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Development Study on ATREX Engine1
PII:
Acta Astronautica Vol. 41, No. 12, pp. 851±862, 1997 # 1998 Published by Elsevier Science Ltd. All rights reserved Printed in Great Britain 0094-5765/98 $19.00 + 0.00 S0094-5765(97)00176-8
DEVELOPMENT STUDY ON ATREX ENGINE{ N. TANATSUGU, T. SATO, V. BALEPIN, Y. NARUO Institute of Space and Astronautical Science, 3-1-1, Yoshinodai, Sagamihara, Kanagawa 229, Japan
T. MIZUTANI, T. KASHIWAGI Ishikawajika Harima Heavy Industries CO., LTD., 3-5, Mukodai-cho, Tanashi-shi, Tokyo 188, Japan
K. HAMABE, J. TOMIKE Kawasaki Heavy Industries, LTD., 2-8-1, Takatsukadai, Nishi-ku, Kobe, 673-02, Japan
and R. MINAMI Mitsubishi Heavy Industries, LTD., 10, Oye-cho, Minato-ku, Nagoya 455, Japan (Received 2 March 1996) AbstractÐThis is the status report of the development study on ATREX engine (Air Turbo Ramjet) that is now under way in the Institute of Space and Astronautical Science (ISAS) cooperation with the Ishikawajima Harima Heavy Industries (IHI), the Kawasaki Heavy Industries (KHI), the Mitsubishi Heavy Industries (MHI). ATREX engine will be applied for the propulsion system of ¯y-back booster of TSTO space plane. ATREX is the combined cycle (a fan-boosted ramjet) engine providing the eective thrust from sea level static to ¯ight Mach number 6. ATREX is worked on the expander cycle with precooling the incoming air as shown in Fig. 1. ATREX employs the tip turbine con®guration which allows the compactness and the light weight of turbo machinery and the variable geometry airintake and plugnozzle which allow the wide range operation conditions. From 1990 to 1992, `` ATREX-500`` has been tested at the sea level static conditions. ATREX-500 is the 1/4-scale model of which fan inlet diameter is 300 mm and overall length 2,200 mm. From 1992 have been performed the wind tunnel tests on the primary components of ATREX, the axisymmetric variable geometry airintakes, the precoolers and the variable geometry plug nozzles. In parallel to the windtunnel tests, the ram combusters have been tested simulating the hypersonic ¯ight conditions and the application studies on advanced carbon-carbon composite for the tip-turbine and fan assembly has been proceeded. In 1994 initiated the ¯ight test plan in which ATREX will be veri®ed in the practical ¯ight conditions by using an unmanned ¯ying test bench. In 1995 will be tested ATREX-500 installing the precooler under the sea level static conditions to examine the engine performance and the icing on the precooler. The present paper addresses the high loading ram combuster experiment using the mixer with skewed lobes to generate swirl ¯ow and the analytical studies and the designs on the precooler and the precooled ATREX engine and the ¯ight test plan. # 1998 Published by Elsevier Science Ltd. All rights reserved
1. HIGH LOADING RAM COMBUSTOR
lobes make the swirl ¯ow of hydrogen and air in the mixing process after injecting into the combuster. This swirl ¯ow improves the combustion performance by prolonging the stay time for mixing and combustion.
The compact ram combuster is required strongly to reduce the size and weight of ATREX. The compactness of ram combustor is accomplished by the high speci®c combustion loading. The combustion performance depends strongly upon the mixer and thus the improvement of mixer is well-worn approach to increase the combustion loading. The mixers with the skewed lobes were designed and tested for the ATREX combuster. The skewed
1.1. Mixer Con®guration with Skewed Lobes
{Paper IAF-95.S5.01 presented at the 46th International Astronomical Congress, October 2±6, 1995, Oslo, Norway 851
Five types of sub-scaled mixer with dierent skewed lobes were tested. Three have dierent skew angles (Type I, Type I-B15, Type I-B20) and another has dierent axial length (Type VI-B15) and the other has dierent lobe shape (Type VIII-B15) as shown in Fig. 2. All mixers have the 16 lobes which divide the ¯ow of hydrogen and air by the ratio of
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Fig. 1. ATREX Engine Flow Diagram
1:4 in cross-sectional area. The mixers tested are scaled down to 1/5 from one of the ATREX-500. 1.2. Test Result of Mixers with Skewed Lobes 1.2.1. Speci®c Combustion Loading. Figure 3 shows the ¯ame temperature distribution indicated
by the nondimensional distance from the combustion chamber normalized by the combustion chamber inner diameter D. They were measured by the dichromatic pyrometer scanning along an axial direction at 0.38D apart from the center of chamber. It is seen from this friuge that the mixers
Fig. 2. Mixer Co®gurations with Skewed Lobes
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Fig. 3. Flame Temperature Distribution Fig. 4. Skewed Lobes Mixer for ATREX-500
with skewed lobes give the higher maximum combustion temperature compared to the non-skewed mixer (Type I), and the maximum temperature point is located around 1.0±1.2D from the combustion chamber inlet which is more upstream compared to 1.8D of the non-skewed mixer. From these results the signi®cant enhancement on speci®c combustion loading can be obtained by the mixers with skewed lobes (especially Type VIII-B15) by a factor of approximately 1.5 compared to the non-skewed mixer (Type I). 1.2.2. Flame Visualization. The combustion ¯ame structure has been analyzed by ultraviolet photography of OH-radical luminescence. The intensity distribution of ultraviolet photography is similar to the temperature distribution. It can be also concluded from this fact that the hydrogen-air mixing is enhanced by the swirl ¯ow made by the mixer with skewed lobes and the speci®c combustion loading increases. 1.2.3. Pressure Loss. The pressure drop through the combustion chamber was not so largely increased by the mixers with skewed lobes compared with non-skewed mixer (Type I). 1.3. Mixer with Skewed Lobes for ATREX-500 Mixer with skewed lobes for ATREX-500 shown in Fig. 4 was designed based on the test results of sub-scaled models. This mixer is integrated in ATREX-500 as well as the Baraban type precooler
and will be tested in September 1995 at Noshiro Testing Center of ISAS.
