IAF-97-S.5.01
DEVELOPMENT STUDY ON ATREX ENGINE SYSTEM
Tetsuya Sato*, Nobuhiro Tanatsugu†, Yoshihiro Naruo±, Hiroshi Hatta†
Institute of Space and Astronautical Science
3-1-1, Yoshinodai, Sagamihara, Kanagawa 229, JAPAN
Junsuke Omi§
Ishikawajima Harima Heavy Industries CO., LTD.
3-5, Mukodai-cho, Tanashi-shi, Tokyo 188, JAPAN
Jun'ichiro Tomike§
Kawasaki Heavy Industries, LTD.
1-1, Kawashki-cho, Akashi, Hyogo, 673 , JAPAN
Ruichi Minami§
Mitsubishi Heavy Industries, LTD.
10, Oye-cho, Minato-ku Nagoya 445, JAPAN
ABSTRACT
Development study on an expander cycle air turbo ramjet engine (ATREX) has been engaged in the Institute of Space
and Astronautical Science since 1988. The ATREX is a combined cycle engine that works as a turbojet in the lower speed
flight and as a ramjet in the higher speed flight up to Mach 6
and thus is a candidate for the propulsion system of the fly
back booster of the TSTO spaceplane. The present paper addresses the following topics. First, firing tests of the precooled
expander cycle ATREX carried out under the sea level static
condition installing a newly designed precooler and a
regeneratively cooled combustion chamber made by DASA.
The improvement of overall engine performance was verified
as well as the individual performance characteristics. Second,
wind tunnel tests on an axisymmetric variable geometry air
intake and a plug nozzle. In this wind tunnel tests were examined the effects of design Mach number, air bleeding mode,
angle of attack and roundness of the center spike tip upon the
total pressure recovery and the mass capture ratio. The boat
tail drag of the plug nozzle with an axisymmetric spike was
measured and its reduction by air injection into the boat tail
was tried. Third, the effect of precompression caused by the
vehicle forebody on the engine performance. Fourth, the application studies on advanced carbon-carbon composite for the
turbo machinery of ATREX. And fifth, flight test planning of
ATREX engine. Here are discussed the engine performance
along the flight trajectory and a conceptual design of the flying
test bed.
study on an expander cycle air-turbo ramjet engine designated
by "ATREX" has been conducted as a candidate for the propulsion system of the fly back booster of the TSTO spaceplane
in the Institute of Space and Astronautical Science in cooperation with the Ishikawajima Harima Heavy Industries, the
Kawasaki Heavy Industries and the Mitsubishi Heavy Industries (1),(2) . ATREX works from liftoff to Mach 6 at an altitude
of 30 km with the advantage thrust and specific impulse due to
installation of an air-cooling system and the tip turbine made
of carbon/carbon composite as well as the variable geometry
air intake and plug nozzle.
ATREX development program initiated in 1988 and
now is in the phase shown in Fig.1. In the first approach, the
sub-scaled engine model (ATREX-500) with fan inlet diameter of 300mm was tested under sea-level static conditions. In
this phase, the turbomachinery composed of the tip-turbine
and fan with ceramic bearings, the internal heat exchanger and
the combustor was evaluated and improved with some modifications. In 1995, ATREX-500 engine was tested with the first
precooler model and thereby the fundamental performance
improvement due to air precooling was verified. However, all
of the tests planned could not be completed due to a little leakFiring Test under
Sea Level Static
Condition
¥Supersonic Wind Tunnel Test
¥Subscale Engine Model
Fan diameter of 300mm
Air precooling system
Tip turbine cofiguration
Regeneratively cooled
combustor
INTRODUCTION
Recently, a fully reusable system is controversial for
the next generation space launch system. Essential requirements for this system are cost reduction in addition to improvement on safety and reliability like an airplane and also
low environmental impact. The reusable spaceplane propelled
by airbreathing propulsion systems seems to be a most preferable candidate to realize such requirements. The development
Copyright © 1997 by the International Astronautical
Federation. All rights reserved.
* Research Associate, Space Propulsion System, † Professor,
Space Propulsion System, ± Research Associate, Space
Applications, § Staff, Aerospace Engineering
Experimental and Numerical
Component Tests
-Axisymmetric air intake
-Plug nozzle
-Interference between forebody
and engine
¥Subscale Test
-Precooler
-Mixer and combustor
¥Numerical Simulation
-Forebody precompression
-Air intake
Examination for
Flight Test
¥Development Study on ACC Turbine
¥Examination of Engine System
¥Conceptual Study of Flight Test
Unmanned Flight
Test with FTB
Fig.1 DEVELOPMENT PROGRAM ON ATREX
1
A flow diagram of ATREX engine system is shown in
Fig.2. ATREX engine is a precooled EXpander cycle Air
Turbo Ramjet engine using liquid hydrogen as a fuel and a
coolant. The air intake under development has an
axisymmetric variable geometry with mixed compression
mode. Because maximum Mach number at the intake entrance
is reduced to approximately 5.3 from 6.0 by utilizing the
forebody precompression, total pressure recovery is supposed
to increase. The air flow passed from the air intake is cooled
down to 160 K at SLS condition by an air precooling system
called "precooler". The precooler increase thrust and specific
impulse and enlarge the region of flight Mach number. The air
flow is compressed by a three-staged fan whose pressure ratio
is about 3.7 at SLS condition, while ram compression is dominant at higher Mach number. Tip turbine configuration assembled on the fan peripheral tip is employed to give compactness and light weight to the turbo machinery. The liquid hydrogen pressurized to about 5MPa by a turbo pump is heated
regeneratively in the precooler, the regeneratively cooled
combustion chamber wall and the heat exchanger. This hot
Precooler
Tip Turbine
ATREX-500 FIRING TEST AT SEA LEVEL STATIC
CONDITION
ATREX-500 Engine Model and Measurements
A schematic figure and a photograph of overall
ATREX-500 engine set on the test stand for thrust measurement are shown in Fig.3 and Fig.4. Full length of the engine
was approximately 5m and the fan inlet diameter was 0.3m. A
bell mouth and an exhaust convergent nozzle were equipped
instead of the axisymmetric air intake and the plug nozzle.
