To realize high-performance cryogenic propulsion systems, the chilldown sequence has to be improved. Because the chilldown is carried out under low gravity, the effect of gravity on the two-phase flow, especially at low flow rate, should be investigated. To understand the physics under low gravity, an experiment was conducted using a sounding rocket. Two identical test sections with different mass flow rates simulated part of a turbopump, each of which has a complex flowpath including slits and a dead end. Using liquid nitrogen, the flight experiment obtained data of temperatures, pressures, void fractions, and video frames of liquid motion. Then, the flight experiment data were compared to the ground data taken under normal gravity, revealing that the slits played an important role in the chilldown process and that the test sections were quickly chilled down under low gravity. The slits of the test sections formed liquid jets, and their behaviors were different from those in the ground experiment. In the flight experiment, the jets easily reached the dead end of the test sections and cooled down the whole walls due to the increase in inertia and wettability; however, such behaviors were hardly observed in the ground experiment. The difference between the ground and flight is significant at lower flow rate.

This manuscript presents the design of a capacitive void fraction sensor for cryogenic LN2 two phase flows and its validation at room conditions. The capacitive void fraction sensor is first designed by means of Electric Field Analysis (EFA) simulations taking into account specific technical constraints coming from the test section in which it should be accommodated. Then it is manufactured and validated using a proper combination of fluids (Polydimethylsiloxane (PDMS) and air) having a dielectric constant ratio similar to the one encountered in LN2/GN2 two phase flows. The validation is performed through comparison with void fraction measured by means of optical visualizations and shows how the capacitive measurement technique robustness allows obtaining reasonable accurate values of void fraction also for the substitute fluid case. The sensor presented in this manuscript was used to evaluate the void fraction during LN2 chilldown of a rectangular cooling channel and the results are presented in the second part of this work.

The hypersonic air-breathing engine, which is currently under development by Japan Aerospace Exploration Agency (JAXA), uses liquid hydrogen as the fuel. In order to accurately control the fuel flow rate during the start-up, it is essential to measure the heat transfer and pressure drop of the two phase flow. These two characteristics depend on void fraction, flow velocity and flow regime; thus, measurement methods for these values are required to be established. In this study, the void fraction measurement method by the high-speed image analyses has been developed. Two images taken from top and side directions by high-speed cameras of 1000 fps are used for "two-direction semi-automatic analysis". We also develop "One-direction full automatic analysis" which uses only the side view as a full-automatic and high sampling rate method with little accuracy deterioration. The preliminary verification test using vertical pipe and acrylic ball shows favorable results within 2.2% error against the theoretical value. A cryogenic experiment using two-phase nitrogen flow was also conducted. Sampling rate of "One-direction full automatic analysis" can be up to 1000 Hz. The difference between the results of two methods was as minor as 5% when the void fraction was below 30%.

Capacitance type void fraction sensors for cryogenic fluid has been developed by our research group. The sensors are roughly categorized into three types; arc type, small-sized arc type, and asymmetric type. In this paper, the three type sensors are compared, and each sensor's merits and demerits are indicated. The arc type sensor was first developed. The sensor is simple, but S/N ration is low and influence of "temperature drift" is large. The small-sized arc type sensor was developed for the sounding rocket S310 test No.43, so the sensor was needed to make as small as possible. The sensor succeeded to measure the void fraction of the liquid nitrogen flow even under microgravity. The asymmetric type sensor was deviced to improve the measurement accuracy. Inaccuracy of the sensor between true void fraction and measured void fraction was only 3% for stratified flow, whereas that of the arc type sensor 30%.

The Japan Aerospace Exploration Agency launched the S-310-43 sounding rocket from the Uchinoura Space Center on Aug.04, 2014 for the purpose of investigating such behavior as boiling and flow of cryogenic liquid rocket propellant in an environment simulating coasting flight on orbit by using the sounding rocket's sub-orbital ballistic flight. In the low-gravity state, the cryogenic fluid (liquid nitrogen) was introduced into the test sections of similar shapes to the flow channels in the cryogenic propulsion systems. The boiling of liquid nitrogen inside the test-sections and the transition of flow regimes from gas/liquid two-phase flow to liquid mono-phase flow were visualized. The temperatures, pressures and void fractions of each channels were measured as well. Development of the experimental equipment for S-310-43 sounding rocket is described in this paper.

The flow around an ONERA-M6 wing, including the effect of wind tunnel wall interference, is computed using CFD analysis with a porous wall model. The computational domain sets porous walls at the top and bottom of the wing section similar to actual wind tunnel experiments. The computational result captures almost the same shock wave shape as the wind tunnel. This could not be computed exactly in previous works that did not include wall effects. The interference of the porous walls reduces the Mach number and attack angle of the flow, and these effects alter the swept angle of front shock wave and the location of rear shock wave. The aerodynamics coefficients are also affected by the wall interference. The lift coefficient becomes smaller due to the reduction in attack angle. The lower Mach number decreases the drag coefficient, while the reduction in attack angle causes additional drag by a mechanism similar to induced drag.