2. AXISYMMETRIC VARIABLE GEOMETRY AIR INTAKE
Axisymmetric air intakes have been studied using supersonic wind tunnel in ISAS from 1993 to establish design method and make clear changes of performance characteristics by bleeding to suppress boundary layers. Design Mach number shifted from Mach 3 in 1993, Mach 3,5 in 1994 to Mach 4.5 in 1995. [1] 2.1. Test Model Figure 5 shows the model for wind tunnel test. Center spike is movable to be adjusted for the given Mach number. Outer diameter of test model is 127 mm which is one sixth of ATREX engine for ¯ight test. Air bleeding is made from small holes on center spike and slits on cowl. Amount of bleeding was ®xed during blow in this test though it has to be adjusted for Mach number of incoming air stream in real ¯ight. Performances decrease abruptly and shift to unstart condition beyond design Mach number especially in the air intake with larger internal compression and therefore design point should be set at maximum ¯ight Mach number. Mach number of 4.5 is set for the present de-
Fig. 5. Axisymmetric Variable Geometry Air Intake Model for Wind Tunnel Tests
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N. Tanatsugu et al. Table 1. Features of Air Intake Models
Air Intake Model Compression Mode External Internal Characteristics Intake Length Boundary Layer Eect Startability Stability Drag in Cowl Lip Pressure Recovery at Mach 3.5 Mass Capture Ratio at Mach 3.5
Type E
Type F
Type G
2 2 High Startability Short Small Good Large 0.56 0.33
1 + Isentropic 3 High Internal Compression Very Long Large Ordinary Small 0.60 0.70
1 + Isentropic 3 High Internal Compression Long Large Ordinary Small 0.49 0.58
sign. Type E, F and G models were designed by means of method of characteristic. Compression is relatively larger in external section of type E model and in internal section of type F and G. Features of these models are summarized in Table 1. 2.2. Test Results Maximum pressure recovery factor is shown in Fig. 6 together with the target curve required by ATREX engine and MIL-spec. curve. All models were tested below Mach 4 which is maximum available blow in ISAS' wind tunnel. Type E model meets the target value below Mach 4. Type F values are smaller in lower and higher Mach number and therefore type F is estimated to be not good beyond Mach 4. Although type G values are relatively lower from 1.5 to 3.5 of Mach number, at Mach 4 it is larger than at Mach 3.5 and estimated to be better beyond Mach 4. Mass capture ratio is 70, 33 and 58 percents at Mach 3.5 for type F, E and G respectively. Eect of bleeding in type F model is shown in Fig. 7. Operation of this test was done as follows: (1) center spike moves backward gradually from foremost forward position and stops at appropriate position, and then (2) increases back pressure gradually by closing plug valve located at discharge port of air intake, and suddenly (3) falls into unstart condition. In type F model, mass capture ratio decreases keeping maximum pressure recovery factor as bleeding in cowl increases. Startablity becomes worse without cowl bleeding and leads to reduced performances. From these results, there is
Fig. 6. Pressure Recovery Factor of Air Intakes
the optimum amount of bleeding in cowl. In case of decreasing air bleeding in center spike, performances become worse due to not decreasing throat area. Boundary layer eects are larger due to larger surface area of center spike. Performances of type G model are best of all models when cowl bleeding is not done. As center spike bleeding increases, performances of type G model become better as well as type F model. 2.3. Conclusion Pressure recovery factor of each type of air intake models are being improved to meet the target of ATREX engine. Bleeding in center spike should be larger in order to suppress boundary layer. Bleeding in cowl should be made with precise amount and position. 3. AIR PRECOOLER
Generally speaking, precooling gives the following bene®ts compared to non-precooled engines: Ð increase of thrust; Ð increase of speci®c impulse; Ð extension of ¯ight range to higher Mach number. These eects could be attained by several manners and/or their combinations: Ð air¯ow increase; Ð increase of fan pressure ratio; Ð decrease of hydrogen ¯ow rate caused by reduction of fan driving power.
Fig. 7. Eect of Air Bleeding in Type F Model
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Fig. 9. Air intake and precoolers intagration: ``blin'' type (upper), ``baraban'' type (lower)
Fig. 8(a). Thrust dependcy on fan inlet temp. and precooler pressure recovery factor (ATREX-500) (b) Speci®c impulse dependcy on fan inlet temp. and precooler pressure recovery factor (ATREX-500)
In order to augment the thrust of ATREX especially below transonic ¯ight, the precooler operation is changed to work continuously from lift-o. In the early design the precooler operates above Mach 2 of ¯ight speed to prevent ice formation on it. On the other hand some problems come out by employing the precooler. Air pressure drop across the precooler has negative eect on engine performance. In order to reduce this pressure loss, the precooler structure becomes larger and heavier. Figures 8a and 8b shows the impact of precooler outlet air temperature on the SL-static performance of ATREX-500 for the precoolers with dierent air pressure recovery factors. ATREX-500 engine with air cooling temperature of 160 K and pressure recovery factor of 0.9 provides twice in the thrust and 1.5 times in speci®c impulse compared to the nonprecooled one. It is due to the fact that air ¯ow rate increases from 6.7 kg/s to 11.5 kg/s and fan pressure ratio from 1.53 to 2.03 by precooling in the same turbomachinery. According to Fig. 8 thrust sensitivity is about 1.0% per 1 K of precooling temperature change and 2.5% per 1% of pressure recovery change around air cooling temperature of 160 K and pressure recovery factor of 0.9. For speci®c impulse these ®gures are 0.4% and 1.4% respectively. It is
seen that the engine performance is aected signi®cantly by cooling temperature and pressure recovery of the precooler. In the beginning of design study of the precooler some key targets were set on precooler performance characteristics and con®guration as follows [2]: Ð the main purpose is air precooling to 160 K with air pressure recovery factor 0.85 and higher at SL-static conditions; Ð precooler should be a shell/tube con®guration with the tube diameter 3 mm and all thickness 0.1± 0.15 mm; Ð precooled ATREX engine should be operated under more stoichiometric mixture ratio; Ð icing free regime of precooler operation should be provided from SL-static conditions. A number of the precooler con®gurations were examined [2], of which two con®gurations were selected for technology development. Figure 9 shows the precoolers integrated in the axisymmetric air intake of ACC ATREX engine. This air intake con®guration is adaptive for the ¯ight of Mach number up to 6. Turbo-machinery is assembled on the right of the precooler shown in Fig. 9. These two precoolers are designated to be ``blin'' and ``baraban'' in Russia, which means ``pancake'' and ``drum'' respectively. ``Blin'' type precooler provides pressure recovery factor 0.86 at design point of 160 K. In ``Baraban'' type this ®gure is 0.94 which is higher than ``Blin'' type. This higher pressure recovery of ``Baraban'' can provide the thrust increase as large as 20% and the speci®c impulse increase as high as 12% compared to ``Blin'' type. Figure 10 shows the performance characteristics of ATREX-500 with ``Baraban'' or ``Blin'' type precooler at SL-static conditions. Thrust (top) and speci®c impulse (bottom) are indicated as contour lines in the coordinates of precooler outlet air temperature and air pressure recovery factor of precooler.
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Fig. 11. Relative Thrust to Weight Ratio based on Nonprecooled ATREX-500
Fig. 10. Performance of precooled ATREX-500 engine at SL-static and characteristics of precoolers
Figure 11 shows the relative thrust to weight ratio of ATREX-500 precooled with Baraban precooler based on non-precooled ATREX-500 (fan inlet air temperature: 288 K, air pressure recovery factor: 1.0) The maximum point locates at fan inlet temperature T around 160 K and air pressure recovery factor s above 0.95. Figures 10 and 11 are the key ®gures to decide the air precooling temperature. It is seen that in ``Baraban'' type precooler the almost maximum thrust and speci®c impulse are given at the design point (T = 160 K, s = 0.94) on the characteristic curves. In this case the decision of design point is optimum. The design point decided for ``Blin'' type precooler (T = 160 K, s = 0.86) provides about maximum thrust, but its speci®c impulse could be
improved by about 3% with a little thrust change if precooling temperature is selected as T = 170± 175 K. By this modi®cation the pressure recovery factor is improved to be 0.90±0.91 and the better engine performance can be attained and its weight could be reduced by about 30%. Needless to say that the con®dential design point should be proved by the tests. It should be noted that the approach mentioned here can provide the selection of maximum engine performance within the geometric restriction required. As the vehicle accelerates, the pressure recovery factor of precooler comes close to 1.0 very quickly and therefore the degree of air precooling is important rather than the pressure recovery factor. Taking into account the total optimization of fuel consumption required in the whole mission, more complicated analysis is necessary for the selection of precooling level at SL-static conditions. It will be one of the subjects in the next step of the study following the veri®cation of design by the SL-static test of precooled ATREX-500 to be held in September and November 1995. More information on precooler con®guration and performance were given in Ref. [3]. Table 2 summarizes the key parameters of SLstatic test results of the nonprecooled ATREX-500 and the estimated parameters of precooled ATREX-500 equipped with ``Blin'' and ``Baraban'' precoolers.