Main components such as the tip turbine are made of metallic
materials not of carbon / carbon composite. Hydrogen flow
supplied from a pressurized LH2 tank is heated in the precooler
tubes. After that it enters the combustion chamber wall in parallel with the internal heat exchanger and drives the threestage turbine. A schematic drawing of the new type precooler
model (Type-II) is shown in Fig.5 and these structural parameters are listed in Table 1 in contrast with Type-I model. Both
models are shell-and-tube heat exchangers made of stainless
steel. Type-II model was designed with some modifications
for easy fabrication and improvement on reliability because
4960
Type II Precooler
2619
Video Camera
Air
Bell
Mouth
760
DESCRIPTION OF ATREX ENGINE SYSTEM
hydrogen drive the tip turbine and is mixed with the air flow
by a lobe type mixer. The mixer which also plays the role of
the frame holder consists of 16 skewed lobes by 15 degrees to
promote mixing and enhance combustion. A gaseous hydrogen/oxygen torch fires during the startup transient and then
combustion is kept without the torch. A variable shaped plug
nozzle is employed to get the effective nozzle expansion over
the wide flight environment. By employing carbon / carbon
composite material to the tip turbine, the regeneratively cooled
combustion chamber, the plug nozzle, etc., simplicity of cooling mechanism and great increase of specific impulse are expected.
1500
age from its thin tubes and brazing parts. In 1996, the new
precooler was designed taking the structural reliability seriously rather than the performance and tested integrating
ATREX-500. The regeneratively cooled combustion chamber
consisting of winding tubes made by Daimler Benz Aerospace
(DASA) was also tested integrating in ATREX engine. In parallel with ATREX-500 development, the key components for
a flight oriented ATREX (e.g. air intake, precooler and plug
nozzle) have been studying experimentally as well as numerical study. More than ten types of axisymmetric variable geometry air intake have been tested in the supersonic wind tunnel
up to Mach 4. Verification tests of air intake beyond Mach 4 is
planned by international cooperation. In a near future are
scheduled some tests concerning with dynamic control of the
air intake working continuously in the time dependently variable airflow condition. In order to make clear the icing or
frosting problem on the precooler surfaces, the fundamental
tests will be carried out with a subscale precooler model. A
variable geometry plug nozzle was conducted in the wind tunnel focusing on the reduction of boat tail drag. The aerodynamic interference between a forebody and an engine have
been examined experimentally in the wind tunnel as well as
the numerical analyses. The precompression effect due to the
forebody on the engine performance was examined by CFD.
Finally, the application studies on advanced carbon-carbon
composite (ACC) for the ATREX tip turbine have been conducting. The manufacture techniques, structural strength and
rotational vibration have been examined experimentally and
analytically.
The flight test of ATREX is also now planning by using
an unmanned flying test bed. The status of the individual studies are summarized in the present paper.
2341
DASA
Combustor
Type II
Heat Exchanger
unit:mm
Fig. 3 Configuration of ATREX-500 Engine
Heat Exchanger
Air Intake
Fan
LH 2
Combustion Chamber
Plug Nozzle
Pump
Fig. 2 ATREX Engine Flow Diagram
Fig. 4 ATREX-500 Engine Set on Test Stand
2
the hydrogen minor leakage had been detected in Type-I after
several tests. The modifications were done in the tube diameter, the wall thickness, the number of the tubes and the number of the coolant paths. These tubes are divided 6 blocks with
3 rows each in the radial direction and they are connected in
Supporting Plate
Coolant
A
Air
Type-I Precooler
φ 300
φ 388
φ 624
φ 645
Type-II Precooler
Air
D
Air
B
C
(
)*
(
)*
Coolant
* ( Parallel Flow Configuration)
Counter Flow Configuration
Fig. 5 Configuration of Baraban Type Precooler
Table 1 Specification of Precooler
Tube Outer Diameter
Wall Thickness
Length
Rows in Circumferential Direction
Rows in Radial Direction
Total Number of Tubes
Heat Transfer Area
Compactness
Number of Coolant Path
Number of Supporting Plates
Angle of Airflow to Tubes
Number of Brazing Points
mm
mm
mm
m2
m2/m3
deg
Hydrogen Inlet
Spiral Tubes
Type-I
3
0.15
820
280
24
6720
51.9
338
8
2
70
13440
Type-II
5
0.3
'
176
18
3168
40.8
265
6
4
90
3168
600
HP/HVOF Splay Jacket
400
φ 760 (Type-II), φ 712 (Type-I)
Air
Combustion
Gas
Manifold
Hydrogen Outlet
Fig. 6 Regerative Cooled Combustion Chamber
Table 2 Specification of Combustion Chamber
Chamber Outer Diameter
400mm
Length
600mm
Weight
54.5kg
Heat Exchange Area 0.754m2
Tubes
Material
Inco-600
Number
125
Section
4x4mm Sq.
Thickness
0.32mm
Max. Pressure
4.0MPa
Max. Temperature
900K
series to make hydrogen flow path. Coolant flows from inner
to outer blocks in turn changing its flow direction. Air flows
through the tube banks from outer to inner in the radial direction along four plates placed to support the tube assembly. The
tubes are bent at the front end of the tube assembly to reduce
brazing number and to eliminate heavy bulk head. Lip plates
were equipped at outer tube assembly to make uniform air admission. The regeneratively cooled combustion chamber was
designed in cooperation with Daimler Benz Aerospace
(DASA) shown in Fig.6 and Table 2. It is formed by spirally
wound 4 mm x 4 mm rectangular tubes covered on the outer
side by sprayed alloy SUS316L(3). This manufacture technique
will be applicable to complicated shapes such as the plug
nozzle. The internal heat exchanger consists of shell and tubes
integrated by brazing technique. Approximately 100 measurement points were equipped such as a thrust, a fan rotational
speed, pressures, temperatures, flow rates and shaft vibration.