In this experimental study, the combined effect of spatial and temporal variations of fuel-air mixture on self-excited combustion instabilities in a gas-turbine model combustor (similar to 60 kW) with a low-swirl injector is reported. Detailed measurements were performed in 4 fuel split (to upstream/downstream injections) conditions while keeping the total equivalence ratio constant. The combustion stability was found to be very sensitive to the fuel split parameter which determined the local equivalence ratio distribution. The majority of the heat-release oscillations was generated in the flame-to-wall impingement region in a manner that satisfied the Rayleigh criterion. The driving force of the instability was considered the periodic interaction between the traveling vortex filled with fresh mixtures and the flame in the near-wall region as reported in the previous study on homogeneous mixture flame. However, the strength of the instability was sensitively modified by the change in the local equivalence ratio distribution. In the strongest oscillation case with inhomogeneous mixture, temporal variations of equivalence ratio exhibited a positive contribution to the thermoacoustic coupling. This suggested that temporal variations in equivalence ratio were enhancing the driving factor of the thermoacoustic instability in addition to the vortex-flame interaction mechanism. (C) 2014 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

Laser Induced Plasma Spectroscopy (LIPS) is a technique aiming at measuring local equivalence ratio inside gas turbine combustors without disturbing the combustion fields. In this study, LIPS is applied to well controlled mixtures of gases at the pressure of 1.5 MPa to obtain plasma spectra as calibration. Two types of estimation methods are examined. One is the peak area ratio method in which ratios between particular atomic emissions in the plasma spectra are used. The other is the correlation method where correlation coefficients between the objective and the database spectra are used. The correlation method showed the best performance in terms of precision in the calibration test. Finally, the technique is applied to the measurement of 2.5 MPa combustion gas. The peak area ratio of H656/ N746 provided a good agreement with the gas sampling data, while the peak area ratio of H656/O777 and the correlation data shifted toward higher equivalence ratios. Oxygen quenching is considered as a dominant cause of the shift. In summary, the results showed promising features of the LIPS technique under high pressure conditions. For improving the accuracy further, it is required to take into account the effect of oxygen quenching.

To model porous walls used in transonic wind tunnels, flow through a hole is investigated using Computational Fluid Dynamics (CFD). First, we analyze the relation between flow rate and differential pressure across the hole. At low differential pressures, such as for wind tunnel porous walls, the flow rate is found to increase linearly with differential pressure. We therefore propose a new model based on a linear relationship between flow rate and differential pressure. The effects of hole shape and boundary layer conditions near the hole are then investigated. In the outflow case (i.e., wind tunnel to plenum chamber), the flow rate increases as the ratio of hole depth to diameter becomes large due to variation of the flow separation area at the hole exit. Boundary layer thickness also affects the flow field: when the ratio of boundary layer thickness to hole diameter becomes small, the flow rate decreases, because the flow along wind tunnel side wall interacts more strongly with the flow through the hole.

A precooled turbojet engine has been developed by JAXA used for the hypersonic airplane and spaceplane. The subscale engine named "S-engine" whose thrust and weight are about 1.2 kN and 100 kg was designed, fabricated and tested. The components and the system firing tests under the sea-level-static condition were successfully conducted. In the next phase, a flight test of the S-engine is planned using a stratospheric balloon in 2010 called balloon-based operation vehicle (BOV). The vehicle is dropped from an altitude of 40 km by a high altitude balloon. After 40 s free-fall, the vehicle is pulled up and the S-engine operates for 30 s at about Mach 2. High-altitude tests of the core-engine verified the performance and healthiness of the engine under the condition corresponding to the BOV flight trajectory. (C) 2009 Elsevier Ltd. All rights reserved.

Precooled turbojet engines are effective propulsion systems for hypersonic aircraft. However, a serious problem is that frost forms on the cooling tubes of the precooler, thereby decreasing the engine performance. This paper presents a new method for defrosting the precooler using jet impingement. The validity of the proposed defrosting method was investigated through fundamental experiments. In the experiments, the frost formed on the cooling tubes of the single-row heat exchanger was removed by jet impingement. We used the jet periodically. The jet interval is 10-50 s. In addition, the jet duration is short (about 0.1 s). Therefore, the consumption of the high-pressure air to make the jet flow is small. The coolant temperature and main How speed influences on the defrosting method effectiveness were assessed. Results show that this defrosting method is effective, especially when the coolant temperature and the main flow speed are low.

Expansion Deflection nozzles present an attractive proposition as a replacement for conventional nozzles on launch vehicles, due to their reduced length, and altitude compensating capability. However, it has long been speculated that they suffer in the latter regard due to aspiration of the low speed flow region inside the nozzle by the supersonic jet surrounding it. This effect is investigated in this paper by direct experimental measurement of base pressures, and found to have little effect on the base pressure of the nozzle within the range of operating conditions investigated. Wall pressures were also used to calculate the efficiency of the altitude compensation within the nozzle, which was found to be between 87 and 100% for the three operating pressure ratios examined. This represents a significant improvement over conventional nozzle performance, and further conformation that wake pressures are indeed close to ambient.

This paper describes a development study of a precooled-cycle hypersonic turbojet engine for the first stage of TSTO space plane and hypersonic airplane. With reflecting the key technologies accumulated from ATREX (expander cycle ATR engine) ground tests, the next flyable subscale engine "S-engine" is now developed. S-engine has 23 cm x 23 cm of rectangular cross-section, 2.2 in of the overall length and about 100 kg of the weight employing a variable-geometry rectangular inlet and nozzle. It produces 1.2 kN of thrust at SLS, which corresponds to (1)/(4) of the ATREX engine. Design of the hypersonic components such as the inlet, precooler and nozzle has been finished and their aerodynamic performances were verified by wind tunnel tests and CFD analyses. A prototype model of the diagonal-flow compressor whose pressure ratio is 6 was manufactured. Its rotating tests under the very-low pressure conditions are now in progress. The reverse-flow annular combustion chamber was successfully tested.
The first flight test of the S-engine is to be conducted in 2008 by the balloon-based operation vehicle (BOV) which is about 5 m in length, 0.55 m in diameter and 500 kg in weight. The vehicle is dropped from an altitude of 40 km by a high altitude balloon. After 40-s free-fall, the vehicle pulls up and S-engine operates for 30s at about Mach 2. High altitude tests of the engine components corresponding to the BOV's flight condition have been conducted. (c) 2007 Elsevier Ltd. All rights reserved.

