小林 弘明

現在JAXA では,高々度気球からの自由落下を利用した微小重力実験装置の開発が進められている.この装置の特徴は,独特の二重殻構造を持つ点にあり,気球からの自由落下中に,内側の実験部を機体内部で浮遊させることで,30 秒から60 秒の良質な微小重力環境が得られる.落下中の機体姿勢制御,および実験部と機体内壁の隙間制御用に,合計16 基の50N 級コールドガスジェットスラスタが搭載されている.本稿では,微小重力実験装置用に開発されたガスジェットスラスタの設計と,その地上性能試験結果,飛行試験による実証結果について示す.

To provide long duration and good quality of micro-gravity environment with moderate cost, we proposed and have been developed an experiment system that is released from a high altitude balloon. The experiment system has a double-shell drag-free structure and it is controlled not to collide with the inner shell to realize good quality of micro-gravity environment. This paper shows the configuration of the experiment system and summarizes its five-year development including three flight test results. The fist stage of the development was successfully completed this year. The next step is micro-gravity fall with engine for longer duration of experiment. Another direction of the development is real operation of the system for micro-gravity scientists. Those future plans are also described.

Rocket-shaped vehicle is developed to conduct microgravity experiment by dropping from the high-altitude balloon. Its design strategy and development status is introduced. Also, the result of its 2nd flight test is summarized to show the feasibility of the balloon-based microgravity experiment.

ドラッグフリー技術に基づいた新しい自由落下カプセルを,2006年5月に高高度気球B200を用いて高度40km から投下し,微小重力実験が行われた。今回の最初の試験飛行により,三陸大気球観測所の制御室とカプセルとの間での無線通信,ドラッグフリー制御,そして飛行シーケンスを分析するための基本的データを得ることに成功した.

The Japan Aerospace Exploration Agency has developed a hypersonic aircraft flying at Mach 5. A precooled turbojet engine is the candidate of the engine for the hypersonic aircraft. The precooled turbojet engine has a heat exchanger(precooler) which cools the breathed air by using cryogenic propellant, such as liquid hydrogen. The precooler has a problem that frost forms on the cooling tubes of the precooler, and the frost decrease the engine performance. Some approaches to deal with the frost formation problem have employed in the development. In this paper, those approaches are introduced and the results of some fundamental studies about frost are also shown.

宇宙航空研究開発機構では,将来の宇宙輸送システムコンセプトの絞り込み,開発すべき技術課題の抽出,および現状技術の改善目標を設定するため,概念設計ツール(Systems Evaluation and Analysis Tool: SEAT)を開発中である。本報告書では,まずSEAT開発の目的,開発シナリオをまとめ,開発した雛形ツールについて報告する。ついで,代表的なエンジンの評価を行うため,雛形ルールを用いて5種類の宇宙輸送システムの概念設計を行った。本報告書では,液体ロケットエンジン,Turbine-Based Combined Cycle(TBCC)エンジン,およびRocket-Based Combined Cycle(RBCC)エンジンを対象とした。概念設計は,全備重量が最小となるように行った。その結果,母機,軌道機共にロケットエンジンを使用する二段式宇宙往還機のコンセプトが,最も軽量となる結果が得られた,本システムはエンジン搭載性についても問題がなく,最も現実的なシステムである。TBCCエンジンやRBCCエンジンといった空気吸い込み式のエンジンを使用する場合には,必要なエンジン基数が多くなり,その搭載性に問題があるという結果が得られた。

A fundamental study of frost formation around a single cold cylinder was conducted using both experimental and numerical methods. We specifically examined the mass transfer around the cylinder under conditions in which a phase change of the vapor occurs in the flow. Through the experimental study, the mass flux to the cold surface of the cylinder was measured at a constant surface temperature (200-250 K). The results show that the mass flux decreases according to the decrease of the wall temperature below 230 K, although it increases above 230 K. This phenomenon cannot be expressed using the common equation with the Sherwood number, which excludes the vapor's phase change (condensation). Numerical studies calculated the flow over the cylinder, including the vapor's phase change. The scheme for compressible, flow was modified to solve lower speed flow. Results of calculations show that we obtained the same tendency as that of the experiment: the mass flux decreases at low temperatures where the phase change occurs.