Thermo-couples were set on the outer wall of the combustion
chamber to watch its over heating. The combustion flame and
the ice formation on the precooler tube surfaces were observed
by a ultraviolet photograph and a miniature video camera.
Purposes and Test Styles
Firing tests conducted in 1996 is the 9th test, so we
called it "ATREX-9". Purposes of ATREX-9 are shown as follows.
(1) Checking structural problems of a new type precooler and
a regeneratively cooled combustion chamber with high pressure and high temperature.
(2) Researching effects of ice formation on the precooler tube
surface and the temperature distortion of the fan inlet upon
performances of the precooler itself and the total engine system.
(3) Verifying effects of the air precooling on the engine thrust
and specific impulse.
(4) Verifying the heat exchange performance by the new combustion chamber.
(5) Verifying the increase of specific impulse with increasing
the turbine inlet temperature by the combustion chamber.
(6) Acquiring the performance data of each component to
check the analysis code of the engine system.
(7) Confirming the operation manner on ATREX engine.
The firing tests were conducted step by step from
checking the structural problem to accomplishment of the best
performance as follows. Test style and typical results are
shown in Table 3.
(1) Non-expander cycle test without air cooling (ATREX9-1)
Liquid hydrogen is directly supplied to tubes of the
combustion wall and ejected outside. The turbine is driven by
the other pressurized gaseous hydrogen of room temperature.
The precooler is installed upstream of the fan without the coolant. The heat exchanger is not employed.
(2) Non-expander cycle test with air cooling (ATREX9-2,9-3)
The turbine is driven by same manner as mentioned
above. Liquid hydrogen is supplied to the precooler and then
the combustion chamber wall as the gaseous coolant and then
discharged outside. The hydrogen flow pattern in the
precooler is changed, that is parallel flow and counter flow
c o n f i g u r a t i o n in ATREX9-2 and in ATREX9-3
respectively.The heat exchanger is not employed.
(3) Expander cycle test (ATREX9-4 to 9-7)
ATREX engine with precooling is operated with the
expander cycle at final test phase. Tests were conducted with
protecting over heating on tubes of the combustion chamber
and the heat exchanger and protecting over spin of the turbine.
3
Table 3 Summary of Test Results
Test Number
ATREX9-1 ATREX9-2 ATREX9-3 ATREX9-4 ATREX9-5 ATREX9-6 ATREX9-7
Duration (sec)
55
65
65
70
50
50
65
Humidity[%]
43
72
38
64
95
57
99
Flow Style of Precooler
Uncooled
Parallel
Counter flow
Test Style
non-Expander
Expander (installed Type-II Heat Exchanger)
Fan Inlet Temperature [K]
294
220
210
203
206
217
226
Turbine Inlet Temperature [K]
292
278
285
575
565
677
669
Thrust [N]
3930
3610
4340
3760
3813
2720
3160
Specific Impluse [N¥sec/kg]
9020
8430
9210
10680
10980
11610
13920
Air Flow Rate [kg/s]
7.37
7.87
8.05
8.57
8.57
7.19
7.24
Hydrogen Flow Rate[kg/s]
0.150
0.358
0.333
0.350
0.340
0.234
0.228
Precooler
0
100
100
100
100
100
100
Each Component Combustor
100
100
100
28
25
18
16
Hydrogen Flow Heat Exchanger
72
75
82
84
/ Total [%]
Turbine
86
91
100
100
Bleed
14
9
0
0
Test Results
Engine Cut off
Ignition
Hydrogen Supply
20000
15000
0.4
Hydrogen Flow Rate (kg/s)
Fan Rotational Speed (rpm), Thrust (N)
The summary of the test results in 1996 shown in Table
3. Tests were conducted in various humidity conditions. Serious ice formation on the precooler tubes did not generate even
if it rained. Test durations were 50-70 seconds depended on
the hydrogen storage tank inventory. Figure 7 shows the typical operation profile. It is important on the starting process that
the upstream of the main valve must be chilled down sufficiently before the hydrogen is supplied to the engine. Ignition
timing is also important to prevent misfire, backfire, blowoff
and strong ignition shock. Fan rotational speed at beginning of
combustion is reduced to about 3000rpm (3sec) by the skewed
type mixer because of improvement of the mixing as compared with about 5000rpm by the straight type mixer. Engine
is controlled automatically by increasing the hydrogen flow
rate on the first 10 seconds and after that controlled manually
up to designed fan rotational speed. Fan operation characteristics agreed well with the planned values in the air precooling
test as well as non-precooling test though the fan inlet temperature has a little incline in radial direction. Although it is
difficult to get the steady state completely because of the metal
heat capacity and the precooler icing, we got the quasi-steady
Rotational Speed
10000
Hydrogen Flow Rate
Thrust
5000
0
-20
0
20
40
60
Time (sec)
80
0.3
0.2
0.1
state after 55 sec. Long duration tests (about 3 minutes) will be
conducted using a new LH 2 storage in 1998 to confirm the icing problem and engine durability in the same period as the
flight test.
Total pressure drop of the air flow side in the precooler
is plotted with predicted values in Fig.8. Test results are much
larger than the predicated values calculated on the assumption
that air flows uniformly within tube banks. In order to make
clear the cause of the larger pressure drop, the air flow in the
precooler was analyzed by means of CFD. The air flow was
assumed to be axisymmetric incompressible and viscous fluid
with constant temperature. The effect of tubes was taken into
account in CFD calculation as an external force term in momentum equations and a production term of k and ε in the
turbulence model, which was given by the empirical equation
to calculate the pressure drop of uniform flow through tube
banks (4). CFD analytical results are shown in Fig.9. Large
pressure drop is caused in the inlet part of the tube banks by the
steep expansion of the flow area. Moreover, incoming air has
tendency to go straight due to its large inertia, so the flow field
in the tube banks is biased strongly to cause additional pressure drop. However, pressure drop in Type-II model could be
about 70% compared with Type-I model. CFD results agree
well with the test results in Fig.8. Figure 10 shows the comparison of the predicted heat transfer rate with the test results.