Development of the Balloon-based Operation Vehicle (BOV) is currently in progress for the first flight scheduled in 2006. In a series of BOV experiments, a vehicle in a wing-body configuration is lifted by a high-altitude balloon and dropped, after which the microgravity experiments will be performed onboard the vehicle under favor of the quasi-free-fall environments. Although the BOV is originally designed for the microgravity experiments, various types of experiments can also be performed in a hypersonic flight at lower altitudes. One candidate currently under review is a flight experiment of a precooled turbojet engine in reduced dimension. In this article, an overview of the BOV experiment is introduced, and the current development status of the BOV and a flight model of the precooled turbojet engine is presented. The aerodynamic load and the aerodynamic characteristics of the BOV are obtained by computational fluid-dynamic analyses and wind-tunnel experiments. (C) 2006 Elsevier Ltd. All rights reserved.

This paper discusses the current R&D status and plans concerning a Mach 6 turbojet engine with an air-precooling system for the first stage of a two-stage-to-orbit space plane (TSTO). An air-turbo ramjet engine with the expander-cycle (ATREX) has been designed and tested at sea level static conditions, which demonstrated the engine system and component performance. Experimental and numerical research of the components have also been conducted to build the fundamental technologies and to improve the engine performance. Some innovative ideas were proposed such as a multi-row-disk inlet and a defrosting method on precooler tubes using methanol. As the next step, the development of a subscale flight-type engine (S-engine) has started. The partial expander cycle was selected instead of the full expander cycle as the prototype engine cycle as a result of optimizing analyses. Total length and weight of S-engine are about 2.2 m and 100 kg respectively including a variable air-inlet and nozzle. Because S-engine will be tested in a flight demonstration program after ground tests, it must be designed with consideration of weight reduction as well as keeping of the high performance. The first engine flight test will be around Mach 2, using a balloon dropped test vehicle. This is scheduled in 2007 FY.

An air-precooling system before compression is indispensable to extend the flight envelope and the improvement of the performance of turbo-based air breathing engines for the space plane. One of the critical problems on a shell-and-tube-type precooler is a deterioration of its heat exchange and pressure recovery performance due to the thick frost formation on its tube surface. An innovative method is proposed to mix a condensable additive like ethanol, methanol, etc. in the airflow as a defrosting system. The defrosting effectiveness and essential factors on the additive were investigated by using a small heat exchanger under two different cooling temperature conditions, that is, lower and higher cooling wall temperatures than the melting point of mixture of the water vapor and the additive. It was cleared in the test that most of the alcohols bad good effectiveness with methanol the best. This methanol addition concept was applied in the precooler of the practical ATREX engine. A methanol injection system worked well and the thick frost layer formed on the tube surface at the entrance side of precooler could be eliminated. The required methanol mass along the ATREX engine flight path is estimated to be less than 3% of fuel hydrogen consumption. (C) 2003 Elsevier Ltd. All rights reserved.

To study the dynamic response of a supersonic airbreathing engine and establish control logic for intake unstart, restart control tests were conducted at Mach 3 using a subscale engine model consisting of an axisymmetric intake (inlet) and a turbojet. Assuming that the combustion flame is blown out by intake unstart, restart control follows a sequence. First, after a flow is started the turbojet engine is ignited. Second, the intake is started while the rotational speed and the combustion gas temperature of the core engine are controlled. Third, the intake spike position and the terminal shock position are controlled, and the intake total pressure recovery achieves the design value (60%). The tests were successful, and engine thrust was recovered at approximately 30-40 s after engine start-up. A sudden increase in combustion gas temperature and rotational speed occurred after intake unstart. To reduce the sudden increase in the gas temperature, a new sequence that involved closing a fuel control valve after detection of intake unstart was implemented, and the increase in gas temperature was reduced. To avoid intake buzz, buzz margin control using a bypass door was successfully implemented.

Feasibility studies were carried out aiming at the application of carbon/carbon (C/C) composites to a turbine disk, heat exchangers, and a plug nozzle for an engine intended for use in a future reusable space vehicle. In these applications, the maximum temperature was estimated to be about 1500degreesC. In order to withstand this high temperature, attempts were made to utilize three-dimensionally reinforced C/C composites. The most serious problem encountered in the application of C/Cs to the turbine disk was the loss of fragments of the composite located near the outer periphery due to strong centrifugal force, which resulted in severe vibration due to rotational imbalance. The heat exchangers and plug nozzle have complex shapes in order to realize a large heat exchanging area. Joined structures were explored for these components. The principal effort in these applications has been placed on finding structures requiring low joining strength and developing materials with low gas leakage.

