[American Institute of Aeronautics and Astronautics 16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference - Bremen, Germany ()] 16th AIAA/DLR/DGLR

American Institute of Aeronautics and Astronautics

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Firing Test of a Hypersonic Turbojet Engine Installed on a Flight Test Vehicle

Hideyuki Taguchi 1, Kenya Harada2 , Hiroaki Kobayashi 3, Takayuki Kojima 4, Motoyuki Hongoh 5, Daisaku Masaki 6, Shujiro Sawai 7, Yusuke Maru 8 Japan Aerospace Exploration Agency, Chofu, Tokyo 182-8522, Japan

and

Tetsuya Sato9 Waseda University, Shinjuku, Tokyo 169-8555, Japan

Hypersonic turbojet engine with pre-cooling system is tested under sea level static condition. The engine is installed on a flight test vehicle, which will fly at Mach 2 speed by a free fall experiment from a stratospheric balloon. Liquid hydrogen fuel and gas hydrogen fuel is supplied to the engine from a tank and cylinders installed in the vehicle. Designated operation of major components of the engine is confirmed. Corrected rotation speed, corrected air flow rate and pressure ratio of the compressor is raised by pre-cooling with liquid hydrogen fuel. Corrected air flow rate and pressure ratio at the pre-cooling operation is reduced comparing from that without pre-cooling on the same corrected rotation speed. There is a deep temperature distortion at the inlet of the compressor and it may cause the performance reduction. Large amount of liquid hydrogen is supplied to the pre-cooler in order to obtain enough pre-cooling performance for Mach 5 flight. Then, fuel rich combustion at the after-burner is adopted. Cowl part of variable geometry nozzle is made with C/C composite material and it has no damage after the combustion test. Operation of the core engine by liquid hydrogen is attained by using a control valve with small effective diameter.

Nomenclature BG = test code for gas hydrogen operation BGP = test code for gas hydrogen operation with pre-cooling BL = test code for liquid hydrogen operation N = mechanical rotation speed Nc = corrected rotation speed Ga = compressor air flow rate Gac = corrected compressor air flow rate Gfmb = main burner fuel flow rate Tpci = coolant inlet temperature Tpco = coolant outlet temperature Tci-1 , Tci-2 = compressor inlet temperatures Tci-mean = compressor inlet average temperature 1 Aviation Program Group, Japan Aerospace Exploration Agency, Chofu, Tokyo 182-8522, Japan, AIAA Member. 2 Aviation Program Group, Japan Aerospace Exploration Agency. 3 Aerospace Research and Development Directorate, Japan Aerospace Exploration Agency. 4 Aerospace Research and Development Directorate, Japan Aerospace Exploration Agency, AIAA Member. 5 Aerospace Research and Development Directorate, Japan Aerospace Exploration Agency. 6 Aerospace Research and Development Directorate, Japan Aerospace Exploration Agency. 7 Associate Professor, Institute of Space and Astronautic Sciences, Japan Aerospace Exploration Agency. 8 Assistant Professor, Institute of Space and Astronautic Sciences, Japan Aerospace Exploration Agency. 9 Professor, Faculty of Science and Engineering, Waseda University, Shinjuku, Tokyo 169-8555, Japan.

16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conferenc AIAA 2009-7311

Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.


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Tmb = main burner temperature Tab = after burner temperature πc = compressor pressure ratio πt = turbine pressure ratio

I. Introduction re-cooled turbojet engines and scramjet engines have been investigated aiming at realization of hypersonic aircrafts. Evaluation methods of these engine performances have been established based on ground tests. X-

43A1 flight experiment was successfully conducted to demonstrate scramjet engine technologies. There are some following plans on the demonstration of hypersonic propulsion systems.

JAXA focused on hypersonic propulsion systems as a key technology of hypersonic transport aircraft2 (Fig. 1). Demonstrations of Mach 5 class hypersonic technologies are stated as a development target at 2025 in the long term vision3.

The design and fabrication of hypersonic turbojet4 that operates from take off to Mach 5 has been advanced. This engine is characterized in cooling the air that becomes a high temperature by a high Mach number flight. A pre-cooler which uses cryogenic liquid hydrogen as coolant is connected in the upstream of the core engine. The high Mach number flight is enabled by protecting the core engine from heat by pre-cooling. Furthermore, the compression power is decreased by pre-cooling so as to realize positive power balance at high Mach number flight.