Experimental studies on telescopic aerospikes for aerodynamic control are reported in this paper. Parametric study on the aerodynamic characteristics of the aerospike has performed including the effect of spike length L, base diameter of tip cone D, spike translating speed and direction. The Axial force coefficient <I>Ca</I> of the aerospike suddenly increases at L/D=3.0 due to flow mode transition from the separation to the reattachment. Reattachment/separation flow mode transition phenomenon can be applicable to a newly invented aerodynamic control device, which is called air-breathing aerospike. In this paper, verification test results of this air-breathing aerospike are also reported. A small solenoid valve in the body cylinder successfully controls reattachment/separation flow mode transition at the angle of attacks from 0 to 12 degree. The spiked bodies’ <I>Ca</I> varies according to the mode transition. As a result, we can control the aerodynamic property of the spiked body by opening/closing the valves periodically.

This paper reports an experimental study on flow oscillation characteristics of an aerodynamic control device that we have proposed. The device can achieve an enhancement of the aerodynamic control ability and a reduction of the flow instability by adding multiple stabilizer disks to a conventional aerospike so as to divide the flow separation region into multiple cavities. In this device, several axisymmetric cavities are formed. It is well known that pressure oscillation is induced around cavities. In this study, the characteristics of the pressure oscillation of several cavities on a cone surface were investigated experimentally by unsteady pressure measurements in a wind tunnel testing. The conical-cavity pressure oscillation had a feature that the oscillation level is large in case that a length-to-depth ratio of the cavity is large; the oscillation frequency can be predicted by the famous Rossiter formula, which is reported in many previous researches on a single rectangular cavity. It was also found that adding thin disks into the large cavity is effective in the reduction of the pressure oscillation level downstream of the cavities. In addition, disk structural vibration measurements were conducted simultaneously with the unsteady pressure measurements, revealing that a flutterlike vibration could occur when the pressure oscillation frequency agrees with the disk eigenfrequency.

This paper reports experimental studies on telescopic aerospikes with multiple disks. The telescopic aerospike is useful as an aerodynamic control device; however, changing its length causes a buzz phenomenon, which many researchers have reported. The occurrence of buzzing might be critical to the vehicle because it brings about severe pressure oscillations on the surface. Disks on the shaft produce stable recirculation regions by dividing the single separation flow into several conical cavity flows. The telescopic aerospikes with stabilizer disks are useful without any length constraints. Aerodynamic characteristics of the telescopic aerospikes were investigated through a series of wind tunnel tests. Transition of recirculation/reattachment flow modes of a plain spike causes a large change in the drag coefficient. Because of this hysteresis phenomenon and the buzzing, the plain spike is unsuitable for fine aerodynamic control devices. Adding stabilizer disks is effective for the improved control of aerospikes.

In this paper, a concept of a new variable-geometry aerodynamics device, which is designated “Multiple-Row-Disk (MRD) device,” is introduced. The MRD device divides large separation region around the shaft of an aerospike into several small cavity flows with multiple disks arranged on the shaft. Experimental studies on aerodynamic characteristics of conical nose with axisymmetric cavities were conducted in order to evaluate a feasibility and a fundamental characteristics of the MRD device. It was found that the MRD device could improve not only drag characteristics compared to the conventional aerospikes, but also static longitudinal stability characteristics compared to the conical nose.

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.

This paper describes design study of the gas turbine engine using PDE as a main combustor of which hydrogen is used as a fuel. In the present analysis, exhaust gas condition of the simple PDE in which no turbine or other obstacles are installed which is estimated by CFD is referred as a turbine inlet condition. About 65 % of enthalpy drop at the exit of the PDE occurs during the exhaust of high temperature combustion gas behind the detonation wave. Furthermore, turbine adiabatic efficiency map of a conventional turbine is taken into account. Approximately, 51-80% of enthalpy drop through the turbine can be extracted as a turbine power. Partial filling of the PDE tube has an effect of smoothing the peak of U/C0. Copyright © 2007 by the American Institute of Aeronautics and Astronautics Inc. All rights reserved.