A feasibility study of three-dimensionally fiber-reinforced carbon-carbon composites (3DC/Cs) for application to a turbine disk of ATREX (Air turbo ramjet engine with expander cycle) was carried out. Spin burst tests at room temperature were conducted using 3D-C/C disks, and the fracture behaviors were characterized. A 3D-C/C disk was totally fractured at a peripheral speed of 516 m/s (r = 150 mm), which is sufficient for the ATREX application. However, fiber bundles at the disk periphery prematurely suffered micro-scale damage, and fragments of the fiber bundle unit flew out before total fracture occurred. In order to prevent the fly-out behavior, the disk was impregnated with Si only near its periphery. Although this treatment increased the initiation speed of the fly-out behavior, this improvement was considered insufficient for purposes of the ATREX application. Next, a simplified analysis was conducted to characterize the fly-out behavior. Based on this analysis, the following three measures were discussed: (1) decreasing bundle thickness (i.e. using fine fiber texture), (2) increasing toughness of the fiber bundle interface, and (3) minimizing local curvature in waviness of the fiber bundles in the circumferential direction.

A feasibility of ATREX (Air-Turbo-Ram Expander cycle) engine with conventional aft-turbine configuration has been studied to be developed in about 10 years, if the development project has started under enough resources. The novel tip-turbine of the original ATREX engine is replaced by a conventional aft-turbine, and the maximum turbine inlet temperature (TIT) is reduced to 1200K, to realize the engine by only using approved metal technologies of modem jet engines. The capability of the performance has been shown by parametric studies by changing components' design parameters. The study shows that the performance of the ATREX engine is not less than that of pre-cooled turbo jet.
Some technical issues on developing the new ATREX engine have been addressed. The most important issue would come from the transient total temperature change due to the rapid acceleration from sea level static (SLS) condition (288K) to Mach 6 at 30km of altitude (1680K) in 6 minutes. The deformation due to transient thermal expansion has to be controlled. Especially, the change of the tip clearance and the clearance between rotors and stators are pointed out to be important design issues. The ATREX engine, which has shorter axial length and simpler rotor, has structural advantage over turbo jet. (C) 2002 International Astronautical Federation. Published by Elsevier Science Ltd. All rights reserved.

A review of the development of a precooler for the air-turboramjet expander-cycle (ATREX) engine is given. Three types of precooler for the ATREX engine ground-test model were designed, manufactured, and tested under sea-level static conditions. The results suggested two problems affecting the precooler performance, heat transfer rate and airflow pressure drop. One is nonuniformity of the airflow through the tube banks. The other problem is frost formation on the heat transfer surfaces. Concerning nonuniformity of airflow, the shell configuration was modified based on analysis by computational fluid dynamics calculation. To improve the precooler performance under frosting condition, a new method to add a condensable gas into the airflow was proposed and examined by experiments on a subscale heat exchanger model. Addition of a small quantity of ethanol can effectively restrain the decline of the precooler performance due to frost formation.

The study on ATREX engine (Air-Turbo Ramjet engine) development is being been conducted in ISAS since 1986 as a candidate fur the propulsion system of the fly-back booster up to Mach 6 on the reusable TSTO space plane. ATREX is a fan-boosted ramjet engine using liquid hydrogen as a fuel and coolant. Sea-level static firing tests of ATREX with precooling were carried out in parallel with the wind tunnel tests on the aerodynamic components such as air intake and plug nozzle. Further studies on precooler have been conducted not only for performance improvement but also for weight reduction to meet flight requirements.
In 1998, a new type precooler (Type-III) was designed and tested. Its heat transfer performance could be improved by increasing its compactness using tubes of smaller in diameter as well as its weight could be reduced. On the other hand, some difficulties increased in its manufacturing due to larger number of tubes required. The new precooler performance on heat transfer and pressure loss compared with the old type of precoolers are presented here. The test results of ATREX engine installed with new precooler is presented.
Some progress in aerodynamic studies on ATREX engine is also presented here. The control of the axisymmetric air inlet by capturing the terminal shock waves acid the reduction on boat tail drag of the plug nozzle were conducted in the wind tunnel. (C) 2001 Elsevier Science Ltd. All rights reserved.

This is the status report of the development study on ATREX engine (Air Turbo Ramjet) that is now under way in the Instutute of Space and Astronautical Science (ISAS) in 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 fly-back booster of TSTO space plane. ATREX is the fan-boosted ramjet producing the effective thrust from sea level static to flight Mach number 6. ATREX is worked on the expander cycle with precooling incoming air as shown in Fig.1. ATREX employs the tip turbine configuration which allows compactness and light weight of turbo machinery and the variable geometry airintake and plug nozzle which allows the wide range flight conditions.
ATREX development study has been conducted with the sea level static tests since 1990. ATREX engine for tests is the scaled model designated by "ATREX-500" 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 aerodynamic components such as the axisymmetric variable geometry air intakes, the precoolers and the variable geometry plug nozzles. The application study on advanced carbon-carbon composite for the tip-turbine and fan assembly has been conducted. This study status is presented in the another session of this IAF congress(1).
The flight test of ATREX is now planning to verify the engine performance and functions in the practical flight conditions by using an unmanned flying test bench.
In 1995 was tested ATREX-500 installing the precooler under the sea level static conditions to examine the engine performance and the icing problem on the precooler.
The present paper addresses primarily the test results of precooled ATREX engine. (C) 1997 International Astronautical Federation. Published by Elsevier Science Ltd.