Table 2. Comparison of ATREX-500 performances at SL-static conditions nonprecooled test results, expected engines values precooled with ``Blin'' and ``Baraban''
Thrust [kgf] Speci®c impulse [sec] Air ¯ow rate [kg/s] H2 ¯ow rate [kg/s] Equiv. mixture ratio Fan pressure ratio Turbine inlet temp. [K] Fan inlet temp. [K] Fan inlet pressure [atm]
ATREX-500 nonprecooled (test results)
ATREX-500 with ``Blin'' precooler (design values)
ATREX-500 with ``Baraban'' precooler (design values)
427 1406 6.74 0.3 1.52 1.53 564 291 1.033
735 1759 10.9 0.416 1.31 2.03 650 160 0.883
909 1975 11.5 0.442 1.32 2.03 650 160 0.968
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Table 3. Correlation coecient of precooler Coecient
Precooler Type ``Blin'' type ``Baraban'' type 0.98 0.82 2.21 0
X1 X2 X3 X4
0.85 0.74 2.21 ÿ0.195
3.1. Quick estimation of precooler performance For quick estimation of the precooler performances, the simple and convenient correlations can be found in the precise calculation results. The precooler temperature eciency Zair is de®ned to be the ratio of the temperature dierence between air inlet and outlet and the temperature dierence between air inlet and hydrogen inlet: Zair
out Tin a ÿ Ta in Ta ÿ Tin h
1
where T: temperature superscript; in:inlet out: outlet subscript; a:air h:hydrogen The air pressure recovery factor sair is de®ned to be the ratio of air stagnation pressures in the outlet and the inlet of precooler: in sair pout a =pa
2
where P:stagnation pressure Superscripts; in:inlet out:outlet subscript; a:air These two parameters correlate closely with the dimensionless ¯ow rates as follows: X2 1 Zair Z0 gX a =gh
Fig. 13. Air pressure Recovery Factor of Baraban Precooler
3
Fig. 12. Air Outlet Temperature of Baraban Precooler
X4 3 sair 1 ÿ 1 ÿ s0 gX acorr gh
ga G0a ; gh Gh =G0h ; gacorr
0 s1 Ga @ p0 Tin a A 0 T0 Ga pin a 4
where Z:temperature eciency s:pressure recovery factor G:¯ow rate P:stagnation pressure T:stagnation temperature superscripts; in:inlet 0:value at design point subscripts; a:air h:hydrogen gacorr:relative corrected air ¯ow rate T0=288.15 K (158C) P0=101.3 kPa X1, X2, X3, X4: exponent coecients determined by the geometric con®guration of precooler.
Fig. 14. Hydrogen Outlet Temp. of Baraban Precooler
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Fig. 15. Con®guration of ATREX-500 with Baraban Precooler set up on test stand
These coecients X1, 2, 3, 4 become almost constant for the speci®c type of precooler over the wide range of conditions from SL-static to hypersonic ¯ight around Mach number of 6.0 at altitude as high as 35 km. The estimation by eqns (3) and (4) is satis®ed with the precise calculation within 3% of deviation. The coecients X1, 2, 3, 4 for the ``Blin'' and ``Baraban'' type precooler are given in Table 3. Figures 12±14 show the contour lines of air outlet temperature, air pressure recovery factor and hydrogen outlet temperature of the Baraban precooler on the air and hydrogen ¯ow rate map which are calculated by the quick coreration equations mentioned above. This calculations are made under the following conditions, air inlet temp.: 288.2 K, air inlet press.: 0.013 MPa, hydrogen inlet temp.: 30 K, hydrogen inlet press.: 4 MPa. Figure 15 shows the con®guration of ATREX500 with Baraban precooler when it will be tested in 1995. 3.2. Precooler Icing The icing problem in the heat exchanger for the atmospheric air cooling was pointed out by the specialists from the beginning of the study. According to the study conducted in CIAM (Russia), icing free operation of the precooled engines is examined with the physical model of ice formation based on the experimental veri®cation [4]. Two ice formation mechanisms have been discovered depending on the steam partial pressure 1. Ice formation on micro level. When steam partial pressure is lower than steam pressure at the triple point (ptr=0.006228 atm), frost is formed avoiding liquid phase. This process is indicated by the arrow crossing sublimation line AO in PT diagram of water shown in Fig. 16. In this case frost layer is growing by molecular level. Formed frosting substance is not so strong and always carried over by air stream.
A special test conducted at pst
Fig. 16. P-T diagram for water stream, AO-sublimation line; OB-solidi®cation line; OC-saturation line.
Development study on ATREX engine
Fig. 17. Seasonal borders of icing. Humidity data for Nemuro area (north-east Hokkaido)
pst
5
Ta < 273K
6
eqns (5) and (6) allow us to draw the borders to prevent icing in the speed-altitude map of ¯ight path. Atmospheric moisture content is maximum at sea level and rapidly decreases as the altitude increases. Icing borders concerning to ¯ight speed-altitude can be decided by the humidity distribution along the altitude and the stagnation temperature of incoming atmospheric air. Figure 17 shows the icing borders caused by the temperature condition given by eqn (6) and the typical trajectories of ¯ying test bed powered by ACC or metal ATREX, where are used the climate data taken at Nemuro which is the nearest city to the Taiki. Icing borders caused by pressure condition given by eqn (5) are located higher as seen in Fig. 17, and therefore the precooler icing is unlikely at the altitude about 0.5 km in November and April and at
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Fig. 18. Times to reach icing border on the seasons
the altitude about 3.5±4.0 km in July and August which is the highest humid season. According to the conclusion given by eqn (6), there is no icing on the precooler in winter season when air temperature is below 273 K at any altitude. This condition takes place from December through March around the city of Nemuro. In the rest of the year there is the icing possible from sea level to the icing borders. Conversion of the trajectory data into the time sequence gives the period of reaching the icing border which is shown in Fig. 18 for each month. It is seen that June to September is a less favorable season for the precooled engine operating for 25±30 sec to reach the icing border in the ¯ight powered by ACC ATREX and for 35±40 sec by the metal ATREX. Further analytical and experimental study is necessary to substantiate the possibility of a precooled engine to be able to operate for 25±40 sec without negative eects by icing. 4. FLIGHT TEST PLAN ON ATREX ENGINE
In the ®nal phase of the ATREX engine development study, the overall system performance, functions and operations of ATREX will be veri®ed by the actual ¯ight test. The ¯ight test will be performed by using the Flying Test Bed (FTB) which is powered by ATREX engine itself after lift-o.