The performance of hypersonic turbojet with pre-cooler is shown by performance analyses4. A supersonic flight of Mach 2 will be achieved by a balloon-based operation vehicle5. The vehicle will be separated from a stratospheric balloon at 40 km altitude. The vehicle will reach the speed of Mach 2 by a free dropping flight. A small hypersonic turbojet6 will be installed on the vehicle and its performance data will be obtained on the flight experiment.

In this study, combustion test of the hypersonic turbojet installed on the balloon-based operation vehicle is conducted on ground (Fig. 2). The experiment aims at the confirmation of equipments on the vehicle and evaluation of the engine performance under sea level static condition. In addition, the start sequence of the core engine is verified using liquid hydrogen fuel.

Figure 1. JAXA Hypersonic Transport Aircraft. Figure 2. Firing test of Hypersonic Turbojet Engine.

II. Characteristics of Hypersonic Turbojet Engine The cross section of the balloon based operation vehicle is shown in Fig. 3. Liquid hydrogen tank, hydrogen gas

cylinder, helium gas cylinder, measurement and control equipments are installed in the experimental vehicle. A hypersonic turbojet is supported with a load cell mount to measure thrust in the direction of the airframe axis under the experimental vehicle. Figure 4 shows cross section of a hypersonic turbojet engine used in this experiment. A pre-cooler using liquid hydrogen fuel is installed in front of the core engine. The size of the engine is determined considering the wind tunnel experiment in the Ramjet Engine Test Facility (RJTF) at JAXA Kakuda Space Center. Table 1 shows specifications of hypersonic turbojet engine. Table 2 shows main materials for each engine components.

The intake is produced by using aluminum alloy to adapt Mach 2 flight condition of balloon-based operation vehicle. It will be replaced to a part produced with C/C composite material for the future flight experiment at Mach 5, because the temperature exceeds the melting point of aluminum alloy.

P


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Pre-cooler heat exchanger is a counter-flow shell and tube type one. In order to avoid brittle destruction by hydrogen and to attain light weight structure, 2mm diameter tube made of thin stainless steel is selected for the heat exchanger. The casing is produced with the aluminum alloy as well as the intake corresponding to Mach 2 flight experiment.

The core engine is an one spool turbojet with the design rotation speed of 80000rpm. The pressure ratio is 6, and the compressor air mass flow rate at the design point is 1.0kg/s. The compressor is produced with titanium alloy to endure the stress at the design rotation speed. A main combustor and the turbine are produced with nickel base alloy. Nominal combustion temperature of a main combustor is limited below 1223K, because the turbine is produced with brisk without cooling holes. The casing and the exit duct of the core engine are made of stainless steels.

The regenerative cooling part of a variable geometry nozzle is made of nickel alloy. The movable ramp part and the external expansion part are made of stainless steels. The fuel injector of the after-burner is placed in the upstream of the exhaust nozzle. Nozzle cowl part is made of C/C composite material with Si impregnated.

Measurement andControl Equipments

LH2 Fuel Tank GHe / GH2 Cylinders

Hypersonic Turbojet Engine

Figure 3. Balloon-based Operation Vehicle.

Air Intake Pre-Cooler Core Engine After-Burner Exhaust Nozzle

Figure 4. Hypersonic Turbojet Engine.

Table 1. Specifications of Hypersonic Turbojet Engine (Sea Level Static Condition).

Length 2.7 m Width, Height 0.23 m Rotation Speed 80000 rpm Air Flow Rate 1.0 kg/s Compressor Inlet Total Pressure 100 kPa Compressor Exit Total Pressure 600 kPa Turbine Inlet Total Pressure 540 kPa Main Burner Temperature 1223 K After Burner temperature 2073 K


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Table 2. Materials for Components.