Marks ⁄ ,¥ indicate the test results measured in the hydrogen
side; Marks ¢ ,£ in the air side. These two values should be
the same in each test, however the hydrogen-side heat transfer
10
Total Pressure Drop, kPa
At first, too much rate of hydrogen was supplied to the combustion chamber and the surplus to drive the turbine was bypassed outside. Hydrogen flow rate of each component
changed test by test with referring the heat exchange performances of the combustion chamber obtained in the previous
test (see Table 3). In the final test , high specific impulse was
obtained without bypass flow.
8
6
4
2
0
0
100
Fig. 7 Tyical Operation Profile on ATREX-9-7 Test
Test Results
Calculation / CFD
Calculation / Uncooled
Calculation / Cooled to 180K
0
2
4
6
8
Air Flow Rate, kg/s
10
Fig. 8 Air Flow Pressure Drop by the Precooer
4
0
0.2
0.1
0
-1000
-1000
-2000
-4000
-5000
-3000
-2000
-4000
-7000
-2000
-6000
-5000
0.2
0.4
-4000
-3000
Total Pressure Drop, Pa
0.6
0.8
Axial Coordinate, m
1
1.2
Velocity Vector
1.4
0.3
time=43s
0.2
0.3
0.4
0.5
0.6
0.7
0.8
Axial Coordinate, m
0.9
1
1.1
1.2
Fig. 9 CFD Result on Pressure Drop by the Precooer
1200
1000
time=65s
800
Fig. 11 Icing Observation by Video Camera
0
0
200 400 600 800 1000 1200
Heat Transfer Rate [Prediction], kW
Fig. 10 Heat Transfer Rate by the Precooler
rate is 5-20% larger than the air-side heat transfer rate. This
discrepancy may be due to the unsteadiness caused by the heat
capacity of the tubes. In some tests conducted in the rain, the
latent heat of the raindrops contained in the incoming air also
brings about the disagreement. All the test results are 10-20%
smaller than the prediction caused by some effects such as nonuniformity of the air flow through the tube banks and thermal
resistance of the frost layer formed on the tube surface. Figure
11 shows photographs of the tube surface recorded by the
video camera set at A-D point in Fig.5. The exfoliation of the
frost layer is observed at point C (upper in Fig.11). It is guessed
due to the large local velocity of air flow. It is observed at point
D that the exfoliated frost clusters are caught between tubes
(lower in Fig.11). However no serious problem, such as terrible
increase of pressure drop, occurred in the test duration.
Performance characteristics of the internal heat exchanger are shown in Fig.12. Heat flux is indicated as a function of the combustion gas flow rate and the temperature difference between the averaged hydrogen temperature and the adiabatic flame temperature of the combustion gas. Marks indicate experimental results with the regeneratively cooled combustion chamber and marks ~ indicate results without it. Both
data are mostly on a linear line except that the data with the
combustion chamber are smaller than the data without it because of cooling the combustion gas beforehand by the combustion chamber. Heat flux performance of the regeneratively
cooled combustion chamber in Fig.13. Both axes indicate as
the same manner of Fig.12. Comparing with the heat exchanger, heat flux of the combustion chamber is smaller be-
1000
Type-II Heat Exchanger
800
]
200
Type-I, Hydrogen side
Type-I, Air side
Type-II,Hydrogen side
Type-II,Air side
2
400
cause the unburnt hydrogen may gather around the chamber
wall and the gas temperature becomes lower. It does not stand
on a linear line because it depends pretty on the hydrogen flow
rate.
Engine performances depend on the fan inlet and turbine inlet temperature (FIT and TIT ). This dependency on
thrust and specific impulse is shown in Fig. 14. Contours in
these figures are analytical results at constant fan rotational
speed of 17,500 rpm and symbols indicate the history of the
600
400
200
0
0
2000
4000
6000
W0.85 x
8000
1 10 4
¢T
Fig. 12 Heat Flux Characteristics on Heat Exchanger
1000
Regenerative Combustion Chamber
800
]
600
Heat Flux [kW/M
Radial Coordinate, m
0
-3000
2
Radial Coordinate, m
0
-1000
-2000
-3000
-4000
-5000
-6000
-7000
-8000
-9000
Total Pressure
Drop, Pa
Wair=7kg/s
-2000
-1000
0
0.3
Heat Flux [kW/M
Heat Transfer Rate [Test Result], kW
0.4
600
400
200
0
0
2000
4000
W0.85 x
6000
8000
1 10 4
¢T
Fig. 13 Heat Flux Characteristics on Combustion Chamber
5
combustion tests in which changed the type of the heat exchanger, the precooler and the combustion chamber. Thrust
increase with decrease of FIT is caused by the following two
reasons. First, air flow rate increases due to increase of its density. Second, fan pressure ratio becomes larger than no precooling case at the same mechanical rotational speed. For example, thrust can be increased approximately 2.6 times by reducing FIT to 160K from room temperature at 288K of TIT.
On the other hand, thrust is independent of TIT because air
flow rate is nearly constant at the same FIT. Specific impulse
depends on both of FIT and TIT. Precooling reduces fan driving power caused by the intermediate cooling effect on the
compression process. In 1992, the turbine inlet temperature
increased by two heat exchangers connected in series. In 1995,
the maximum thrust of 4,960N was obtained by reducing FIT
to 175K by the precooler. In 1996, the maximum specific impulse of 14,400 N-sec/kg was obtained by increase of the turbine inlet temperature to 670K by the regeneratively cooled
290
2500
Fan Inlet Temperature [K]
270
3000
250
230
3500
210
4000
190
4500
170
5000
5500
150
150 250 350 450 550 650 750 850
Turbine Inlet Temperature [K]
Thrust (N)
0
0
1400
M1 = M 0
Pt 1 = Pt0
ρu1 = ρu0
00
Ft
Fd
Fuselage
20
00
0
170
*
Captured Flow
180
190
ξ = ξ ( M1 )
a) No Precompression
160
0
120
100
0
0
00
8000
6000
Fan Inlet Temperature [K]
210
In higher Mach number flight condition, larger total pressure losses occur in the compression processes of intake and
r e d u c e the p r o p u l s i v e p e r f o r m a n c e . ' F o r e b o d y
precompression' is one of the solutions to reduce total pressure
losses. Figure 15 shows the concept of the precompression.