Fig. 19. Flight Pro®le of ATREX-500 Flight Test
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Fig. 20. Con®guration of FTB Table 4. Speci®cations of FTB for ATREX Flight Test Weight [kg] ATREX Engine Structure Electronics Fuel Supply System Control Drive System Recovery System Fuel (LH2) Total Max. Flight Mach Number Max. Altitude [km] Max. Dynamic Pressure [kPa] Wing Loading [kg/m2] Fuel Tank Pressure [MPa] Fuel Tank Volume D L [m]
370 458 80 112 90 100 190 1400 >6 >30 50 600 0.3 0.8 6.6
The entire ¯ight pro®le is shown in Fig. 19. The FTB takes o horizontally with the take-o assist system running on the rail up to 0.4±0.5 of Mach number. The takeo assist system is powered by a conventional turbo jet engine in parallel with ATREX engine. In the early phase of the ¯ight test, the maximum ¯ight speed is up to 4.5 of Mach number with the metal turbo machinery. The ®nal target speed is Mach 6 which can be attained with the carbon-carbon composite turbo machinery. After reaching the maximum speed planned, FTB is decelerated by throttling engine power and glides down and ®nally lands on the sea by parachute for recovery. The FTB ¯ies by autonomous and programmed control in addition to the radio guidance system with GPS. The data obtained during ¯ight are transmitted by means of telemeter. Con®guration of FTB and its speci®cations are shown in Fig. 20 and Table 4.
Fig. 21. Improvement of Thrust of ACC and Metal ATREX by Precooling at SL-Static Conditions
Two ATREX engines are designed for ¯ight tests; one is the metal ATREX which is available for ¯ights up to Mach 4.5 and the other is the ACC ATREX which can attain the ®nal goal up to Mach 6. They have same size as in ATREX-500 developed for SL-static tests. Their design data are shown in Table 5. Engine performances are improved in ACC ATREX primarily due to the improvement of turbo-machinary caused by its higher peripheral speed at higher temperature compared to metal ATREX. Their performances are also improved by precooling shown in Figs 21 and 22 compared to the test results of non-precooled ATREX-500 indicated in these ®gures. The eects of air precooling can be seen on the ¯ight conditions shown in Figs 23 and 24, where the thrust and speci®c impulse are indicated for two air precooling temperatures 160 K and 220 K at SL-static conditions. The bene®t of precooling temperature 160 K is 18±25% in speci®c impulse and 25±40% in thrust compared to 220 K. Figure 25 shows air inlet and outlet temperatures and wall temperature of precooler in ¯ight conditions along the typical ¯ight path of FTB. Decrease of speci®c impulse beyond Mach 3.5 is caused by increase of hydrogen ¯ow rate for cooling purposes although thrust is increased. Typical ¯ight path of FTB powered by ACC and metal ATREX, arrival time of FTB to a given ¯ight speed and dynamic pressure along its ¯ight path are
Table 5. Design Data of ATREX Engine Made of ACC and Meatl for Flight Test Component Precooler Fan
Turbine
Combuster Heat Exchanger Thrust Nozzle
Design Parameter Air Pressure Drop [%] Outlet Air Temperature at S.L. Static [K] Pressure Ratio Eciency [%] Peripheral Speed [m/s] Inlet Diameter [mm] Inlet Hydrogen Temperature [K] Pressure Ratio Eciency [%] Peripheral Speed [m/s] Pressure Drop [%] Combustion Eciency [%] Combustion Gas Pressure Drop [%] Discharge Gas Velocity Coecient
ACC Engine 16 160 3.8 80 420 300 1500 6 40 500 5 80 8 0.97
Design Value
Metal Engine 12 160 2.7 80 320 300 650 5 40 380 5 80 8 0.97
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Fig. 22. Improvement on Speci®c Impulse of ACC and Metal ATREX by Precooling at SL-Static Conditions
Fig. 23. Thrust of ACC ATREX in Flight Conditions
Fig. 24. Speci®c impulse of ACC ATREX in Flight Conditions
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Fig. 26. Typical Flight Path of FTB Powered by ACC and Metal ATREX Engine
Fig. 27. Arrival Time of FTB Powered by ACC and Metal ATREX Engine along Flight Path Indicated in Fig. 22
Fig. 28. Dynamic Pressure along Flight Path Indicated in Fig. 22
shown in Fig. 26±28. This ¯ight test needs approximately 170 to 180 km in range and 190 to 260 seconds in time to reach the target ¯ight speed for ACC and metal ATREX powered ¯ight respectively. Dynamic pressure is kept 50 kPa from 1 to 5 of Mach number and reduced gradually beyond Mach 5 in order to suppress combustion pressure. We would like to proceed with the ¯ight test under international cooperation. 5. CONCLUSIONS
Fig. 25. Air Inlet and Outlet Temperature and Maximum Wall Temperature of Precooler in Flight
1. Mixer of ATREX combustor has two functions, that is mixing and ¯ame holding without any other ¯ame holder, and therefore combustion
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performances depend only on mixer con®guration. Six dierent con®guration mixers were examined in sub scale tests as well as CFD analyses. It could be found from this study that the mixer to generate swirl mixing ¯ows by its skewed lobes could reduce combustion chamber. Of these mixers with dierent skewed angles, the skewed angle of 15 degrees gives best performances. In this case combustion is completed at distance of 1.3 D from mixer outlet end. (D: combustion diameter) The swirl mixer with 15 degrees of skewed angle is applied to ATREX500 and will be veri®ed by test in 195. 2. Three types of axisymmetric variable air intake were tested in a wind tunnel up to Mach 4 to make clear their performances (pressure recovery factor, mass capture ratio and eect of air bleeding) and also to establish the design method. Pressure recovery factors of each model are being improved to meet the target of ATREX engine. Bleeding in center spike should be larger in order to suppress boundary layer. Bleeding in cowl should be made with precise amount and position. 3. Analytical and design study of air precooling for the ATREX engine showed that the outlet air temperature T = 160 K and the air pressure recovery s = 0.9 are almost optimum at equivalence mixture ratio of 1.3±1.4 under SL-static conditions. 4. When the fan inlet air temperature is cooled down to 160 K with the pressure recovery factor of s = 0.9 in the precooler, the thrust and speci®c impulse of the existing ATREX-500 engine are expected to be increased by 2 and 1.5 times respectively at SL-static conditions compared to the nonprecooled one.
5. Target thrust and speci®c impulse of the modi®ed ACC ATREX engine with fan inlet diameter of 0.3 m for ¯ight test are estimated to be 1730± 1930 kgf and 3150±3300 sec at SL-static condition. 6. Two types of precooler, designated ``Blin'' and ``Baraban'', to integrate in the existing ATREX500 engine were designed. Of which ``Baraban'' type precooler were manufactured and will be tested by integrating to ATREX-500 in September of 1995. 7. Based on the physical model of ice formation which was veri®ed by the fundamental experiment, the icing borders in the ¯ight speed-altitude map were estimated. In December to March no-icing window was expected around the launch site candidate of Taiki. Maximum duration of icing formation along the typical ¯ight trajectory was estimated to be as long as 25±40 sec in the most humid season. 8. We would like to proceed with the ¯ight test under the international cooperation.
REFERENCES
1. Tanatsugu, N., Sato, T. et al., DEVELOPMENT STUDY ON ATREX ENGINE, 45th IAF Congress, Jerusalem, Oct., 1994. 2. Balepin, V. V., Tanatsugu, N., Some Considerations of Precooler for ATREX Engine, Proceedings of the Space Transportation Conference, ISAS, Jan 1995. 3. Balepin, V. V., Tanatsugu, N., Sato, T., Mizutani, T., Hamabe, K. and Tomike, J., Development Study of Precooling for ATREX Engine, XII ISABE Symp., Melbourne, 1995. 4. Balepin, V. V., Folomeev, E. A., Galkin, S. M. and Tjurikov, E. V., Rocket Based Combined Cycles for Vertical Take-o Space Vehicles, AIAA Paper 956078, 1995.