This paper proposes a new micro gravity experimental system called BOV (Balloon-based Operation Vehicle). BOV uses a free-fall capsule with double-shell structure to prevent influence of aerodynamic disturbance. Additionally, BOV is raised to 40km by a high altitude balloon to extend micro gravity duration to 30(or possibly 60) seconds. Thus we realize a medium duration micro gravity system with good micro gravity environment. In this system, the most characteristic point is double-shell structure. The inner shell can fall freely since the outer shell measures the relative position with laser displacement sensors and is controlled by gas-jet thrusters not to collide the inner shell. Therefore the inner shell can be uninfluenced of the dynamic pressure and other aerodynamic disturbances ideally. The BOVs project has run since 2004. The first flight to check the whole system was accomplished in 2006. The aim of this flight was test of a high altitude balloon, communication and data handling system, control system, onboard electronics and operation. The second flight expected to achieve 30 seconds micro gravity was also accomplished on May in 2007. This paper presents the development of BOV's control system and shows the experimental results of micro gravity and consideration for effectiveness of the proposed system. Copyright 2007 by the IAF or the IAA. All rights reserved.

In this paper, experimental studies on telescopic aerospikes with multiple disks are reported. The telescopic aerosoike is useful as an aerodynamic control device; however, changing its length causes a buzz phenomenon, which many researchers have reported. The occurrence of buzzing might be critical to, the vehicle because it brings about severe pressure-oscillations on the surface. Disks on the shaft produce stable recirculation regions by dividing the single separation flow into several conical cavity flows. Therefore, the telescopic aerospikes with stabilizer disks are useful without being any length constraints. Aerodynamic characteristics of the telescopic aerospikes were investigated using wind tunnel tests. Transition of recirculation/reattachment flow modes of a plain spike causes a large change in the drag coefficient. Because of this hysteresis phenomenon, the plain spike is unsuitable for fine aerodynamic control devices. Adding stabilizer disks is effective for the improved control of aerospikes.

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. © 2006 Elsevier Ltd. All rights reserved.

A study on mass flux of water vapor and mist around a cold cylinder (120–250K) was conducted by means of both experimental and numerical methods. The cylinder was placed in a forced convection air flow at a speed of 1m/sec. The experimental study revealed that the mass flux of the cylinder decreases rapidly under the temperature of about 200K. The mass flux at the cylinder temperature of 120K is one-sixth of that of 240K. The numerical study could simulate the mass flux around the cylinder in which we used a new phase change model with considering the transfer of the particles of mist. By this calculation we found some characteristics of the mass transfer of the cold cylinder which is unique when the temperature of the cylinder becomes cold enough to occur condensation.

A fundamental study for frost formation around a single cold cylinder was conducted using an experimental and numerical method. We focused on the mass transfer around the cylinder under the condition where phase change of the vapor in the flow occurs. By the experimental study, the mass transfer rate on the cold surface of the cylinder at a constant surface temperature (200–250K) was measured. The results show that the mass transfer rate decreases according to the decrease of the wall temperature below 230K, while it increases above 230K. This phenomenon can not be expressed by the common equation of Sherwood number in which the phase change of the vapor (condensation) is excluded. In the numerical study, we calculated the flow around the cylinder including the phase change of the vapor. The scheme for compressible flow was modified to be able to solve lower speed flow. As a result of the calculation we obtain same tendency as that of the experiment that the mass flux decreases at low temperatures where the phase change occurs.