The engine located underneath fuselage can swallow the incoming air compressed in front of air intake. By utilizing the
forebody precompression, total pressure recovery and mass
capture ratio increase. These effects lead to augmentation of
engine thrust, however drag of fuselage also increases. Therefore, precompression effect can be either advantageous or disadvantageous to the propulsive performance.
The precompression effect for spaceplane was analyzed by
CFD calculation using 3-dimensional compressible NavierStokes equations. The equations are solved with implicit TVD
scheme. The turbulence eddy viscosity is calculated by
Baldwin and Lomax algebraic model.
Figure 16 shows the model of this analysis supposing
FTB(flying test bed) of ATREX engine, where both fuselage
and engine are axisymmetric. The length of the fuselage and
nose are 7D(D:diameter of fuselage) and 3D respectively. The
diameter of the engine is 0.7D.
Total pressure recovery and mass capture ratio are the
functions of Mach number at the entrance of intake. With
precompression, these characteristics are expressed as eq's.(1)
and (2). M, Pt and m denote Mach number, total pressure and
mass flow rate respectively. Subscript '1' and '0' denote the
entrance of air intake and the free stream conditions respectively.
P
Total Pressure Recovery η* = η( M1 ) t1
(1)
Pt 0
Mass Capture Ratio
270
230
Numerical Analysis on Forebody Precompression
m1
(2)
m0
Optimization of the shape of vehicle from a view point
of precompression was done here. It is found that the shape of
axisymmetric vehicles should be designed so as to reduce the
290
250
combustion chamber. This ATREX-500 engine can be thrust
of 5,600N and specific impulse of 17,000 N-sec/kg by improvement of the precooler.
150
M0, Pt 0, ρu0
150 250 350 450 550 650 750 850
M1, Pt 1, ru1
Engine
Turbine Inlet Temperature [K]
Specific Impulse(N• sec/kg)
Mark
Heat Exchanger
FY
1991
Type-I
1992
Type-II
1992
Type-IÅ~2
1992 Type-I+Type-II
1995
Type-II
1996
Type-II
Precooler
Type-I
Type-II
b) Precompression
M1 < M 0
Pt 1 < Pt0
ρu1 > ρu0
Captured Flow
Cooled C/C
Type-I
Ft
Fd
M0, Pt 0, ρu0
Fuselage
M1, Pt 1, ρu1
Engine
Fig. 14 Effect of Fan and Turbine Inlet Temperature on
Engine Performance Characteristics
Fig. 15 Concept of the Forebody Precompression
6
3.0 D
7.0 D
‡D
Fuselage
Nose Shape =
Straight Cone
0.2 D
Engine
7.0 D
‡0.7D
3.0D
Fig. 16 Base Model
Mass Capture Ratio (
)
Total Pressure Recovery ( ¯)
1.2
¯
¯*
1
*
0.8
0.6
0.4
mixed compression type of two stage external (including isentropic) and four stage internal compression has been selected
as a result of some wind tunnel tests. The spike is moved back
and forth by an electromotor as well as the back pressure control valve. Figure 19 shows the tip configuration of the test
model (Type-K) which cowl inlet diameter is 123mm. The
supersonic diffuser were designed by the method of characteristics under the maximum design Mach number of 5.3 in consideration of the forebody precompression effect. Boundary
layers are bled from 2280 holes opened on the spike surface
and 3000 holes on cowl surface which diameter of 1mm. Because the throat height of the axisymmetric air intake is shorter
than that of two-dimensional intake, the boundary layer has
strong harmful effects. The flow rate and the location of bleed
are adjustable. As for the spike bleed, the bleed region is divided into three parts (Zone-1, 2, and 3) and the bleed rate is
individually controlled by the orifice. Cowl bleed region is
also changed by closing bleed holes. The intakes with five
kinds of round nose were tested as the measure against the
aerodynamic heating of high Mach number. The performance
characteristics of air intake against angle of attack were also
investigated. Air intake performances are evaluated by total
pressure recovery and mass capture ratio.
0.2
Test Results
0
0
1
2
3
4
Flight Mach Number
5
6
Fig. 17 Precompression Effect for Flight Mach Number
useless compression at nose. For instance, the air intake
should be located far from the fuselage surface unless it interferes with shock wave, the nose length should be longer and
the nose tip should be a little blunt.
The i n f l u e n c e of f l i g h t Mach number on
precompression effect is shown in Fig.17. The angle of attack
is 5 degree. It is found that precompression effect is remarkable at flight Mach number beyond 4. It is due to the fact that at
lower flight Mach number, the angle of oblique shock is larger
and thus the inlet position of air intake is located relatively
near the fuselage surface, which results in breathing air with
large total pressure drop. Accurate estimation of
precompression effect is indispensable especially for hypersonic vehicles.