Acta Astronautica Vol. 41, No. 12, pp. 851±862, 1997 # 1998 Published by Elsevier Science Ltd. All rights reserved Printed in Great Britain 0094-5765/98 $19.00 + 0.00 S0094-5765(97)00176-8
DEVELOPMENT STUDY ON ATREX ENGINE{ N. TANATSUGU, T. SATO, V. BALEPIN, Y. NARUO Institute of Space and Astronautical Science, 3-1-1, Yoshinodai, Sagamihara, Kanagawa 229, Japan
T. MIZUTANI, T. KASHIWAGI Ishikawajika Harima Heavy Industries CO., LTD., 3-5, Mukodai-cho, Tanashi-shi, Tokyo 188, Japan
K. HAMABE, J. TOMIKE Kawasaki Heavy Industries, LTD., 2-8-1, Takatsukadai, Nishi-ku, Kobe, 673-02, Japan
and R. MINAMI Mitsubishi Heavy Industries, LTD., 10, Oye-cho, Minato-ku, Nagoya 455, Japan (Received 2 March 1996) AbstractÐThis is the status report of the development study on ATREX engine (Air Turbo Ramjet) that is now under way in the Institute of Space and Astronautical Science (ISAS) cooperation with the Ishikawajima Harima Heavy Industries (IHI), the Kawasaki Heavy Industries (KHI), the Mitsubishi Heavy Industries (MHI). ATREX engine will be applied for the propulsion system of ¯y-back booster of TSTO space plane. ATREX is the combined cycle (a fan-boosted ramjet) engine providing the eective thrust from sea level static to ¯ight Mach number 6. ATREX is worked on the expander cycle with precooling the incoming air as shown in Fig. 1. ATREX employs the tip turbine con®guration which allows the compactness and the light weight of turbo machinery and the variable geometry airintake and plugnozzle which allow the wide range operation conditions. From 1990 to 1992, `` ATREX-500`` has been tested at the sea level static conditions. ATREX-500 is the 1/4-scale model of which fan inlet diameter is 300 mm and overall length 2,200 mm. From 1992 have been performed the wind tunnel tests on the primary components of ATREX, the axisymmetric variable geometry airintakes, the precoolers and the variable geometry plug nozzles. In parallel to the windtunnel tests, the ram combusters have been tested simulating the hypersonic ¯ight conditions and the application studies on advanced carbon-carbon composite for the tip-turbine and fan assembly has been proceeded. In 1994 initiated the ¯ight test plan in which ATREX will be veri®ed in the practical ¯ight conditions by using an unmanned ¯ying test bench. In 1995 will be tested ATREX-500 installing the precooler under the sea level static conditions to examine the engine performance and the icing on the precooler. The present paper addresses the high loading ram combuster experiment using the mixer with skewed lobes to generate swirl ¯ow and the analytical studies and the designs on the precooler and the precooled ATREX engine and the ¯ight test plan. # 1998 Published by Elsevier Science Ltd. All rights reserved
1. HIGH LOADING RAM COMBUSTOR
lobes make the swirl ¯ow of hydrogen and air in the mixing process after injecting into the combuster. This swirl ¯ow improves the combustion performance by prolonging the stay time for mixing and combustion.
The compact ram combuster is required strongly to reduce the size and weight of ATREX. The compactness of ram combustor is accomplished by the high speci®c combustion loading. The combustion performance depends strongly upon the mixer and thus the improvement of mixer is well-worn approach to increase the combustion loading. The mixers with the skewed lobes were designed and tested for the ATREX combuster. The skewed
1.1. Mixer Con®guration with Skewed Lobes
{Paper IAF-95.S5.01 presented at the 46th International Astronomical Congress, October 2±6, 1995, Oslo, Norway 851
Five types of sub-scaled mixer with dierent skewed lobes were tested. Three have dierent skew angles (Type I, Type I-B15, Type I-B20) and another has dierent axial length (Type VI-B15) and the other has dierent lobe shape (Type VIII-B15) as shown in Fig. 2. All mixers have the 16 lobes which divide the ¯ow of hydrogen and air by the ratio of
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Fig. 1. ATREX Engine Flow Diagram
1:4 in cross-sectional area. The mixers tested are scaled down to 1/5 from one of the ATREX-500. 1.2. Test Result of Mixers with Skewed Lobes 1.2.1. Speci®c Combustion Loading. Figure 3 shows the ¯ame temperature distribution indicated
by the nondimensional distance from the combustion chamber normalized by the combustion chamber inner diameter D. They were measured by the dichromatic pyrometer scanning along an axial direction at 0.38D apart from the center of chamber. It is seen from this friuge that the mixers
Fig. 2. Mixer Co®gurations with Skewed Lobes
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Fig. 3. Flame Temperature Distribution Fig. 4. Skewed Lobes Mixer for ATREX-500
with skewed lobes give the higher maximum combustion temperature compared to the non-skewed mixer (Type I), and the maximum temperature point is located around 1.0±1.2D from the combustion chamber inlet which is more upstream compared to 1.8D of the non-skewed mixer. From these results the signi®cant enhancement on speci®c combustion loading can be obtained by the mixers with skewed lobes (especially Type VIII-B15) by a factor of approximately 1.5 compared to the non-skewed mixer (Type I). 1.2.2. Flame Visualization. The combustion ¯ame structure has been analyzed by ultraviolet photography of OH-radical luminescence. The intensity distribution of ultraviolet photography is similar to the temperature distribution. It can be also concluded from this fact that the hydrogen-air mixing is enhanced by the swirl ¯ow made by the mixer with skewed lobes and the speci®c combustion loading increases. 1.2.3. Pressure Loss. The pressure drop through the combustion chamber was not so largely increased by the mixers with skewed lobes compared with non-skewed mixer (Type I). 1.3. Mixer with Skewed Lobes for ATREX-500 Mixer with skewed lobes for ATREX-500 shown in Fig. 4 was designed based on the test results of sub-scaled models. This mixer is integrated in ATREX-500 as well as the Baraban type precooler
and will be tested in September 1995 at Noshiro Testing Center of ISAS.