Aerodynamic performances of a rectangular intake were investigated experimentally. After a tradeoff study of rectangular intakes whose operative Mach number is from 0 to 6, 20% external compression intake is selected as the best intake from the viewpoint of low number of actuators. Intake performances such as total pressure recovery and mass flow ratio are evaluated by wind tunnel tests. The free stream Mach number of the wind tunnel was M5.1. The size of the intake was 75mm in cowl capture height. Low ramp driving force was achieved by connecting links of the second ramp and third ramp. After the first wind tunnel test that is performed to evaluate the basic performance of the intake, the configuration of the intake is modified. Ramp length of the first ramp and the second ramp were changed to improve the total pressure recovery. Bleed from the second ramp is added. Seal mechanism between the variable ramps and the sidewall is modified. Total pressure recovery is improved from 9.9% to 21.7% by the modifications.

A new reprocessing technology, FLUOREX was proposed for thermal reactors cycle and future thermal/fast reactors (coexistence) cycle. The proposed system is a hybrid system that combines fluoride volatility and solvent extraction methods. Spent fuel will be sheared and cladding material will be removed by dry oxidation/reduction method such as AIROX process. Fluorination and purification of most uranium can be easily achieved by fluoride volatility method with compact facility. About 10% residues including plutonium can be treated in well-established PUREX method, which means this facility load will be about 1/10 of the conventional PUREX facility with same capacity. Between fluorination process and PUREX process, there is a pyrohydrolysis process where the fluoride compounds from fluorination process are converted to the oxides. Pure mixture of Pu and U can be obtained by solvent extraction method without separating Pu and U, which is suitable for conventional MOX fuel fabrication. The system can recover pure U and MOX with the decontamination factor of over 107 and can drastically reduce the cost and waste generation compared with the conventional one. Semi engineering scale experiments for the fluorination, pyrohydrolysis, and dissolution of Pu containing materials were carried out. From those experimental results, key elemental processes were fundamentally proofed. © 2005 Published by Elsevier Ltd. All rights reserved.

パルスデトネーションエンジンを航空宇宙用空気吸い込み式推進機関へ利用することを想定し，空気吸い込み式ＰＤＥの推力性能を見積もった．ＰＤＥは，供給圧が外気圧（大気圧）まで膨張するため，高空で作動する場合，推重比が低下する．この推力低下を軽減するため，新概念ＰＤＥを提案した．ＰＤＥの出口に高速開閉弁を設け，燃料充填時に弁を閉じることにより，高速飛行時のラム圧力をＰＤＥの推力に有効に利用する．さらに，インテーク捕獲流を燃料を利用して予冷却することにより，エンジン飛行可能領域がM0.5上昇する．これらのシステム解析の結果を踏まえ，PDRJE(Pulse Detonation Ramjet Engine)を提案した．さらに，ＰＤＥの基本的な推進性能について調査することを目的として，水素・酸素ＰＤＥの燃焼実験を実施した．ＰＤＥ燃焼器に作用する熱流束は5kW/m²/Hzであった．

In this paper, a new concept concerning the effective use of conical cavity flow is presented. Because the flow over cavities may cause noise, oscillation, and resonance vibration, most of the researches focus on the negative effect of the cavity flow however, there are also some studies concerning the effective use of cavities. For example, spike at nose of body is used as an aerodynamic device for drag reduction and/or thermal protection system of the vehicle. In 2003, we proposed an advanced aerodynamic device with conical cavities. This device, which is designated multi-rowdisk (MRD), has a cone and some disks that are placed in the axial direction. Because flows over deep cavities formed between disks are stable, this structure can replace the solid structure without deteriorating aerodynamic performance of streamline objects. We continue the research on the MRD structure for applying to supersonic inlets or airplanes. In the first part of this paper, examination on the conical cavity flow is reported. In the second part, we show wind tunnel results of the supersonic inlet with the MRD structure. © 2005 Published by Lister Science.

Six Sigma is the management strategy developed by Motorola to reduce defects in products. Design for Six Sigma (DFSS) is a methodology for determining the values of the design parameters, which maximize the performance of some system without tightening the material, manufacturing or environmental tolerances. This paper presents the implementation of DFSS for redesign of the LE-7 engine. Uncertainties with design parameters and operational conditions are considered in evaluating thrust performance, thrust chamber life, turbo-pump cavitation, and combustion stability. Traditional deterministic optimization results and probabilistic optimization results are compared. It is found that robustness of rocket engine is not always consistent with the extension of thrust chamber life.

