is clearly seen asshown in Fig. 5. In this test series, the engine, fuel supply,control and measurement systems were verified. The wayof the start-up sequence was constructed when GH2 wasused for the core-engine. However, the start-up using LH2has not been matured yet.

Fig. 5. S-engine firing test.

3. Engine flight demonstration and preliminaryexperiments

3.1. Engine flight demonstration by balloon-based operation

vehicle (BOV)

The engine flight demonstration using a stratosphericobservation balloon called balloon-based operation vehi-cle (BOV) is planned in 2010 as a low-cost flight way ofthe S-engine. Originally, BOV was proposed by ISAS/JAXAas a test bed for micro-gravity experiments using a free-fall flight produced by the balloon. Two flight experiments(BOV-1 and BOV-2) were successfully conducted in May2006 and June 2007. More than 30 s of micro-gravityconditions were attained in these tests.

The configuration of the BOV-3 which is the flight testbed of S-engine is much different from others. Aschematic drawing of the vehicle installing the S-engineis shown in Fig. 6. The vehicle made of CFRP isapproximately 4.6 m in length, 0.55 m in diameter andmore than 500 kg in weight. The vehicle is separatedfrom the balloon at an altitude of 40 km. MaximumMach number is 2.0 and the duration of the enginetest is about 1 min. Details of the BOV are described inRefs. [3,4].

The flight condition of the BOV-3 is different from thatof the practical vehicle as follows: first, the engine start-up time must be shorter because of the short test durationof BOV, second, the flight conditions on the velocity andatmospheric pressure change drastically. The engineignition must occur under the quite small dynamicpressure and the fuel flow rate should be changeddrastically. Hence, some preliminary experiments of thecore-engine have been conducted under the low ambientpressure conditions corresponding to the BOV-3 flight.

The compressor rig test was conducted at the low-speed wind tunnel of ISAS/JAXA [3]. The ambient pressurewas set from 1 to 100 kPa. The rotor was driven by a 2 kWelectrical motor which is used as the starter. Theperformance of the compressor decreased under the lowambient pressure. When the ambient pressure is 1 kPa,

Fig. 6. Schematic drawing of BOV-3.

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the pressure ratio and the corrected flow rate are reduced,compared with these in case of the atmospheric ambientpressure by 10% and 40%, respectively. This is because thecompressor breathes the thick boundary layer.

Some additional tests, such as the ignition test of thecombustor and the engine start-up tests of the core-engine system under the sea-level and low ambientpressures, are discussed in the following sections.

Fig. 7. Apparatus of the ignition test of the combustio

Fig. 8. Result of the comb

3.2. Ignition test of the combustor under the high-altitude

condition

The ignition and flame-holding tests of the reverse-flow annular-type combustor under high-altitudecondition were conducted at Akiruno Testing Center ofJAXA as shown in Fig. 7. GH2 is supplied to the combustorconnected to the vacuum chamber whose pressure is

n chamber under the low-pressure conditions.

ustor ignition test.

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T. Sato et al. / Acta Astronautica 66 (2010) 1169–11761174

preset. The temperature rise of the thermo-couples set inthe combustion chamber judges whether the ignition issuccessful or not. The test result is shown in Fig. 8. Thecondition of the BOV-3 is indicated as the area surroundedbetween the red and blue lines which show the pressures

0

1 104

2 104

3 104

4 104

5 104

6 104

-20 0 20

Test_ID: BG

rotational speed

turbine inlet temp.

Rot

atio

nal s

peed

[rpm

]

(5%

)st

arte

r pow

er-u

p(2

0 %

)

(GH

2 su

pply

)

Fig. 9. Core-engine start-up test und

Fig. 10. Apparatus of the core-eng

in the combustor in 100% and 60% of the compressorrotating speed cases, respectively. Marks O and X meanssuccess and failure of the ignition. The repeatability issatisfied in the test. The test result shows the ignition isnot successful in the original BOV-3 flight conditions.

0

500

1000

1500

40 60 80

-75_6

Tem

pera

ture

[K]

Time [sec]

er the atmospheric pressure.

ine test under low pressure.

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T. Sato et al. / Acta Astronautica 66 (2010) 1169–1176 1175

Some measures such as change of the igniter position,addition of the pilot burner and/or change of the flighttrajectory are discussed.

Fig. 11. Core-engine start test under the high-altitude condition.

0

1 104

2 104

3 104

4 104

5 104

6 104

-20 0 20

Test_ID:L

Rot

atio

nal s

peed

[rpm

]

win

dmill

con

ditio

n

(GH

2 su

pply

)

Fig. 12. Rotational speed and turbine inlet temperature on

3.3. Core-engine start-up test at ground and high-altitude

conditions

The system firing tests were successfully performed asmentioned above. However, the time the compressorreaches to the idling rotational speed (about 40,000 rpm)is too long to satisfy the BOV-3 flight sequence. To reducethis time, the core-engine start-up test using GH2 wasconducted under the atmospheric pressure at AP7 testfacility of IAT/JAXA in April 2008. One of the test results isshown in Fig. 9. The time up to the idling rotational speedis 28 s which is about a half of that in case of the systemfiring test as shown in Fig. 4. The automatic start-upcontrol was also established by this test.