Total pressure recovery and mass capture ratio for each
Mach number in Fig.20 and Fig.21 respectively and the specification of each intake type is shown in Table 4. These results
are not obtained in the optimum bleed pattern test and total
pressure recovery could be increased about 5‘10 % by employing optimum bleed. For example, Type-C model achieves
total pressure recovery of 77% in case of the optimum bleed at
Mach 3.5 (Mark ¡ in Fig.20) compared with 69% in improper
one (Mark ). Generally total pressure recovery and mass
capture ratio decrease with increasing maximum design Mach
number. On the other hand large total pressure drop comes
about when free stream Mach number exceeds the design
Mach number. Mass capture ratio is diminutive at transonic
condition. However, it is not a critical problem because the
isentropic
compression
normal
shock
spike
bleed
bypass
flow
spike moves
Study on Axisymmetric Variable Geometry Air Intake
oblique
shock
A development study about a variable geometry
axisymmetric air intake for ATREX have been conducted
since 1993. ATREX engine is operated from lift-off to Mach 6
under accelerating flight conditions without cruising and thus
the air intake needs to be controlled always so as to adapt to the
changing flight conditions. The throat area is controlled by
moving its centerbody (called "spike") back and forth. The
axisymmetric configuration is adopted from the view point of
simplicities on the variable mechanism. The air intake is required to accomplish lower total pressure loss in the compression process (high total pressure recovery) and larger air mass
capture to meet the demand of turbo-machinery (high mass
capture ratio). This study has been conducted by the supersonic wind tunnel tests as well as by the numerical analyses.
cowl
bleed
spike bleed
flow
Fig. 18 Schematic Draw of Axisymmetric Air Intake
p
Cowl bleed zone
Spike bleed zone
(Zone-1) (Zone-2) (Zone-3)
Air intake models for the wind tunnel tests
A sketch of the air intake is shown in Fig.18. The
Fig. 19 Detail on the Air Intake Spike
7
Table 4 Specification on Air Intake Model
Air Intake Model
Type-C
Type-E
Type-F
Compression Mode
A
a
A
A
Test PY
1994
1995
1995
1996
Design Maximum Mach Number
3.5
4.5
4.5
5.3
Half Cone Angle of Center Spike (deg.)
10
17
8
10
2.62
2.34
Total Pressure Recovery
i
¯j
Spike Length^Cowl Diameter
2.02
1.23
Compression Mode AFExternal 1+Isentropic+Internal 4
Compression Mode BFExternal 2+ Internal 3
Type-K
1
Total Pressure Recovery i
(Zone1, Zone2, Zone3, Cowl)
0.7
(*,2.7,*,Front14)
Group B
(*,2.6,*,Full)
0.6
(*,2.5,*,Front10)
(*,1.9,1.2,Full)
(*,2.1,*,Close)
(-0.2,2.0,*,Full)
Group A
0.5
(*,0.9,2.3,Full)
(-0.6,1.5,0.7,Full)
(0.0,0.4,0.7,Full)
0.4
0.52 0.54 0.56 0.58
0.6
0.62 0.64 0.66
Mass Capture Ratioi
j
Fig. 22 Effect of Bleed Pattern on Intake Performance
bleed from just upstream of the throat is effective and that
from the downstream decrease mass capture ratio on both of
spike and cowl. As for the spike nose bluntness, the largest
nose radius clearly reduces total pressure recovery because a
separation of the oblique shock wave from the spike tip
changes the shock and flow pattern considerably. However, a
proper bluntness could improve pressure recovery. Total pressure decrease with increasing the angle of attack because the
pressure side falls into the unstart condition prior to the opposite side. The intake with smaller spike angle and/or blunter
nose is sensitive to the angle of attack.
Numerical Analysis on Air intake and Future Plan
Optimum bleed
of Type-C
0.9
0.8
0.7
MIL.Spec.
Type-C
Type-E
Type-F
Type-K
0.6
0.5
0.4
0.3
¯j
flow rate required for the fan is also little at transonic. Figure
22 shows an effect of the bleed pattern on total pressure recovery and mass capture ratio. Percentage of bleed flow rate from
spike bleed zone 1, 2, and 3 and rows of holes opened on cowl
surface are shown in parentheses. For example (*, 2.7, *,
Front14) indicates bleed ratio of 2.7% from zone-2 on the
spike and first 14 rows of bleed holes from cowl tip are
opened. Group A in this figure shows an effect of bleed from
spike with the same cowl bleed area and Group B shows an
effect of the cowl bleed with same spike bleed pattern respectively. It is found that the bleed from the spike mainly influences on total pressure recovery, while that from the cowl influences on mass capture ratio. It is caused by following reason. The boundary layer on the spike is thicker than that on the
cowl, so the insufficient bleed from the spike causes the large
pressure drop by the boundary layer separation. On the other
hand, the excessive bleed from the cowl remove not only the
boundary layer but the main air flow. It is also found that the
0
1
2
3
Freestream Mach Number
4
Fig. 20 Total Pressure Recovery Performance
CFD with compressible viscous fields has been used to
make clear the internal flow fields and to compare with the
experiment. Figure 23 shows Mach number contours of typeC model at the freestream Mach number of 3.5, where both
axes are devided by the cowl entrance radius. Boundary layers
are bled from X/R=4.8-5.2 on the spike and 4.7-5.2 on the
cowl that is almost same as the wind tunnel test. A flow separation on the spike surface downstream of the normal shock
wave can be recognized. Total pressure recovery and mass
captured ratio almost agree with the experimental results.
As the future plan, studies on drag measurement and
automatic control of the intake are scheduled in supersonic
wind tunnel. Because of the large mount of spillage at transonic region, precise prediction of the spillage drag is important to estimate the overall engine performance. The unstart
1
Cowl
Type-C
Type-E
Type-F
Type-K
0.8
0.6
r/R
Mass Capture Ratio
i
j
1
Boundary layer
separation
0.5
0.4
0.2
Spike
0
0
0
1
2
3
Freestream Mach Number
4
0
1
2
3
4
5
6
7
x/R
Fig. 23 Numerical Results of Flowfield around Air Intake
Fig. 21 Mass Capture Ratio Performance
8
phenomenon is one of the most critical problems for the air
intake. It must be controlled as the increasing Mach number
and/or engine internal variation to prevent the unstart or to restart immediately. To begin with, the control of the normal
shock position against the back pressure change is planned.