2. AXISYMMETRIC VARIABLE GEOMETRY AIR INTAKE
Axisymmetric air intakes have been studied using supersonic wind tunnel in ISAS from 1993 to establish design method and make clear changes of performance characteristics by bleeding to suppress boundary layers. Design Mach number shifted from Mach 3 in 1993, Mach 3,5 in 1994 to Mach 4.5 in 1995. [1] 2.1. Test Model Figure 5 shows the model for wind tunnel test. Center spike is movable to be adjusted for the given Mach number. Outer diameter of test model is 127 mm which is one sixth of ATREX engine for ¯ight test. Air bleeding is made from small holes on center spike and slits on cowl. Amount of bleeding was ®xed during blow in this test though it has to be adjusted for Mach number of incoming air stream in real ¯ight. Performances decrease abruptly and shift to unstart condition beyond design Mach number especially in the air intake with larger internal compression and therefore design point should be set at maximum ¯ight Mach number. Mach number of 4.5 is set for the present de-
Fig. 5. Axisymmetric Variable Geometry Air Intake Model for Wind Tunnel Tests
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N. Tanatsugu et al. Table 1. Features of Air Intake Models
Air Intake Model Compression Mode External Internal Characteristics Intake Length Boundary Layer Eect Startability Stability Drag in Cowl Lip Pressure Recovery at Mach 3.5 Mass Capture Ratio at Mach 3.5
Type E
Type F
Type G
2 2 High Startability Short Small Good Large 0.56 0.33
1 + Isentropic 3 High Internal Compression Very Long Large Ordinary Small 0.60 0.70
1 + Isentropic 3 High Internal Compression Long Large Ordinary Small 0.49 0.58
sign. Type E, F and G models were designed by means of method of characteristic. Compression is relatively larger in external section of type E model and in internal section of type F and G. Features of these models are summarized in Table 1. 2.2. Test Results Maximum pressure recovery factor is shown in Fig. 6 together with the target curve required by ATREX engine and MIL-spec. curve. All models were tested below Mach 4 which is maximum available blow in ISAS' wind tunnel. Type E model meets the target value below Mach 4. Type F values are smaller in lower and higher Mach number and therefore type F is estimated to be not good beyond Mach 4. Although type G values are relatively lower from 1.5 to 3.5 of Mach number, at Mach 4 it is larger than at Mach 3.5 and estimated to be better beyond Mach 4. Mass capture ratio is 70, 33 and 58 percents at Mach 3.5 for type F, E and G respectively. Eect of bleeding in type F model is shown in Fig. 7. Operation of this test was done as follows: (1) center spike moves backward gradually from foremost forward position and stops at appropriate position, and then (2) increases back pressure gradually by closing plug valve located at discharge port of air intake, and suddenly (3) falls into unstart condition. In type F model, mass capture ratio decreases keeping maximum pressure recovery factor as bleeding in cowl increases. Startablity becomes worse without cowl bleeding and leads to reduced performances. From these results, there is
Fig. 6. Pressure Recovery Factor of Air Intakes
the optimum amount of bleeding in cowl. In case of decreasing air bleeding in center spike, performances become worse due to not decreasing throat area. Boundary layer eects are larger due to larger surface area of center spike. Performances of type G model are best of all models when cowl bleeding is not done. As center spike bleeding increases, performances of type G model become better as well as type F model. 2.3. Conclusion Pressure recovery factor of each type of air intake models are being improved to meet the target of ATREX engine. Bleeding in center spike should be larger in order to suppress boundary layer. Bleeding in cowl should be made with precise amount and position. 3. AIR PRECOOLER
Generally speaking, precooling gives the following bene®ts compared to non-precooled engines: Ð increase of thrust; Ð increase of speci®c impulse; Ð extension of ¯ight range to higher Mach number. These eects could be attained by several manners and/or their combinations: Ð air¯ow increase; Ð increase of fan pressure ratio; Ð decrease of hydrogen ¯ow rate caused by reduction of fan driving power.
Fig. 7. Eect of Air Bleeding in Type F Model
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Fig. 9. Air intake and precoolers intagration: ``blin'' type (upper), ``baraban'' type (lower)
Fig. 8(a). Thrust dependcy on fan inlet temp. and precooler pressure recovery factor (ATREX-500) (b) Speci®c impulse dependcy on fan inlet temp. and precooler pressure recovery factor (ATREX-500)
In order to augment the thrust of ATREX especially below transonic ¯ight, the precooler operation is changed to work continuously from lift-o. In the early design the precooler operates above Mach 2 of ¯ight speed to prevent ice formation on it. On the other hand some problems come out by employing the precooler. Air pressure drop across the precooler has negative eect on engine performance. In order to reduce this pressure loss, the precooler structure becomes larger and heavier. Figures 8a and 8b shows the impact of precooler outlet air temperature on the SL-static performance of ATREX-500 for the precoolers with dierent air pressure recovery factors. ATREX-500 engine with air cooling temperature of 160 K and pressure recovery factor of 0.9 provides twice in the thrust and 1.5 times in speci®c impulse compared to the nonprecooled one. It is due to the fact that air ¯ow rate increases from 6.7 kg/s to 11.5 kg/s and fan pressure ratio from 1.53 to 2.03 by precooling in the same turbomachinery. According to Fig. 8 thrust sensitivity is about 1.0% per 1 K of precooling temperature change and 2.5% per 1% of pressure recovery change around air cooling temperature of 160 K and pressure recovery factor of 0.9. For speci®c impulse these ®gures are 0.4% and 1.4% respectively. It is
seen that the engine performance is aected signi®cantly by cooling temperature and pressure recovery of the precooler. In the beginning of design study of the precooler some key targets were set on precooler performance characteristics and con®guration as follows [2]: Ð the main purpose is air precooling to 160 K with air pressure recovery factor 0.85 and higher at SL-static conditions; Ð precooler should be a shell/tube con®guration with the tube diameter 3 mm and all thickness 0.1± 0.15 mm; Ð precooled ATREX engine should be operated under more stoichiometric mixture ratio; Ð icing free regime of precooler operation should be provided from SL-static conditions. A number of the precooler con®gurations were examined [2], of which two con®gurations were selected for technology development. Figure 9 shows the precoolers integrated in the axisymmetric air intake of ACC ATREX engine. This air intake con®guration is adaptive for the ¯ight of Mach number up to 6. Turbo-machinery is assembled on the right of the precooler shown in Fig. 9. These two precoolers are designated to be ``blin'' and ``baraban'' in Russia, which means ``pancake'' and ``drum'' respectively. ``Blin'' type precooler provides pressure recovery factor 0.86 at design point of 160 K. In ``Baraban'' type this ®gure is 0.94 which is higher than ``Blin'' type. This higher pressure recovery of ``Baraban'' can provide the thrust increase as large as 20% and the speci®c impulse increase as high as 12% compared to ``Blin'' type. Figure 10 shows the performance characteristics of ATREX-500 with ``Baraban'' or ``Blin'' type precooler at SL-static conditions. Thrust (top) and speci®c impulse (bottom) are indicated as contour lines in the coordinates of precooler outlet air temperature and air pressure recovery factor of precooler.
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Fig. 11. Relative Thrust to Weight Ratio based on Nonprecooled ATREX-500
Fig. 10. Performance of precooled ATREX-500 engine at SL-static and characteristics of precoolers
Figure 11 shows the relative thrust to weight ratio of ATREX-500 precooled with Baraban precooler based on non-precooled ATREX-500 (fan inlet air temperature: 288 K, air pressure recovery factor: 1.0) The maximum point locates at fan inlet temperature T around 160 K and air pressure recovery factor s above 0.95. Figures 10 and 11 are the key ®gures to decide the air precooling temperature. It is seen that in ``Baraban'' type precooler the almost maximum thrust and speci®c impulse are given at the design point (T = 160 K, s = 0.94) on the characteristic curves. In this case the decision of design point is optimum. The design point decided for ``Blin'' type precooler (T = 160 K, s = 0.86) provides about maximum thrust, but its speci®c impulse could be
improved by about 3% with a little thrust change if precooling temperature is selected as T = 170± 175 K. By this modi®cation the pressure recovery factor is improved to be 0.90±0.91 and the better engine performance can be attained and its weight could be reduced by about 30%. Needless to say that the con®dential design point should be proved by the tests. It should be noted that the approach mentioned here can provide the selection of maximum engine performance within the geometric restriction required. As the vehicle accelerates, the pressure recovery factor of precooler comes close to 1.0 very quickly and therefore the degree of air precooling is important rather than the pressure recovery factor. Taking into account the total optimization of fuel consumption required in the whole mission, more complicated analysis is necessary for the selection of precooling level at SL-static conditions. It will be one of the subjects in the next step of the study following the veri®cation of design by the SL-static test of precooled ATREX-500 to be held in September and November 1995. More information on precooler con®guration and performance were given in Ref. [3]. Table 2 summarizes the key parameters of SLstatic test results of the nonprecooled ATREX-500 and the estimated parameters of precooled ATREX-500 equipped with ``Blin'' and ``Baraban'' precoolers.