Component Part Material Air Intake Intake Ramp, Cowl Aluminum Alloy Pre-Cooler Heat Exchanger Stainless Steel

Casing of Heat Exchanger Aluminum Alloy Core Engine Compressor Titanium Alloy

Combustor, Turbine Nickel Alloy Casing of Core Engine Stainless Steel

After-Burner Exhaust Nozzle

Nozzle Cooling Wall Nickel Alloy Nozzle Plug, External Wall

Stainless Steel

Nozzle Cowl C/C-Si impregnation

III. Experimental Setup The combustion experiment is executed in the JAXA Taiki Aerospace Proving Ground in Hokkaido, Japan.

Figure 5 shows the outline of experimental diagram. Liquid hydrogen is supplied to the experimental vehicle from an external container. Then, it is supplied to the engine by pressurizing the liquid hydrogen tank using helium gas. The liquid hydrogen passes through the pre-cooler and regenerative cooling channel of variable geometry nozzle. Then, the fuel is injected into the after-burner to obtain high temperature combustion gas. Fuel rich combustion is selected in order to attain deep cooling at the pre-cooler. Helium gas is supplied for both pressurization and purge of the fuel. The helium gas is supplied from external gas container because the required amount at the ground test exceeds the capacity of installed gas cylinder.

Hydrogen gas for core engine operation is supplied both from external gas containers and installed gas cylinders. The starting sequence of the core engine is established using gas hydrogen. On the other hand, when the core engine is started with the liquid hydrogen, the control of the flow rate is difficult because the fuel can be vaporized by the thermal capacity of tubing and the control valve. Then, the starting experiment using liquid hydrogen fuel is also conducted.

PreCooler

MainBurner

LH2Tank

Compressor Turbine

Air Intake

GH2 Cylinder

GH2 Cylinder

AfterBurner

Variable Geometry

Nozzle

Evaporator

GHe

Control Valve

Shut Off Valve Shut Off Valve

Shut Off Valve

Control Valve

Regulator

LH2 GH2Balloon-based Operation Vehicle

Figure 5. Outline of Experimental Diagram.


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IV. Experimental Methodology Typical experimental conditions are shown in Table 4. In the experimental code BG, control methodology of the

core engine by using gas hydrogen (GH2) is established. The mechanical rotation speed is limited below 75% (60000rpm) because of the outer shell vibration. In the experimental code BGP, gas hydrogen is supplied to the core engine, and liquid hydrogen (LH2) is supplied to the pre-cooler and the after-burner. In the experimental code BL, liquid hydrogen is supplied to the core engine.

Table 5 shows typical experimental cases for the data evaluation in the present study. Case 1 is an experiment to operate only the core engine by using gas hydrogen, and the mechanical rotation

speed of the case reached 76%. Main burner fuel flow rate is controlled so that main burner temperature (Tmb) should not exceed the designed value.

In Case2, gas hydrogen is supplied as the main burner fuel of the core engine. Liquid hydrogen is supplied to the pre-cooler and the after-burner for 20 seconds while the core engine is operating by the mechanical rotational speed about 60%. The main combustion temperature is controlled at a constant value before pre-cooling. Then, the fuel flow control valve for the core engine is set constant during the pre-cooling operation. The corrected rotation speed reaches 75% because the compressor inlet temperature decreases.

In Case 3, main burner temperature (Tmb) is controlled all the time even after the pre-cooling operation is started. In this experiment, the flow rate of the main burner fuel increased automatically by the fuel control during the pre-cooling operation. The mechanical rotation speed reaches 69%, and the corrected rotation speed reaches 91%. Because the outer shell vibration level reaches the limiting value on the acceleration, the experiment is stopped before the end of sequence. Then, the time of pre-cooling and the after-burning is 10 seconds.

Table 4. Experimental Conditions.

Experimental Code BG BGP BL Main Burner Fuel GH2 GH2 LH2 After Burner Fuel - LH2 - Number of Experiments 10 5 4

Table 5. Experimental Cases for data analysis.