This 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 fly-back booster of TSTO space plane. ATREX is the combined cycle (a fan-boosted ramjet) engine providing the effective thrust from sea level static to flight 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 configuration 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 flight conditions and the application studies on advanced carbon-carbon composite for the tip-turbine and fan assembly has been proceeded. In 1994 initiated the flight test plan in which ATREX will be verified in the practical flight conditions by using an unmanned flying 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 flow and the analytical studies and the designs on the precooler and the precooled ATREX engine and the flight test plan. © 1998 Published by Elsevier Science Ltd. All rights reserved.

We developed a dynamic simulator for the hypersonic precooled turbojet including the unsteady models of the inlet unstart phenomenon and flow properties of two-phase hydrogen flow. Wind tunnel tests and numerical simulation clarify that the amplitude and frequency of the inlet buzz are determined by the inflow and outflow balance of the inlet and the duct volume located after the inlet. A capacitance-type void fraction meter was newly developed and relationship between the quality and heat transfer rate of the nitrogen was researched. The dynamic simulation agrees with the engine firing test

The most important problems in supersonic combustion ram jet engines (SCRAM) which are suitable for hypersonic transports and space planes are the phenomena such as thermal choking and compound choking. In the case of integrated SCRAM engines, the boundary layer flow which is developed along the vehicle fuselage is sucked into engines with the main flow. The flow tube of the boundary layer and that of the main flow have the same total temperature but different total pressures and Mach numbers. When they coexist in the engine they behave quite queerly against the cross-sectional area change in the engine and heat addition in each flow tube. The main flow tube is easily choked at supersonic conditions because of the existence of the boundary layer flow, where the static pressure balance is kept between two flow tubes. This phenomenon is called compound choking.The purpose of the present research is to investigate the details of choking phenomena in SCRAM engines and to find out the control rule which is necessary for stable operation of SCRAM engines.In order to realize choking in SCRAM engines, an engine module of 30mm×32.6mm frontal cross section was put in the supersonic wind tunnel of 100mm×100mm cross section driven by the vitiated air of temperature 2000K (maximum) at Mach number 2.0. Self-ignition was realized by injecting hydrogen fuel into the engine module. From the pressure signals measured on the engine wall, it was recognized that supersonic combustion was realized while the injection pressure of hydrogen was low. But when the injection pressure was increased the pressure inside the engine increased indicating subsonic combustion. In such circumstances the region of high pressure was observed to extend in sequence from downstream to upstream in the engine and it finally fell into the unstart condition

Wind tunnel test methods on the air-breathing engine for the reusable space planes were researched under the actual flight conditions. A control method of the air-inlet by adjusting its shape was proposed and confirmed in the ONERA S3 supersonic wind tunnel. This test was successfully conducted in the variable Mach number conditions, which simulated the accelerating and decelerating conditions of the actual flight. An unsteady control method when the compression form changed was established. Some problems were cleared by comparing with the CFD analysis. Now the inlet control test with a small turbojet engine for a model plane in the supersonic wind tunnel is planned. The control scquences including in case of the inlet unstart and restart will be constructed in the test. As for the total engine system, the firing test of the air-turbo-ram jet engine was conducted under the sea-level static. The frost formation problem was especially investigated in the test and countermeasures against it were proposed and verified. Some lobe type mixers were tested in the ram combustor test. Platinum catalytic igniters were also tried in this combustion test, which was cooperated with ENSMA in France. We reserached some wind tunnels and measurement systems in USA, France and also discussed with foreign research on the future overall engine testing in the supersonic and hypersonic wind tunnel

A feasibility study of three-dimensionally fiber-reinforced carbon-carbon composites (3D-C/Cs) for application to a turbine disk of ATREX (Air-turbo-ramjet engine with expander cycle) was carried out. Spin burst tests at room temperature were conducted using 3D-C/C disks, and the fracture behaviors were characterized. A 3D-C/C disk was totally fractured at a peripheral speed of 516 m/s (r = 150 mm), which is sufficient for the ATREX application. However fiber-bundles at the disk periphery prematurely suffered micro-scale damage, and fragments of the fiber bundle unit flew out before total fracture occurred. In order to prevent the fly-out behavior, the disk was impregnated with Si only near its periphery. Although This treatment increased the initiation speed of the fly-out behavior, this improvement was considered insufficient for purposes of the ATREX application. Next, a simplified analysis was conducted to characterize the fly-out behavior. Based on this analysis, the following three measures were discussed: (1)decreasing bundle thickness, (i.e.,using fine fiber texture),(2)increasing toughness of the fiber bundle interface, and (3)minimizing local curvature in waviness of the fiber bundles in the circumferential direction. The third countermeasure was found to be most effective and disks minimizing local curvature were verified to satisfy ATREX requirement.In addition to above study, experiments to confirm durability of 3D-C/Cs were carried out. This category of study included, high temperature behavior, creep, fatigue, and notch insensitivity. The 3D-C/Cs were found to possess sufficient above durability for the ATREX application