Table 2. Comparison of ATREX-500 performances at SL-static conditions nonprecooled test results, expected engines values precooled with ``Blin'' and ``Baraban''
Thrust [kgf] Speci®c impulse [sec] Air ¯ow rate [kg/s] H2 ¯ow rate [kg/s] Equiv. mixture ratio Fan pressure ratio Turbine inlet temp. [K] Fan inlet temp. [K] Fan inlet pressure [atm]
ATREX-500 nonprecooled (test results)
ATREX-500 with ``Blin'' precooler (design values)
ATREX-500 with ``Baraban'' precooler (design values)
427 1406 6.74 0.3 1.52 1.53 564 291 1.033
735 1759 10.9 0.416 1.31 2.03 650 160 0.883
909 1975 11.5 0.442 1.32 2.03 650 160 0.968
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Table 3. Correlation coecient of precooler Coecient
Precooler Type ``Blin'' type ``Baraban'' type 0.98 0.82 2.21 0
X1 X2 X3 X4
0.85 0.74 2.21 ÿ0.195
3.1. Quick estimation of precooler performance For quick estimation of the precooler performances, the simple and convenient correlations can be found in the precise calculation results. The precooler temperature eciency Zair is de®ned to be the ratio of the temperature dierence between air inlet and outlet and the temperature dierence between air inlet and hydrogen inlet: Zair
out Tin a ÿ Ta in Ta ÿ Tin h
1
where T: temperature superscript; in:inlet out: outlet subscript; a:air h:hydrogen The air pressure recovery factor sair is de®ned to be the ratio of air stagnation pressures in the outlet and the inlet of precooler: in sair pout a =pa
2
where P:stagnation pressure Superscripts; in:inlet out:outlet subscript; a:air These two parameters correlate closely with the dimensionless ¯ow rates as follows: X2 1 Zair Z0 gX a =gh
Fig. 13. Air pressure Recovery Factor of Baraban Precooler
3
Fig. 12. Air Outlet Temperature of Baraban Precooler
X4 3 sair 1 ÿ 1 ÿ s0 gX acorr gh
ga G0a ; gh Gh =G0h ; gacorr
0 s1 Ga @ p0 Tin a A 0 T0 Ga pin a 4
where Z:temperature eciency s:pressure recovery factor G:¯ow rate P:stagnation pressure T:stagnation temperature superscripts; in:inlet 0:value at design point subscripts; a:air h:hydrogen gacorr:relative corrected air ¯ow rate T0=288.15 K (158C) P0=101.3 kPa X1, X2, X3, X4: exponent coecients determined by the geometric con®guration of precooler.
Fig. 14. Hydrogen Outlet Temp. of Baraban Precooler
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Fig. 15. Con®guration of ATREX-500 with Baraban Precooler set up on test stand
These coecients X1, 2, 3, 4 become almost constant for the speci®c type of precooler over the wide range of conditions from SL-static to hypersonic ¯ight around Mach number of 6.0 at altitude as high as 35 km. The estimation by eqns (3) and (4) is satis®ed with the precise calculation within 3% of deviation. The coecients X1, 2, 3, 4 for the ``Blin'' and ``Baraban'' type precooler are given in Table 3. Figures 12±14 show the contour lines of air outlet temperature, air pressure recovery factor and hydrogen outlet temperature of the Baraban precooler on the air and hydrogen ¯ow rate map which are calculated by the quick coreration equations mentioned above. This calculations are made under the following conditions, air inlet temp.: 288.2 K, air inlet press.: 0.013 MPa, hydrogen inlet temp.: 30 K, hydrogen inlet press.: 4 MPa. Figure 15 shows the con®guration of ATREX500 with Baraban precooler when it will be tested in 1995. 3.2. Precooler Icing The icing problem in the heat exchanger for the atmospheric air cooling was pointed out by the specialists from the beginning of the study. According to the study conducted in CIAM (Russia), icing free operation of the precooled engines is examined with the physical model of ice formation based on the experimental veri®cation [4]. Two ice formation mechanisms have been discovered depending on the steam partial pressure 1. Ice formation on micro level. When steam partial pressure is lower than steam pressure at the triple point (ptr=0.006228 atm), frost is formed avoiding liquid phase. This process is indicated by the arrow crossing sublimation line AO in PT diagram of water shown in Fig. 16. In this case frost layer is growing by molecular level. Formed frosting substance is not so strong and always carried over by air stream.
A special test conducted at pst
Fig. 16. P-T diagram for water stream, AO-sublimation line; OB-solidi®cation line; OC-saturation line.
Development study on ATREX engine
Fig. 17. Seasonal borders of icing. Humidity data for Nemuro area (north-east Hokkaido)
pst
5
Ta < 273K
6
eqns (5) and (6) allow us to draw the borders to prevent icing in the speed-altitude map of ¯ight path. Atmospheric moisture content is maximum at sea level and rapidly decreases as the altitude increases. Icing borders concerning to ¯ight speed-altitude can be decided by the humidity distribution along the altitude and the stagnation temperature of incoming atmospheric air. Figure 17 shows the icing borders caused by the temperature condition given by eqn (6) and the typical trajectories of ¯ying test bed powered by ACC or metal ATREX, where are used the climate data taken at Nemuro which is the nearest city to the Taiki. Icing borders caused by pressure condition given by eqn (5) are located higher as seen in Fig. 17, and therefore the precooler icing is unlikely at the altitude about 0.5 km in November and April and at
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Fig. 18. Times to reach icing border on the seasons
the altitude about 3.5±4.0 km in July and August which is the highest humid season. According to the conclusion given by eqn (6), there is no icing on the precooler in winter season when air temperature is below 273 K at any altitude. This condition takes place from December through March around the city of Nemuro. In the rest of the year there is the icing possible from sea level to the icing borders. Conversion of the trajectory data into the time sequence gives the period of reaching the icing border which is shown in Fig. 18 for each month. It is seen that June to September is a less favorable season for the precooled engine operating for 25±30 sec to reach the icing border in the ¯ight powered by ACC ATREX and for 35±40 sec by the metal ATREX. Further analytical and experimental study is necessary to substantiate the possibility of a precooled engine to be able to operate for 25±40 sec without negative eects by icing. 4. FLIGHT TEST PLAN ON ATREX ENGINE
In the ®nal phase of the ATREX engine development study, the overall system performance, functions and operations of ATREX will be veri®ed by the actual ¯ight test. The ¯ight test will be performed by using the Flying Test Bed (FTB) which is powered by ATREX engine itself after lift-o.