Case Case 1 Case 2 Case 3

Code Number BG50-4 BGP50-4 BGP50-5 Main Burner Fuel GH2 GH2 GH2 After Burner Fuel - LH2 LH2 After Burning Time 0 s 20 s 10 s Pre-Cooling Off On On Max. N 76 % 58 % 69 % Max. Nc 76 % 75 % 91 % Main Burner Fuel Control Tmb Control Tmb Control /

Constant RateTmb Control

Table 6 shows the experimental sequence of Case 2. The exit torch burner to burn exhausted hydrogen gas is

operated. Then, the oil supply for the bearing lubrication starts. Afterwards, an automatic sequence is started from -25s, and the core engine is rotated to 20% rotation speed by an electric motor. The main combustion starts from 0 s, and the rotation is accelerated. The engine enters the self-sustaining condition with exceeding 30% rotation speed. Pressurizing of the liquid hydrogen tank is started at 20 s. Then, the main fuel flow rate is increased while controlling so that the main burner temperature should not exceed the design temperature. It reaches 60% rotation speed at 50 s. The liquid hydrogen is supplied to the pre-cooler and the after-burner during the time. The height of the movable nozzle throat is reduced at 70 s. The supply of the main fuel is stopped at 85 s. The automatic sequence is stopped at 90 s.


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Table 6. Experimental Sequence (Case 2).

Time Event Preparation Exit Torch Burner Start

Oil Supply Start Data Acquisition Start

-25 s Automatic Sequence Start -14 s Starter Motor Start

-2 s Igniter Start -1 s 20% Rotation (16000 rpm) 0 s Main Burner Ignition (GH2) 8 s Igniter Stop

30% Rotation (24000rpm) Starter Motor Stop

20 s LH2 Tank Pressurize Start 50 s Pre-Cooling Start (LH2)

After Burner Ignition (LH2) 70 s Exhaust Nozzle Throat Height Change (from 26 to 19 mm) 80 s Exhaust Nozzle Throat Height Change (from 19 to 26 mm)

LH2 Tank Pressure Release After Burning / Pre-Cooling Stop

85 s Main Burner Fuel Stop 90 s Automatic Sequence Stop

Termination Line Purge / Bearing Cooling Data Acquisition Stop Exit Torch Burner Stop Oil Supply Stop

V. Experimental Results Stable engine operation is confirmed under the test condition of Table 5. In Case 1, the engine operation that accelerated in the restriction of the design main burner temperature is proven

through the adjustment of the fuel control variable in the preliminary experiment. The increase of the outer shell vibration is avoided by limiting the mechanical rotation speed at about 75%.

In Case 2 and Case 3, the liquid hydrogen tank and the gas hydrogen gas cylinder installed in the experimental vehicle supplied the fuel to the hypersonic turbojet, and normal engine operation is confirmed. The nozzle cowl made of C/C composite material is not damaged after the after-burning operation. The flow rate of afterburner fuel decreased after 20 seconds from ignition time because the liquid hydrogen is nearly empty in the tank at that time. Effective data of the change of operating point by movable nozzle is not acquired because the nozzle is driven after the fuel is decreased.

Starting sequence of the core engine with liquid hydrogen is confirmed by using a small diameter fuel flow valve. In the experimental code BL, the effective diameter of the fuel control valve is reduced to start the core engine using liquid hydrogen based on the results of previous experiments. A fine regulation that allows safe engine acceleration without exceeding design limit of turbine inlet temperature is attained. However, the time history of the fuel flow rate changed when the initial temperature condition of the tubing and the fuel control valve is changed. Therefore, when the core engine is operated with the liquid hydrogen, it is necessary to construct the fuel control rule that considers the temperature distribution of the fuel control valve and the fuel tubing.

Hereafter, acquired data in Case 2 with 20 seconds pre-cooling and after-burning is evaluated in detail.


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A. Rotation Speed Figure 6 shows the time history of rotation speed in Case 1. Mechanical rotation speed (N) is 20% and 16000

rpm at 0 s, 30% and 24000 rpm at 20 s, 75% and 60000 rpm at 55 s. The turbine power is relatively low in the operation below 20000 rpm because the turbine inlet pressure is not enough. Then, the rate of increase of the rotational speed is low at such rotation speed. The main burner temperature tends to rise because the air mass flow does not increase so much even if the main fuel flow rate is increased at the rotation speed.

However, the operation that exceeded the design temperature of the turbine is allowed for several seconds before it entered the self-sustaining condition because it did not enter the condition if the main fuel flow rate is not enough. Corrected rotation speed (Nc) using the average temperature of the compressor inlet shows almost the same time history as the mechanical rotation speed.