Wind Tunnel Test on Boat Tail Drag of Plug Nozzle
The thrust nozzle of ATREX needs variable geometry
to adjust throat area as well as the air intake in order to adapt
the accelerating flight condition. The nozzle pressure ratio
must be changed from 3 to 400. The axisymmetric convergent
plug nozzle is employed for ATREX to achieve the requirements. In the axisymmetric convergent plug nozzle, throat
area can be controlled by moving plug back and forth against
cowl or by using variable cowl flaps. And plug nozzle is able
to achieve relatively higher performance under any ambient
pressure even if nozzle pressure ratio is off-design. But boat
tail drag at outer surface of cowl is so large, shown in Fig. 24
and it is severe problem especially with transonic flight. The
objective of this study is to reduce boat tail drag by blowing
secondary flow to outer surface of cowl.
nozzle expansion ratio.
Fideal = Wnozzle
κ −1


κ
2κ
1




RT 1−  
κ − 1 tc   ε  
(4) Thrust efficiency that display total nozzle performance is
defined as follow.
ηn =
Fnet
Fideal
In this experiment, nozzle design expansion ratio of
ε=9.86 is fixed and mass flow rate of secondary flow is 9%
compared to the exhaust flow rate.
Effects of the secondary flow on the thrust efficiency is
shown in Fig.26. In case of Mach number of 1.5, the thrust
efficiency increases more than 5% by secondary air flow.
In this study, it is shown that secondary air flow from
cowl holes is effective to reduction of boat tail drag in case of
low nozzle expansion ratio. This effect in subsonic and transonic conditions will be investigated.
Development Study of ACC Turbine Disk
Wind Tunnel Test
A plug nozzle model as shown in Fig.25 was tested in
the supersonic wind tunnel. Compressed air is supplied to the
chamber of the model instead of the combustion gas. The
nozzle pressure ratio is changed by changing the chamber
pressure and free stream conditions. This model has 120 holes
whose diameters are 1mm on the cowl surface through which
the chamber air is blown to the free stream (called 'secondary
flow'). This secondary flow is injected vertically against thrust
direction in order to prevent the momentum thrust generation.
In this study, it is expected that the boat tail drag can be decreased by static pressure rise on the boat tail surface.
Nozzle performances are evaluated as following
'Thrust Efficiency'.
(1) Nozzle pressure ratio (ε) is defined by chamber pressure
(Pc) and free stream static pressure (Pa).
ε=
Requirement for the tip-turbine structure of the
ATREX engine, shown in Fig.27, is extremely severe due to
high temperature environment up to 1500°C and high level of
centrifugal force. Because of this, metallic material, such as
super alloy, cannot be applied without a cooling system. Advanced Carbon-Carbon Composites (ACCs) are attempted to
be employed as the material for the tip turbine structure.
Free Stream
Cowl
Secondary Flow from
W˙ cowl Holes on Cowl
W˙nozzle
Pc
Pa
Plug
Nozzle Exhaust
Flow
(2) Net thrust (Fnet) measured by a load-cell is differnce between the gross thrust (F) and the boat tail drag (D).
Fnet = F − D
Fig. 25 Configration of Plug Nozzle
0.2
1
0.15
0.9
Thrust Efficiency
Nozzle Boat-tail Drag / Ideal Thrust
(3) Ideal thrust(Fideal) is defined by total temperature of chamber air(T tc ), mass flow rate of nozzle exhaust(W nozzle ) and
0.1
0.05
Cowl Closed
Cowl Open
0.8
0.7
Free Stream M1.5
0
0
2
4
ATREX Flight Mach Number
Fig. 24 Boat Tail Drag of Plug Nozzle
0.6
6
6
7
Nozzle Expansion Ratio
Fig. 26 Effect of Secondary Flow on Thrust Efficiency
9
5
As shown in Fig.27, the turbine disk has a complex shape, that
consists of 2 rings, the turbine ring and the fan disk, and the fan
blade connecting the 2 rings. This type of a shape is ordinarily
rather difficult to form by use of a continuous fiber reinforced
composite material. This is due to difficulty to arrange fiber
directions optimally in accordance with varying loading directions from place to place. Thus main problem in this development study of the ACC turbine disk is how to form optimum
reinforcement and design, as shown in Fig.28.
formed separately and then the fabricated each component is
assembled into the final structure. The strength of each component can be easily optimized in accordance with each stress
distribution. In order to assemble the turbine disk structure,
two types of joint configurations were employed. One of the
joints is a dovetail joint between the fan blade and the fan disk,
and the other is a pin joint between the fan blade and the turbine ring. Simplified component tests and analytical studies
have been carried out on both types of joints and satisfactory
results were obtained.
Configuration of Turbine Disk
The load and shape of the turbine disk are essentially
axi-symmetric. Therefore it is not difficult to estimate roughly
the stress distribution and the locations of stress concentrations in the turbine disk. It is obvious that the hoop stress in the
turbine ring is highest and the value of the stress is estimated to
be about 560 MPa at 440m/s, which is the rotational speed
when the flight speed is Mach 6.
To make complicated turbine structure into the optimum arrangement of the reinforcing fibers, a joined structure was attempted to be adopted. In this structure, the fan disk, fan
blades, and turbine ring + tip turbine blades were at first
Tip-Turbine Blade
Fan Blade
Fan Disk
High Speed Rotation Test
High speed rotation tests have been also carried out to
confirm the stress distributions and to establish the failure criterion of the tip turbine structure under rotation.
The test on a flat ACC disk with the quasi-isotropic stacking
sequence of carbon fiber resulted in the peripheral burst velocity equal to 427 m/s. Comparison between the results from the
rotation test and that of static tensile tests implied the failure of
the ACC during a rotation test obeying the average stress criterion over the radius of the disk rather than the maximum stress
condition. This average stress criterion was assumed to be effective in the case of stress concentration being around 3. This
is known to be one of merits ACCs have and we confirm that
ACCs have notch insensitivity for the specimen with mild
stress concentration sources also in static tests.
FEM analyses of turbine disk structures indicate that
stress concentrations are predicted to be rather mild, about 3.