Fig. 19. Flight Pro®le of ATREX-500 Flight Test
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Fig. 20. Con®guration of FTB Table 4. Speci®cations of FTB for ATREX Flight Test Weight [kg] ATREX Engine Structure Electronics Fuel Supply System Control Drive System Recovery System Fuel (LH2) Total Max. Flight Mach Number Max. Altitude [km] Max. Dynamic Pressure [kPa] Wing Loading [kg/m2] Fuel Tank Pressure [MPa] Fuel Tank Volume D L [m]
370 458 80 112 90 100 190 1400 >6 >30 50 600 0.3 0.8 6.6
The entire ¯ight pro®le is shown in Fig. 19. The FTB takes o horizontally with the take-o assist system running on the rail up to 0.4±0.5 of Mach number. The takeo assist system is powered by a conventional turbo jet engine in parallel with ATREX engine. In the early phase of the ¯ight test, the maximum ¯ight speed is up to 4.5 of Mach number with the metal turbo machinery. The ®nal target speed is Mach 6 which can be attained with the carbon-carbon composite turbo machinery. After reaching the maximum speed planned, FTB is decelerated by throttling engine power and glides down and ®nally lands on the sea by parachute for recovery. The FTB ¯ies by autonomous and programmed control in addition to the radio guidance system with GPS. The data obtained during ¯ight are transmitted by means of telemeter. Con®guration of FTB and its speci®cations are shown in Fig. 20 and Table 4.
Fig. 21. Improvement of Thrust of ACC and Metal ATREX by Precooling at SL-Static Conditions
Two ATREX engines are designed for ¯ight tests; one is the metal ATREX which is available for ¯ights up to Mach 4.5 and the other is the ACC ATREX which can attain the ®nal goal up to Mach 6. They have same size as in ATREX-500 developed for SL-static tests. Their design data are shown in Table 5. Engine performances are improved in ACC ATREX primarily due to the improvement of turbo-machinary caused by its higher peripheral speed at higher temperature compared to metal ATREX. Their performances are also improved by precooling shown in Figs 21 and 22 compared to the test results of non-precooled ATREX-500 indicated in these ®gures. The eects of air precooling can be seen on the ¯ight conditions shown in Figs 23 and 24, where the thrust and speci®c impulse are indicated for two air precooling temperatures 160 K and 220 K at SL-static conditions. The bene®t of precooling temperature 160 K is 18±25% in speci®c impulse and 25±40% in thrust compared to 220 K. Figure 25 shows air inlet and outlet temperatures and wall temperature of precooler in ¯ight conditions along the typical ¯ight path of FTB. Decrease of speci®c impulse beyond Mach 3.5 is caused by increase of hydrogen ¯ow rate for cooling purposes although thrust is increased. Typical ¯ight path of FTB powered by ACC and metal ATREX, arrival time of FTB to a given ¯ight speed and dynamic pressure along its ¯ight path are
Table 5. Design Data of ATREX Engine Made of ACC and Meatl for Flight Test Component Precooler Fan
Turbine
Combuster Heat Exchanger Thrust Nozzle
Design Parameter Air Pressure Drop [%] Outlet Air Temperature at S.L. Static [K] Pressure Ratio Eciency [%] Peripheral Speed [m/s] Inlet Diameter [mm] Inlet Hydrogen Temperature [K] Pressure Ratio Eciency [%] Peripheral Speed [m/s] Pressure Drop [%] Combustion Eciency [%] Combustion Gas Pressure Drop [%] Discharge Gas Velocity Coecient
ACC Engine 16 160 3.8 80 420 300 1500 6 40 500 5 80 8 0.97
Design Value
Metal Engine 12 160 2.7 80 320 300 650 5 40 380 5 80 8 0.97
Development study on ATREX engine
Fig. 22. Improvement on Speci®c Impulse of ACC and Metal ATREX by Precooling at SL-Static Conditions
Fig. 23. Thrust of ACC ATREX in Flight Conditions
Fig. 24. Speci®c impulse of ACC ATREX in Flight Conditions
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Fig. 26. Typical Flight Path of FTB Powered by ACC and Metal ATREX Engine
Fig. 27. Arrival Time of FTB Powered by ACC and Metal ATREX Engine along Flight Path Indicated in Fig. 22
Fig. 28. Dynamic Pressure along Flight Path Indicated in Fig. 22
shown in Fig. 26±28. This ¯ight test needs approximately 170 to 180 km in range and 190 to 260 seconds in time to reach the target ¯ight speed for ACC and metal ATREX powered ¯ight respectively. Dynamic pressure is kept 50 kPa from 1 to 5 of Mach number and reduced gradually beyond Mach 5 in order to suppress combustion pressure. We would like to proceed with the ¯ight test under international cooperation. 5. CONCLUSIONS
Fig. 25. Air Inlet and Outlet Temperature and Maximum Wall Temperature of Precooler in Flight
1. Mixer of ATREX combustor has two functions, that is mixing and ¯ame holding without any other ¯ame holder, and therefore combustion
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performances depend only on mixer con®guration. Six dierent con®guration mixers were examined in sub scale tests as well as CFD analyses. It could be found from this study that the mixer to generate swirl mixing ¯ows by its skewed lobes could reduce combustion chamber. Of these mixers with dierent skewed angles, the skewed angle of 15 degrees gives best performances. In this case combustion is completed at distance of 1.3 D from mixer outlet end. (D: combustion diameter) The swirl mixer with 15 degrees of skewed angle is applied to ATREX500 and will be veri®ed by test in 195. 2. Three types of axisymmetric variable air intake were tested in a wind tunnel up to Mach 4 to make clear their performances (pressure recovery factor, mass capture ratio and eect of air bleeding) and also to establish the design method. Pressure recovery factors of each model are being improved to meet the target of ATREX engine. Bleeding in center spike should be larger in order to suppress boundary layer. Bleeding in cowl should be made with precise amount and position. 3. Analytical and design study of air precooling for the ATREX engine showed that the outlet air temperature T = 160 K and the air pressure recovery s = 0.9 are almost optimum at equivalence mixture ratio of 1.3±1.4 under SL-static conditions. 4. When the fan inlet air temperature is cooled down to 160 K with the pressure recovery factor of s = 0.9 in the precooler, the thrust and speci®c impulse of the existing ATREX-500 engine are expected to be increased by 2 and 1.5 times respectively at SL-static conditions compared to the nonprecooled one.
5. Target thrust and speci®c impulse of the modi®ed ACC ATREX engine with fan inlet diameter of 0.3 m for ¯ight test are estimated to be 1730± 1930 kgf and 3150±3300 sec at SL-static condition. 6. Two types of precooler, designated ``Blin'' and ``Baraban'', to integrate in the existing ATREX500 engine were designed. Of which ``Baraban'' type precooler were manufactured and will be tested by integrating to ATREX-500 in September of 1995. 7. Based on the physical model of ice formation which was veri®ed by the fundamental experiment, the icing borders in the ¯ight speed-altitude map were estimated. In December to March no-icing window was expected around the launch site candidate of Taiki. Maximum duration of icing formation along the typical ¯ight trajectory was estimated to be as long as 25±40 sec in the most humid season. 8. We would like to proceed with the ¯ight test under the international cooperation.
REFERENCES
1. Tanatsugu, N., Sato, T. et al., DEVELOPMENT STUDY ON ATREX ENGINE, 45th IAF Congress, Jerusalem, Oct., 1994. 2. Balepin, V. V., Tanatsugu, N., Some Considerations of Precooler for ATREX Engine, Proceedings of the Space Transportation Conference, ISAS, Jan 1995. 3. Balepin, V. V., Tanatsugu, N., Sato, T., Mizutani, T., Hamabe, K. and Tomike, J., Development Study of Precooling for ATREX Engine, XII ISABE Symp., Melbourne, 1995. 4. Balepin, V. V., Folomeev, E. A., Galkin, S. M. and Tjurikov, E. V., Rocket Based Combined Cycles for Vertical Take-o Space Vehicles, AIAA Paper 956078, 1995.