Figure 7 shows the time history of rotation speed in Case 2. Mechanical rotation speed (N) of 20% is attained by an electric motor before the main burner fuel injection. The engine is accelerated by the power of main burning and the mechanical rotation speed becomes about 47000 rpm at 50 s. The mechanical rotation speed becomes 42000 rpm at 70 s, 40000 rpm at 80 s, 45000 rpm at 85 s. This speed drop around 50 s is caused by the equivalence ratio decrease. Air flow rate increases by pre-cooling because the density increases. On the other hand, fuel flow rate is nearly constant because the fuel control valve is set constant. Then, the main burner temperature decreases and mechanical rotation speed decreases. Corrected rotation speed (Nc) after pre-cooling rises and it reaches 60000 rpm. This is because the temperature at the compressor inlet largely decreases by pre-cooling.

At 70 s, the supply of liquid hydrogen is decreased by lack of residual liquid hydrogen in the fuel tank. At the same time, the nozzle throat height is changed by the movable ramp. The former affects increase of rotation speed because the main burner temperature is raised by the decrease of air flow rate. The latter affects decrease of rotation speed because of the pressure rise of turbine exit. In this case, the effect of the latter seems larger because the rotation speed decreased. At 85 s, the nozzle throat height is 26 mm and as same as that at 70 s. However, the rotation speed at 85 s is higher than that at 70 s because of the higher main burner temperature.

Corrected rotation speed (Nc) after pre-cooling rises despite the decrease of mechanical rotation speed. The rise is occurred by the decrease of the compressor inlet temperature. The corrected speed is 60000 rpm at 70 s, 49000 rpm at 80 s, and 50000rpm at 85s, respectively. The effect of pre-cooling almost becomes the maximum at 70 s. The compressor inlet temperature rises gradually since 70 s because the mass flow rate of the liquid hydrogen in the pre-cooler decreases. Difference between the corrected rotation speed and the mechanical rotation speed becomes small, after 70 s.

Figure 6. Rotation Speed (Case 1). Figure 7. Rotation Speed (Case 2).

B. Air Flow Rate and Fuel Flow Rate Figure 8 shows the time history of the air mass flow rate in Case 2. Actual air mass flow rate (Ga) is calculated

based on pressure and temperature data that had been obtained with the pitot tubes and the thermo-couples installed in the compressor inlet. The measurement data with the pitot tubes is validated in the former experiment with the measurement value of a V-corn type air flow meter, which is connected in the series. However, the validation of the


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measuring error when there is a temperature distortion is not executed. The actual air mass flow rate rises largely after pre-cooling, and it reaches about 2 times larger than that before pre-cooling.

After the pre-cooling, corrected compressor air flow rate (Gac), which is corrected at the average compressor inlet temperature, reaches a value that is smaller than the actual air flow rate. This is because the compressor inlet temperature decreases by pre-cooling, and the volume flow rate decreased by the density rise. The time history of the corrected compressor air flow rate shows a qualitative tendency similar to the history of the corrected speed.

Figure 9 shows the time history of main burner fuel flow rate (Gfmb) in Case 2. Positive flow rate is calculated by using the upstream pressure of the choke orifice that is inserted in the supply line. The flow rate is calculated even before fuel injection at 0 s. This is occurred by the measurement noise when the measurement value is very small. The main burner fuel flow rate increases from 0 s to 20 s, and the increasing rate become smaller after 20 s. The rate change happens by the fuel control to keep the burning temperature below the design temperature, after the engine become self-sustaining operation.

After about 40 s, the experiment is conducted with the constant opening area of the fuel control valve. However, the main burner fuel flow rate changes slightly. This may occurred by temperature change of main burner fuel, which absorbs the heat of combustion gas at the downstream of the turbine by an evaporator. The evaporator is installed for the gasification of liquid hydrogen fuel.

Figure 8. Compressor Air Flow Rate (Case 2). Figure 9. Main Burner Fuel Flow Rate (Case 2).