ターボジェットエンジンの飛行領域を極超音速領域まで拡大し、比出力、比推力を向上させるために、取り込んだ空気を熱交換器(空気予冷却器)で冷却してファンに送り込む空気予冷却というシステムが有望である。しかし、このシステムにおいて、主流中の水分が空気予冷却器の冷却面で凝結して伝熱性能を劣化させることが問題となっている。本研究ではこの問題に対し、主流空気中に凝縮性ガスを混入することにより、霜層内部の空隙を充填し密度を増加させ、霜層厚さの成長を緩和する方法を提案し、要素モデルにより実験した。まず、数種類の凝縮性ガスを主流空気と混入することによって、混入ガスのどの様な性質が重要であるかを調査し、アルコールが効果的であり、中でもメタノールが最適であることを示した。また、着霜低減のメカニズムを解明し、物質の親水性と融点降下作用が非常に重要であること、冷却面の温度によって着霜軽減メカニズムが異なることを明示した。次に、メタノールの主流に対する適正混合比を把握するために、メタノールの混合比を変化させて実験を行い、主流中の水蒸気量に対して、圧力損失の低減という観点からは0.5倍、伝熱性能低下の防止という観点からは1.2倍のメタノールを混合すればよいことが分かった。また、主流の相変化が着霜へどのように影響するかを調べるため、当実験の冷却面温度でも相変化が起きる二酸化炭素を主流として用いて実験を行い、主流の相変化により着霜量は増加するが、これに対してもメタノールの混合は効果的であることを確認した。最後に、具体的なエンジンに搭載する方法についても検討を行い、液体水素を燃料とするスペースプレーン用ターボジェットエンジンにおいては、燃料の約3%のメタノールが必要であることを示した

Supersonic Biplane Theory is a concept of eliminating wave drag and sonic boom by using shock wave interaction between two airfoils facing each other. The following research topics are investigated in this research to construct and validate the theory : Definition of three-dimensional wing configuration and its optimization, Investigation of low speed characteristics of supersonic biplane by using CFD and EFD, Supersonic wind tunnel test of starting process of three-dimensional supersonic biplane, Free flight experiment of supersonic biplane using ballistic range

現在JAXAで研究開発が進められている静粛超音速機用インテークバズ現象について、数値流体解析による現象解明を行った。インテークバズは低流量作動領域で発生する衝撃波の自励振動で、インテーク形状や主流条件、後方の圧力条件等、様々な要因によって特性が異なり、バズの発生条件やメカニズムに関する知見は十分ではない。本研究の結果、始動状態からバズに遷移する途中に微小振動を伴う剥離状態があることを見出した。また、機体とインテークを統合した数値解析をJAXAで実施した風洞実験と比較し、バズが発生する条件が一致し、インテーク単体の場合とは条件が異なることを確認した。

地上静止状態からマッハ5の広範囲を単一機構で運用できる極低温空気予冷器（プリクーラ）を搭載した予冷ターボジェットエンジンを提案しているが，プリクーラへの着霜が問題となっている．本研究課題においては陽極酸化処理とフッ素処理により超撥水性を付与したアルミニウム製伝熱管を用いて，その着霜抑制の有効性を確認する事を目的とし，また独自の2次元着霜モデルを提案し霜層の成長を数値解析にて予測する事を目的とした．低速風洞での着霜試験では着霜質量が10～17%程度減少した一方で，撥水性が霜層密度を減少させることも分かった．数値解析では低コストで精度の良いモデルを開発した．今後は実機材質のSUS316L材へと拡張させて研究を行う．

本研究では、二段式スペースプレーン（TSTO）機に搭載が検討されているロケット複合サイクル（RBCC）エンジンのエジェクタジェットモード作動時とインテーク不始動時におけるインテーク性能と流れ場の様子をRANSによる粘性数値解析を用いて調査した。主流条件としては、離陸速度である主流Mach数0.5およびインテークが不始動状態となるMach 2.0、動圧50kPaとした。その結果、主流Mach数0.5においてエジェクタ効果により吸い込み空気量が11.9％増加することおよびランプ面静圧が5〜10%低下していることを確認した。一方、主流Mach 2.0においては、従来システム検討に使用していた簡易推算法に比べて、6.7％空気吸い込み流量が低くなった。その理由として、インテークが不始動に陥ることにより、ランプ面途中で大規模な剥離が生じていることを明らかにした。また不始動の剥離点の圧力上昇について前向きステップや過膨張ノズルと同様にArensの自由干渉理論による結果とよく一致することがわかった。

極超音速空気吸込式エンジンを用いた航空宇宙輸送機の開発をブレークスルーするため、飛行実験による機体／推進統合制御技術の実証を提案している。本特定課題研究においては、鍵技術である極超音速矩形インテークに対して、数値解析および超音速風洞試験によって、空力性能を取得することを目的とした。飛行実証機と同サイズの小型模型のMach4超音速風洞試験を実施し、従来の模型と比較することで、サイズの違いによるRe数の効果を調査した。流量捕獲率と全圧回復率の最高性能に関しては、ほぼ同じになるが、作動範囲が狭くなることがわかった。今後は、データベースの構築と飛行実験の詳細検討を実施する。