Thus the average stress criterion might be applied to the
present structure. Combining the average stress criterion, tensile test results of constituent C/C composites, stress distributions by FEM analyses, and simplified model experiment, we
can estimate the burst rotational speeds of the structures,
474m/s. This result leads to a tentative conclusion that the
joined structure now we discussing is estimated to have sufficient strength but for actual design more strength should be
required because we must have margin strength for yielding
the safety factor.
Turbine Disk(Ring)
Fig.27 Schematic Drawing of Turbine Disk
M a te r i a l P r o p e r t i e s
Future Flight Test Plan on ATREX Engine
In the final phase of ATREX engine development
study, the overall system performance, functions and opera-
C o nc e p t ua l D e s i g n
E Thermal Properties
E Axial or Centrifugal Turbine
E Mechanical properties (R.T.---2300K)
E Monolithic or Joined
M a te r i a l S e l e c ti o n
E Comparion of High Temp. Properties
C/C, CMC, HTM
E Thermo-Mechanical Properties
F ab ri cat i o n
E Simplified Medels: Monolithic & Joined
E Strength & Vibration
Hi g h S t r e n g t h
A n t i - o x id a t i on S iC C oa t in g
E Spollation-Bonding strength
E Cracks in Coating-Sealing
O p ti m um S tr uc t ur e
E Densification of 3D C/C Composite (1) Reinforcement Design
(2) Joint Structure
E DoveTail-Optimal geometry
E Fiber Orientation
& Fiber Orientation
E Fiber volume fraction
C e n t r i f i c a t io n
Turbine Ring: 3D, r, ˘
, z E Pin-Notch sensitivity
Fan Blade: r, 3D, ˘
, z, } ˘ E Vibration-Joint Structure
E High Temp. Spin Burst Test
Fan Disk: Quasi-isotropic
@@
Italic
Underline
Completed
Underway
Planning
Fig.28 Principal problems and accomplished technology in the development study of the turbine disk structures
10
2.7 m
ISAS
12.5 m
LH2 Tank
ISAS
LH2 Tank
0.9 m
5.6 m
1.2 m
ATREX Engine
6.2 m
0.8 m
Fig. 29 Configration of ATREX Flying Test Bed
Thrust (N)
Specific Impulse (N-sec/kg)
50000
40000
30000
20000
10000
0
1
2
3
4
5
6
Mach Number
Fig. 30 Thrust and Specific Impulse of the ATREX Engine
along Typical Flight Path
Reusable
Reusable
Booster
40.6 m
Orbiter
References
17.5 m
5.4 m
1) T.Sato, N. Tanatsugu, et al., "Development Study on
ATREX Engine for Future Spaceplane", 7th International
Space Conference of Pacific-Basin Societies, 1997
2) N. Tanatsugu, T. Sato et al., "Development Study on
ATREX Engine", 47th International Astronautical Congress,
1996
3) Joachim Kretschmer, "Injection and Combustion Chamber
Technologies for Hypersonic Proplusion", 5th International
Symposium "Propulsion in Space Transportation", 1996
4) A. Zukauskas, R. Ulinskas, "Banks of Plain and Finned
Tubes," Heat Exchanger Design Handbook, 2, Fluid Mechanics and Heat Transfer, pp.2.2.4-1 - 2.2.4-17, Hemisphere Publishing Corp., Washington, D. C., 1983
ISAS
20.0 m
ISAS
1.7 m
ISAS
4.0 m
tions of ATREX will be verified under the real flight conditions by the flight test. The flight test will be performed by
using a flying test bed (FTB) which is powered by ATREX
engine itself after lift-off as shown in Fig.29. The FTB takes
off horizontally with the take-off assist system up to 0.25 - 0.5
of Mach number. The takeoff assist system is powered by a
conventional turbo jet engine or rocket engine or solid rocket
in parallel with ATREX engine. After reaching to the maximum speed planned, FTB is decelerated by throttling thrust
and glides down and finally lands on the sea by parachute for
recovery. The FTB flies by the autonomous and programmed
control in addition to the radio guidance system with GPS.
Figure 30 shows thrust and specific impulse of the ATREX
engine along the typical flight path of the FTB. Thrust of
ATREX increases almost in proportion to the flight dynamic
pressure. These engine performances are improved by using
advanced carbon/carbon composite turbine and fan. Beyond
Mach 3.5 the specific impulse decreases due to the increase of
the hydrogen consumption as a cooling for the structure materials of engine, on the other hand the thrust increases. This
flight test needs approximately 200 km in range and 200 seconds in time to reach the target flight velocity. Dynamic pressure is kept 50 kPa from 1 to 6 of Mach number.
Fig.31 and Table 5 show an example of practical application of ATREX for fully reusable TSTO putting the payload
of 600kg into the LEO. Gross lift-off mass is 61ton and total
thrust of four ATREX engines is 57tonf. In case of the expendable orbiter, the gross lift-off mass decreases to 39ton.
Key technologies on the ATREX engine are summarized as
follows.
-Improvement of total pressure recovery and mass capture in
the air intake
-Application of ACC material to the hot components such as
the turbine, the nozzle surface and the regeneratively cooled
combustion chamber wall, etc.
-Reduction of the precooler mass by using small diameter
tubes
-Reduction of the boat tail drag of plug nozzle
-Control of the intake and the nozzle adaptable to flight conditions changing from time to time
ISAS
Fig. 31 TSTO Space Transport System using ATREX
Table 5 Specification on TSTO System
Gross Liftoff Mass [ton]
Payload Mass[kg]
1st Stage
Gross Mass [ton]
Propellant Mass [ton]
Gross ATREX Thrust [tonf]
2nd Stage
Gross Mass [ton]
Propellant Mass [ton]
Gross Rocket Thrust [tonf]
11
Fully
Reusable
61
600
Expendable
2nd Stage
39
600
50
15.7
57
31
9.1
36
11
9.8
12.5
7.1
6.2
8.4