C. Pre-cooling Temperatures Figure 10 shows the time history of coolant temperatures in the pre-cooler in Case 2. Coolant inlet temperature

(Tpci) becomes the liquid hydrogen temperature, 5 seconds after the supply of liquid hydrogen at 50 s. K type thermo-couple is used in the experiment. Then, it becomes below the usual range of the measurement in such as vicinity of 20K that is the temperature of the liquid hydrogen. Coolant outlet temperature (Tpco) becomes around 130K and almost stable.

Figure 11 shows the time history of compressor inlet temperatures (Tci-1,Tci-2) in Case 2. The compressor inlet temperature is measured at two places, the upper side and the lower side of the duct. There is a difference in the measurement value of the upper side and the lower side though both of the compressor inlet temperatures decrease after pre-cooling. There is a possibility of the liquefaction of a part of air because of the air temperature decreases to the vicinity of liquefaction temperature. Mean value of the two points (Tci-mean) is used as a reference temperature when the corrected rotation speed and the corrected compressor air flow rate is calculated. However, because the temperature measurement points are only two, this temperature has a possibility that it is different from the actual average temperature.


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Figure 10. Coolant Temperature (Case 2). Figure 11. Compressor Inlet Temperature (Case 2).

D. Combustion Temperature Figure 12 shows the time history of combustion temperature in Case 2. Main burner temperature (Tmb) exceeds

design temperature (1223K) for a few seconds in the vicinity of 20 s, at which the engine shifts to the self-sustaining operation. During pre-cooling from 50 s to 70 s, air mass flow is increased and the combustion temperature is reduced due to the equivalence ratio decrease.

After-burner temperature (Tab) is lower than the main burner temperature before pre-cooling start at 50 s. This is because there is a total temperature drop at the turbine. Moreover, it is thought that a value is lower than the actual one, because the temperature is measured by B type thermo-couple inserted in the gas flow downstream of the nozzle. The after burner temperature is from 1900 to 2000K, which is close to the design temperature, during the pre-cooling operation from 50 s to 70 s. After 72 s, the temperature decreases suddenly. Judging from the movie of the experiment, the flame in the afterburner is blown off by the influence of decreasing after burner fuel flow rate or the change in the nozzle throat area.

Figure 13 shows the time history of the total equivalence ratio in Case 2. Before 50 s, the equivalence ratio is calculated from the main burner fuel flow rate and air flow rate. It is qualitatively consistent with the historical trend of the main burner temperature. Equivalence ratio of the main burner is from 0.2 to 0.6, and it is lean combustion. During pre-cooling and after-burning after 50 s, the equivalence ratio is calculated from the main burner fuel flow rate, after-burner fuel flow rate and air flow rate.

This engine adopts fuel rich combustion in order to cool the air sufficiently and to enable the engine operation at Mach 5 flight condition. Then, the hydrogen fuel is excessively supplied comparing to the theoretical mixture ratio. The equivalence ratio becomes about from 2.0 to 2.3 in Case 2, and normal after-burning is confirmed on the condition. After 70 s, equivalence ratio decreases because the flow rate of the liquid hydrogen supplied by the liquid hydrogen tank decreases.

Figure 12. Burner Temperatures (Case 2). Figure 13. Total Equivalence Ratio (Case 2).


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E. Pressure Ratio and Performance Maps Figure 14 shows the time history of the pressure ratio in Case 2. Both compressor pressure ratio (πc) and turbine

pressure ratio (πt) change corresponding to the change in the corrected rotation speed. Each pressure ratio increases during the pre-cooling operation. Engine pressure ratio is estimated as 1.2 before pre-cooling and 1.4 after pre-cooling. The ratio is obtained by the ratio of compressor pressure ratio and turbine pressure ratio. Judging from the pressure ratio, the exhaust nozzle seems not to be choked in the experiment.

Figure 14. Pressure Ratio (Case 2).

Figure 15 shows the compressor operating line of Case 1 to 3 plotted in the performance map obtained by CFD

analyses. Both pressure ratio and the corrected mass flow rate at 75% operating point of Case 1 reaches the smaller value compared with the equivalent part on the performance map with 75% rotation. The symptom of surge is not seen from the pressure vibration measured with a high frequency pressure sensor installed on the wall surface of the compressor inlet and outlet. However, the operating line entered the surge area on the performance map. Considering from these facts, there is a possibility that both actual pressure ratio and actual corrected mass flow rate are smaller than the predicted values on the performance map with CFD.