　本研究では、三次元Busemannインテークの等エントロピ圧縮による高性能という長所を活かし、始動性の悪化や長さが長くなるという短所を克服すべく、新形状のインテークを設計し、風洞実験およびCFD解析によって性能の実証を行うことを目的とした。今年度の研究成果を以下に示す。　第一に、Taylor-Maccoll方程式およびStreamline traced techniqueを用いて、性能を落とすことなく、長さを短縮できるインテーク形状を設計した。従来の入口楕円-出口矩形インテークと比べ、入口部を半楕円、半矩形型に変えることにより、二次流れを軽減させることができた。　第二に、非粘性計算によって、５種類の形状の異なるインテークを解析した。Streamline traced technique を用いて設計した切り込みを持つインテークは、設計の元となった軸対称形状と同等の性能が得られることが確認された。また、改良型である前述の楕円、半矩形型インテークは、矩形入口形状に比べ長さを16.3 %短縮でき、同じ長さの楕円入口形状に比べ、全圧損失を8.9 %減少させることができた。非設計点マッハ数においては、いずれの形状も良い性能を示した。　第三に、超音速風洞実験（JAXA相模原キャンパスの超音速風同実験装置を借用）によって、楕円、半矩形型インテークの性能を検証した。複雑形状である本インテークを光造型技術を用いて製作することによりコストを低減した。金属材料に比べて、強度が弱い部分を金属板で補強することにより、風洞実験に耐えうる強度を確保した。風洞装置の故障により、Mach 2条件のみで実験を行った。下部カウルの位置による離脱衝撃波の位置や振動の様子などの流れ場の変化を確認した。風洞実験を粘性および非粘性の数値解析と比較し、衝撃波の位置が異なることにより、最大壁面静圧で、約10%の差異を生じた。　以上、ほぼ計画通りに進み、外部講演３件、修論１編、卒論１編の成果をあげることができた。この結果を踏まえ、更なる改良設計と高マッハ数での性能評価を実施する予定である。

　本研究では、ロケットエンジンや極超音速予冷ターボジェットエンジン（PCTJ）の燃料である極低温液体水素の二相流状態における流動特性を解明するため、主要パラメタであるボイド率計を開発、高度化することを目的とし、実験的に検証した。研究成果を以下に示す。　第一に、これまで研究を進めていた静電容量型ボイド率メータ（液相と気相の誘電率の違いを利用している）に関して、課題となっていた温度ドリフトを是正する手法（配管材質の選定と電極サイズの調整）を確立し、温度ドリフトをおよそ1/10に低減することに成功した。　第二に、観測ロケットを用いた液体水素流動試験に搭載するための小型ボイド率を開発した。主な課題である、小型化、部品点数の削減（耐振動、耐加速度）、ノイズの減少、極板位置の精度向上を実施した。また、自作Cメータの回路設計、製作を実施し、十分な精度が得られることを確認した。　第三に、これまで使用していたボイド率計測の理論式について検証を行うために、電場解析を実施し、理論値、実験値と比較した。また、内部の流動様式をトモグラフィの原理を用いて、静電容量から逆投影法による解析手法を検討し、予備実験を行った。　第四に、市販流体解析ソフトopen formを用いて、二次元の気液二相流の解析を行い、実験結果を補間した。今後、自作の二相流解析スキームを構築する予定である。　第五に、本静電容量型ボイド率計を応用した、クオリティーメータを提案、試作した。クオリティを直接計測する方法は、これまで実施例がみられず、画期的なものである。　第六に、ステレオ撮影によって得られた画像に関して、光学式疑似三次元画像解析手法を確立し、光学的にボイド率を計測した。次にあげる液体窒素を用いた実験で、静電容量型ボイド率計と比較を行い、計測精度を確認し、本手法が時間分解能の高いことを実証した。　上記で構築した技術は、液体窒素を用いた流動実験（東北大学の試験装置を使用）により、確認し、問題点を抽出した。　以上、研究計画で示されていた項目はほぼ実施され、所定の成果をあげ、講演１件を行い、査読論文を投稿中である。今後は、この結果を踏まえ、小型ボイド率計、トモグラフィ手法、クオリティ計測手法を洗練し、実用化に向けた研究を実施する予定である。

　本研究では、極超音速／超音速空気吸込み式エンジンに使用するエアインテーク（空気取入口）の遷音速時における非定常特性を解明することを目的とし、超音速風洞実験によってインテークバズという非定常現象を実験的に調査した。　インテークバズとは、遷音速領域（Mach 1.3-2.0）において、インテーク表面の境界層剥離によって起こる自励振動で、エンジンの停止や構造破壊を起こす危険性があるため、エアインテークの作動領域が狭くなるという問題が生じている。そこで、JAXAで開発中の予冷ターボジェットエンジンの矩形インテークモデルを新規に設計、製作し、Mach 2-3の風洞実験によって、遷音速領域における空力特性データを取得するとともに、バズの発生しない安定作動領域を調査した。　実験の結果、本エアインテークモデルにおいて、始動状態または不始動状態からバズへと遷移する現象を観測した。このとき、性能曲線上において流量捕獲率は減少し、全圧回復率は上昇することがわかった。シュリーレン撮影と非定常圧力計によって、バズの振動周波数は20 ～ 30 Hzであることが観測され、ランプ側の流れが剥離していることより、Dailey型バズであることを確認した。また、インテークのスロート面積を小さくしていくにつれて、バズの周波数が減少し、最終的にバズが停止することを確認した。さらに、亜音速ディフューザの後ろ側にバイパスドアを設け、そこから抽気をすることによって、バズを停止させる手段を提案し、実証した。　現在は、実験と並行して、CFD解析によって風洞実験で見えなかったインテーク内部流れの様子を検証することを試みている。今後の課題としては、バズの発生タイミングを明確化すること、抽気量のバズに対する影響を調査すること、バズの周波数の支配パラメタを明らかにすることがあげられる。