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 had 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. © 2003 Elsevier Ltd. All rights reserved.

The way to demonstrate the ATREX engine in

A nozzle of the ATREX Engine must have variable geometry to change pressure ratio from 3 (SLS) to 550 (Mach 6). An axisymmetric plug nozzle will be employed for the ATREX Engine. Throat area of the axisymmetric plug nozzle can be controlled by moving plug back and forth in the axial direction. The plug nozzle shows relatively higher performance than conventional C-D nozzle under any ambient pressure. The objective of this study is to estimate thrust efficiency and boat tail drag of the plug nozzle. Several types of the plug nozzle were tested in a supersonic wind tunnel. Injecting secondary flow thorough the cowl was devised and shown to be effective in the reduction of boat tail drag.

Axisymmetric and three dimensional CFD simulations were performed to predict and improve internal and external performance of supersonic inlet of ATREX engine. Aerodynamic performances such as total pressure recovery, mass capture ratio and intake external drag were in good agreement with the experimental results over a Mach number range from 1.5 to 5.5 and over a angle of attack from 0 deg to 4 deg. However, because the CFD result estimates the flow separation caused by the subsonic diffuser, the static pressure of the CFD result was lower than the experimental data. Furthermore, some ideas to improve the performances are shown, that is to increase the bleed flow rate of the spike and decrease the area gradient of the subsonic diffuser.

This paper describes the study of frost formation on the precooler tube surfaces. Frosting rates are quantitatively shown by experimental and numerical methods, which depends on the temperature and humidity of the air flow as well as the cooling wall temperature. The effectiveness of a method to improve the precooler performance under frosting condition was investigated by experiments using a sub-scale heat exchanger model. Addition of a methanol proved to be most effective compared with other possible substances in both cases of lower and higher cooling wall temperature. Then the effectiveness of the methanol addition was ascertained for the practical condition that means the same tube configuration and flow velocity as the precooler designed for the ATREX engine firing test model. The result showed that the addition of the same quantity as the water vapor could restrain the frost layer from choking the flow in the duration of 300 seconds, which is sufficient time for precooler operation. The required methanol mass along the ATREX engine flight path was estimated to be less than 3 % of fuel hydrogen on board. Accordingly, the method came to be promising candidate for practical application.

Here is presented an experimental and analytical study on a precooler for hypersonic air-breathing engines. Precooling of the incoming air breathed by an air-inlet gives extension of the flight envelope and improvement of the thrust and specific impulse. Three precooler models were installed into an air-turbo ramjet engine and tested under the sea level static condition. When the fan inlet temperature was down to 160K, the engine thrust and specific impulse increased by 2.6 and 1.3 times respectively. parametric studies on the precooler design values and a sizing analysis were also performed, Decrease of tube outer diameter on the precooler is only way to increase heat exchange rates without increase of its weight and pressure loss.

The present paper addresses the development study of the air turbo ramjet engine with expander cycle (ATREX) being conducted since 1986 in the Institute of Space and Astronautical Science in cooperation with the industries (IHI, KHI, MHI and SHI). The ATREX is expected as one of the most promising candidates for the propulsion system of a future space plane. The ATREX is the combined cycle engine performing like a turbojet at lower flight speed and a fan-boosted ramjet at higher flight speed beyond Mach 3 to 6. The 1/4-scaled model of ATREX whose fan inlet diameter of 300mm was built for system verification under sea level static conditions. Carious components such as turbo-fan, precooler, mixer, regeneratively cooled combustor and heat exchanger as well as the expander cycle have been developed and verified in the firing test. Number of 63 tests with 3,300 sec of the total duration have been conducted step by step since 1990 at Noshiro Testing Center of ISAS. The recent study focused on the precooler. The latest model (Type-III) was designed taking into consideration the reduction of size and weight as well as the heat-dynamic performance aiming at the flight model/ Engine performance is improved such that the Thrust and specific impulse are increased 1.8 and 1.2 times respectively by reducing the fan inlet temperature to 180K. The heat exchange performance of precooler was 80-90 % of the design value and the large pressure losses of the air flow occurred due to the frost formation on the tube surfaces. This frost formation phenomenon has been made clear analytically and experimentally and several methods eliminating it have been devised and tested.