In Case 2, the corrected rotation speed after pre-cooling becomes 75% as same as the maximum speed in Case 1. However, the corrected air flow rate and pressure ratio is lower than those in Case 1. There is a possibility to calculate higher corrected rotation speed than the actual value, because the predicted average temperature of the compressor inlet is lower than the actual value. In addition, there is a possibility that the compressor efficiency is decreased by a strong temperature distribution at the inlet of the compressor.

In Case 3, the corrected rotation speed after pre-cooling becomes 90%. However, the corrected air flow rate and pressure ratio is much lower than the values on the performance map. The reason for this difference seems as same as that in Case 2. In this case, the experiment is terminated because the vibration on the outer shell exceeded the regulated value in spite of the mechanical rotation speed of 70%. Because the operating line on the high rotation side of Case 3 comes close to the surge side in the performance map, there is a possibility that the compressor started entering to surge region.

Figure 16 shows the turbine operating line of Case 1 to 3 plotted in the performance map obtained by CFD analysis. In all cases, experimental values are almost identical to operate the turbine. Then, it is deduced that the pre-cooling and after burning has less effect on the turbine performance. However, the corrected mass flow rate of the experimental value becomes smaller in the choke area, at which pressure ratio is over 1.8, compared with the performance map predicted with CFD. The decrease may be occurred by decreasing the effective flow area in the turbine blades by separation or secondary flow.


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Figure 15. Compressor Map and Operating Line. Figure 16. Turbine Map and Operating Line.

VI. Conclusion

A hypersonic turbojet engine is installed on a flight test vehicle, and a ground firing test is conducted. Then, the following results are obtained.

- Corrected rotation speed, corrected air flow rate and pressure ratio of compressor rises by pre-cooling using liquid hydrogen.

- Corrected air flow rate and pressure ratio of compressor at the operating point with pre-cooling operation is lower than those without pre-cooing on the same corrected rotation speed. There is a possibility that the temperature distortion affects on this difference.

- Fuel rich combustion for 20 seconds and 2000K is demonstrated as after-burning during pre-cooling operation.

- The liquid hydrogen can be used for the core engine operation by selecting appropriate effective diameter of the fuel control valve.

Acknowledgements The authors acknowledge the cooperation of related researchers in JAXA. The authors also acknowledge

collaboration of professors and students in the University of Tokyo, Waseda University, Science University of Tokyo, Gunma University, and Muroran Institute of Technology.

References 1 Marshall, L. A., et. al., "Overview With Results and Lessons Learned of the X-43A Mach 10 Flight," AIAA 13th

International Space Planes and Hypersonic Systems and Technologies Conference, AIAA 2005-3336, May 2005. 2 Taguchi, H., Murakami, A., Sato, T. and Tsuchiya, T., Conceptual Study on Hypersonic Airplanes Using Pre-Cooled

Turbojet Engine, AIAA-2008-2503, 15th International Space Planes and Hypersonic Systems and Technologies Conference, 2008.

3 JAXA Long Term Vision -JAXA2025-, 2005 . 4 Taguchi, H., Futamura, H., Yanagi, R. and Maita, M., “Analytical Study of Pre-Cooled Turbojet Engine for TSTO

Spaceplane,” AIAA/NAL-NASDA-ISAS 10th International Space Planes and Hypersonic Systems and Technologies Conference, AIAA 2001-1838, 2001.

5 Fujita, K., et. al., “Precooled Turbojet Engine Flight Experiment using Balloon-based Operation Vehicle,” IAC-05-C4.5.01, 2005.

6 Taguchi, H., Sato. T., Kobayashi, H., Kojima, T., Okai, K. and Fujita, K.: Design Study on a Small Pre-Cooled Turbojet Engine for Flight Experiments,13th AIAA/CIRA International Space Planes and Hypersonic System and Technologies Conference, AIAA 2005-3419, 2005.

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[American Institute of Aeronautics and Astronautics 16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference - Bremen, Germany ()] 16th AIAA/DLR/DGLR