　本研究では、将来の極超音速輸送機に向けてJAXAで研究開発中の極超音速予冷ターボジェットエンジンを対象とし、非定常シミュレータを構築することを目的とした。　まず、非定常的要素の強い、液体水素の相変化を含めた熱流動特性を解明し、モデル化を行うため、気液二相流に関する実験を行なった。静電容量型ボイド率計と高速度ビデオ撮影（画像判別法）によるボイド率計測という二つの手法を構築し、水-空気、軽油-空気について適用した。その結果、5%以下の精度でボイド率計測が可能であることを確認した。さらに、これらを水素を用いた予備実験で適用したところ、二相状態においては定量的にボイド率が計測されることや、超臨界時の密度変化による静電容量の変化を捉えることができた。一方、極低温による静電容量の温度ドリフトという問題点が明らかになった。また、同時に圧力損失を計測し、水素二相流におけるクオリティと摩擦係数に関する定性的な評価を行った。今後は、実験の精度を向上させ、また、伝熱特性を評価することで、二相流状態の熱流体特性のモデル化を行なう予定である。　第二に、汎用性のある非定常シミュレータの枠組みを構築し、実際の予冷ターボエンジン燃焼実験との比較を行なった。解法としては、エンジン内の各流体機器を１ないし２つの要素としてモデル化し、流体の保存方程式を解くボリュームジャンクション法を採用した。圧縮機はP-Qマップを用いてモデル化し、バルブやタービン等においてはチョーク計算を導入した。実験結果と比較したところ、定量的にもおおよそ一致し、この手法が有効であることが確かめられた。また、エンジンがウインドミル作動している状態についても、圧縮機の流体抵抗をモデル化し、組込むことによって再現することができた。　本研究によって、エンジン非定常シミュレータの完成に近づくとともに、液体水素燃料のマネジメントに関する貴重なデータを取得することができた。

　本研究では、小型の液体水素ターボエンジンの起動制御手法を確立するために、熱物性モデルを組込んだ非定常シミュレーションを実施し、実験と比較した。　対象としたエンジンは、JAXAで研究開発中の予冷ターボエンジン（Sエンジン）である。Sエンジンの燃料制御、特に主燃焼器の起動時には、無冷却タービンの温度制限値（950 ℃）を越えないように、かつ、目標回転数へと短時間で到達させなければならない。さらに、液体水素が供給されると配管やバルブの熱容量により蒸発し、流調弁において、気相、気液二相、液相と変化し、密度が急激に変動する。この急激な密度変動を制御することが、極めて難しく、これまでのエンジン試験でも運転シーケンスは確立していない。　本解析では、市販のプラント用非定常解析コード（DYNSIM; Invensys Process Systems社製）に自作の液体水素伝熱流体モデルを組込んだ。配管やバルブなどを要素に区切り、物質収支、熱収支、圧力バランス等の微分方程式をオイラー法で解くものである。各要素には、材質、形状、CV値、熱物性値、摩擦係数等を与える。　熱物性モデルに関しては、流調弁の上流側は、超臨界の単相熱伝達を仮定し、変化の激しい下流側は配管を10個のセグメントに分割し、それぞれのセグメントごとに、流体のクオリティ、伝熱面過熱度に応じて場合分けし、熱伝達率を計算した。蒸発器の出口以降は、完全に気相であるため単相強制対流熱伝達とした。なお、圧力損失に関しては、全ての領域で単相を仮定した。　実験は、JAXA大樹宇宙実験場において2008年11月に行なわれた。水素流量の時間変化を比較したところ、定性的な傾向はあっているものの、定量的には大きくずれている。ずれの要因としては、二相状態の熱伝達解析の誤差や流調弁のCV値の推測誤差が考えられる。また、シミュレーション結果によると流調弁の位置とオリフィスの位置で、非定常的に流量が一致していないことがわかった。今後、物性モデルの高精度化を実施する。

液体水素燃料を冷熱源とした熱交換器（プリクーラ）の問題である冷却管への着霜問題を解決することを目的とし、空気ジェットによって霜層を吹き飛ばす手法を提案して、実験的に実証した。過去の研究より、層の性状は冷却面温度に大きく影響を受けることが分かっており、冷却面温度をパラメタとして、霜の性状とジェットによる吹き飛び効果の関係を調査した。実験方法を以下に述べる。SUS製の熱交換器に冷媒を供給し、調温調湿空気供給装置から供給される湿り空気と熱交換させる。熱電対と差圧計により熱交換器の熱交換量と全圧損失を計測することにより、熱交換器の性能を定量的に評価した。冷媒としては代替フロン系冷媒および液体窒素を使用し、冷却管表面温度は、Tw =250, 220, 83 Kの3種類とした。空気側の速度は着霜していないときに1 m/sとし、ジェットの速度は、冷却管の前面において36 m/sとした。以下に、実験結果を示す。ジェットによる除霜をしない場合、全圧損失は冷却面温度Twが低いほど大きく、実験開始後100秒時において、Tw = 80 Kの場合は、Tw = 250 Kの場合の約10倍の全圧損失上昇となった。ジェットを間欠的（50秒に1度）に噴射し除霜を試みた結果、Tw = 80 Kの場合は、噴射した瞬間には着霜のないときとほぼ同じ圧力損失に回復した。すなわち、本手法の除霜の効果は確認された。一方、Tw = 220 Kの場合には除霜の効果は弱く、Tw = 250 Kに至っては効果は見られなかった。これは、表面温度が高い場合には、霜層の表面温度が上がり、液相となり、それが霜層内部に浸透することにより、密度が大きく固い霜層になることが原因であると予想した。しかしながら、先に述べたように、Tw = 250 Kの場合には、霜層による圧力損失自体が小さいことより、問題は小さいと考える。今後は、実際の熱交換器に近い形で定量的な調査を行う一方で、エンジンの圧縮機からの抽気をジェットとして利用するシステムを構築する予定である。