The flight of spaceplane is always under accelarating in the assent way and always under decelerating in the desent way and yet cruising in the return way. Besides, its flight envelope is considerably wider than that of airplane. Thus the integrated design method is required to build the best transportation system optimized taking into account the propulsion system and the airframe under the entire flight conditions. In this paper its shown an optimization method on TSTO spaceplane system. Genetic algorithm (GA) was applied to optimize design parameters of engine, airframe, and trajectory simultaneously. Several types of engine were quantitatively compared using payload ratio as an evaluating function. From a Viewpoint of the relation between performance and weight, it was concluded that the precooled turbojet is the most promising engine for TSTO among Turbine Based Combined Cycle (TBCC) engines.

A flight-type ATREX engine consists of many components; inlet, precooler, fan, turbine, combustor, nozzle, and so on. These components have been developed by a number of sea-level combustion tests and wind tunnel tests in ISAS. In this paper, a detailed characteristic model of each component is described. These characteristic models are integrated into a flight-type ATREX engine simulator.

In this paper, specification for TSTO spaceplane powered by airbreathing engines is presented. A Multi-criteria Trade-off Analysis (MTA) of Turbine Based Combined Cycle (TBCC) engines was made to estimate the performance of an flight-type engine under the current level of technology. The cycle analysis has been performed after all the component models are reviewed. The flight-type ATREX engine can meet all demands from the TSTO spaceplane but Isp. The following calculation results showed the possibility of ATREX engine providing higher Isp by improving the performance of some components.

The ATREX engine has been developed by ISAS for the propulsion system of a fly-back booster of a future TSTO space plane. Summary of the engine system, current R&D status and plans on the ATREX are presented in this paper, Several studies on system optimization, precooled expander cycle, thermal-fluid performance, control and composite materials have been performed by the system firing tests and wind tunnel tests, etc. As the next step, we propose a development plan in the next decade, in which a half-scale prototype engine with the flight weight will be produced for the demonstration flight test.

The flight of Spaceplane is always under accelarating in the assent way and always under decelarating in the desent way and yet cruising in the return way. Besides, its flight envelope is considerably wider than that of airplane. Thus the integrated design method is required to build the best transportation system optimized taking into account the propulsion system and the airframe under the entire flight conditions. In this paper it is shown an optimization method on TSTO spaceplane system. Genetic algorithm (GA) was applied to optimize design parameters of engine, airframe, and trajectory simultaneously. Several types of engine were quantitatively compared using payload ratio as an evaluating function. It was concluded that precooled turbojets is the most promising engine for TSTO among Turbine Based Combined Cycle (TBCC) engines.

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 modern 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. © 2002 International Astronautical Federation. Published by Elsevier Science Ltd. All rights reserved.

Here is presented an experimental and analytical study on a precooler for hypersonic air-breathing engines. Precooling of the incoming air breathed by an air-inlet gives extension of the flight envelope and improvement of the thrust and specific impulse. Three precooler models were installed into an air-turbo ramjet engine and tested under the sea level static condition. When the fan inlet temperature was down to 180K, the engine thrust and specific impulse increased by 2.0 and 1.2 times respectively. Thick frost formed on the tube surfaces at the entrance part of the precooler blocked the air-flow passage. On the other hand, very thin frost formed at the exit part because the water vapor included in the air was changed to mist particles due to the low temperature of the air in this part. Parametric studies on the precooler design values and a sizing analysis were also performed. Decrease of tube outer diameters on the precooler is only way to increase heat exchange rates without increase of its weight and pressure loss.