The core-engine test under the high-altitude conditionis currently underway. The test apparatus is shown in Figs.10 and 11. The same test facility as in the ignition test ofthe combustor is used. The core-engine is connected to thevacuum chamber. Airflow is breathed from the atmosphereand the control valve regulates the engine inlet pressure.Typical test results are shown in Figs. 12 and 13. This testcondition corresponds to 1.95 in flight Mach number and19.2 km in altitude on the BOV-3 trajectory. Fig. 12 showsthe time-variation of the compressor rotational speed and

0

500

1000

1500

40 60 80

G-9

rotational speed

turbine inlet temp.

Tem

pera

ture

[K]

Time [sec]

the start-up test under the high-altitude condition.

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-90

-80

-70

-60

-50

-40

-20 0 20 40 60 80

Test_ID:LG-9

Comp. inComp. outTurbine inTurbine out

Pre

ssur

e [k

Pa-

gaug

e]

Time [sec]

Fig. 13. Pressure distribution on the start-up test under the high-

altitude condition.

Fig. 14. Hypersonic turbojet experimental vehicle (HYTEX).

T. Sato et al. / Acta Astronautica 66 (2010) 1169–11761176

turbine inlet temperature. Before the control valve opens,the compressor already rotates at 6000 rpm by the smallleakage flow through the control valve. When the controlvalve opens at �15 s, the airflow breathed from theatmosphere drives the compressor up to 24,000 rpm,which is called ‘‘windmill start’’. This rotational speedalmost equals that produced by the electric starter in thesystem firing tests. Hence, the electric starter may beunnecessary on the BOV-3 flight test. GH2 is supplied tothe combustor at 0 s and the PID controller is used to keepthe turbine inlet temperature at 1100 K in this test. Theengine was ignited smoothly and accelerated very fast. Thesmall perturbation of the turbine inlet temperatureremained due to the change of the airflow rate.

Fig. 13 shows the time-variation of the pressures at thecompressor inlet (engine inlet), outlet, turbine inlet andoutlet (engine outlet). The turbine outlet pressure nearlyequals the pressure of the vacuum chamber and it risesgradually due to increase of the pressure of the vacuumchamber. Before the engine is ignited (�15–0 s), thepressure of the compressor outlet is lesser than that of thecompressor inlet. That is, both the compressor and turbineare driven by the pressure difference between the upstreamand downstream of the core-engine, which is called ‘‘wind-mill condition’’. When the engine is ignited, the compressorinlet pressure decreases and outlet pressure increases,which means the engine works normally. More number oftests will be conducted along the flight trajectory.

4. Conclusion

JAXA performs the development study of the precooledturbojet engine for the flight demonstration after theground tests. Presently, the subscale engine including thevariable geometry inlet and nozzle has been developed.The engine will be demonstrated in the ground and flighttests in few years.

After that, the demonstration under the higher Machnumber condition up to Mach 5 will be conducted byapplying composite materials to the engine. The hyper-sonic turbojet experimental vehicle (HYTEX) installing theS-engine has also been designed and researched as shownin Fig. 14 [5]. The configuration is designed by anintegrated optimization analysis on the aerodynamicsand trajectory. The flat-shaped vehicle whose length is4.5 m installs 1 or 2 sets of the S-engine. Several secondsof the Mach 5 cruise by using auxiliary accelerationdevices such as the high-altitude balloon and/or solidrocket. Aerodynamic performance of the vehicle has beenanalyzed by CFD and tested in the subsonic, supersonicand hypersonic wind tunnels.

References

[1] T. Sato, H. Taguchi, H. Kobayashi, T. Kojima, Development study ofMach 6 turbojet engine with Air-precooling, Journal of the BritishInterplanetary Society 58 (7/8) (2005) 231–240.

[2] N. Tanatsugu, T. Sato, Y. Naruo, et al., Development study on ATREXengine, Acta Astronautica 40 (1996) 165–170.

[3] T. Sato, H. Taguchi, H. Kobayashi, T. Kojima, et al., Developmentstudy of precooled-cycle hypersonic turbojet engine for flightdemonstration, Acta Astronautica 61 (2007) 367–375.

[4] K. Fujita, S. Sawai, R. Kobayashi, et al., Precooled turbojet engineflight experiment using balloon-based operation vehicle, ActaAstronautica 59 (1–5) (2006) 263–270.

[5] T. Tsuchiya, Y. Takenaka, H. Taguchi, Multidisciplinary designoptimization for hypersonic experimental vehicle, AIAA Journal 45(12) (2007) 3021–3023.