Shinichiro Tokudome

Highlights • Strategy formulated under Inter-University Research Institute System unique to Japan. • To develop a competitive space transportation system by using advanced technologies. • To “start small” for fulfilling our target to achieve innovations in human society. • “Fusion of transportation with exploration” in deep space exploration missions. • “Small flying test bed system” to conduct flying technology demonstrations. Abstract The Space Transportation System Committee of the Institute of Space and Astronautical Science (ISAS) of the Japan Aerospace Exploration Agency (JAXA) has been continuously formulating medium-to long-term strategies in the field of space transportation systems under the Inter-University Research Institute System of ISAS since FY2018. This committee is considering the role of ISAS in cooperation with the organization-wide activities of JAXA to formulate strategies in the field of space transportation systems. Among its previous achievements, the committee assembled a strategic target and scenario for the space transportation system research field at the end of FY2018 and has been continuously revising it. Based on the formulated mission scenario, the committee identified three priority areas related to system technologies that must be tackled. These are a “reusable orbit transportation system” aimed for highly frequent mass transportation from Earth to low Earth orbits, “deep space interorbital transportation system” aimed for a marked improvement in space science and exploration missions in terms of frequency and flexibility, and “small flying test bed system” for flight demonstrations, which is indispensable in the research and development of space transportation systems. In this paper, the authors summarize the medium-to long-term strategies and their concrete implementation measures over the next two decades. Keywords Strategic R&D, Space transportation system, Deep space exploration, Small flying test bed, Start small

The Space Transportation Committee of the Institute of Space and Astronautical Science (ISAS) of the Japan Aerospace Exploration Agency (JAXA) has been continuously formulating medium-to long-term strategies in the field of space transportation system (STS) under the Inter-university Research Institute System since FY2018. And the mission related to the strategic research and development in the STS field at ISAS is summarized as “To build an interorbital transportation network system meeting the future diverse needs of the space science field”. One of the most important technical challenges toward achieving the mission is to realize the orbit STSs capable of frequent flight and one of its keys is the hypersonic flight technology. In this paper, the authors introduce a conceptual study of hypersonic flying test bed vehicles to acquire the key technology.

In order to use laser ignition systems for solid rocket motors operating in deep space environments, it is necessary to elucidate the laser ignition characteristics of the ignition charge in low-temperature environments. This study aims to design an experimental system that can confirm the ignition threshold, ignition delay, and ignition temperature by irradiating an ignition charge with a diode laser in a low-temperature environment. Ignition experiments at room temperature were conducted. The data were evaluated statistically to obtain an ignition threshold with the maximum likelihood method. The relationship between the laser irradiation duration and the laser power with respect to the ignition threshold was obtained. The target value of the low-temperature environment temperature was determined as -50 °C. We examined the requirements of the experimental system and conceptually designed the system to simulate the low-temperature environment. It was confirmed that the constructed experimental system cooled the ignition charge to -50 °C. In the ignition experiment, the ignition charge was successfully ignited at room temperature and at low temperature. The ignition delay, the ignition temperature, and the high-speed image were obtained. Eventually, the validity of the experimental system was confirmed through the function tests.

DESTINY+ requires a high-performance kick stage with a high mass ratio and high launch system safety. The kick stage requires a reliable laser ignition system, which is the subject of this study. Our development efforts included componentizing and dual redundancy elements, such as capacitors and laser diodes in the laser firing unit, which is designated LUUS. LUUS has two types of laser-initiated devices for motor ignition and separation device operation. An optical fiber path also enables a continuity check by an optical frequency domain reflectometry device. Further, we conducted continuity checks and laser ignition tests to validate the design in procedures simulating assembly- and launch-site operations. These tests successfully executed inspections and ignitions as planned. These results confirm the validity of our design, bringing us closer to realizing a reliable laser ignition system for launch operations applicable to DESTINY+. This contribution paves the way for future advances in high-performance, safe launch systems for solid rockets in space exploration.

Laser-initiated ignition systems (LIISs) have been developed with the aim of providing essential immunity to electrical disturbances. In the basic configurations of such systems, the electric circuits for generating the laser and detonator signals are electrically separated by a non-conductive optical fiber, increasing the resistance to ignition stimuli other than laser light on the detonator side. In this study, the various environmental resistance tests required for the detonators currently used in rockets were conducted for a new laser-initiated detonator (LID). As all of the tests show results satisfying the requirements, it is considered that the LID has reached the stage of practical use.

The S-520-RD1 was developed as a test-bed small launcher using the guidance and control system based on SS-520-5 for a cube-sat launcher and successfully launched a supersonic Combustion Tester on July 24, 2022. One of the key points of this launch success is the Rhumb-line control system to achieve the guidance target and the payload test conditions. This paper shows the outline of the guidance and control system and their development, and the flight results.

To reduce the cost of space transportation, JAXA continues studying fully reusable space transporters with hypersonic airbreathing engines. To fulfill the requirement of acceleration capability from zero to hypersonic speed, the rocket-based combined cycle (RBCC) engine is one of the promising candidates and has been studied in JAXA Kakuda Space Center. In the researches of the engine, the specialized wind tunnel for the airbreathing engines (Ramjet Engine Test Facility, RJTF) has been always a practical tool to investigate the performance and detailed flow phenomena inside the engine. However, to simulate the hypersonic, high enthalpy flow, the wind tunnel uses gas hydrogen and oxygen to heat the flow, resulting the air contamination by water which affects to the combustion phenomenon. To clarify the effect of the contamination, series of the cooperative researches have been conducted in JAXA and universities. The final goal of the researches are to create the adjustment tools for contamination effects, which ensure more accurate estimation for the performances of engine in the real flight. For validating the tools, the real flight data are quite essential. In the researches, a flight experiment had been also planned to get combustion data during hypersonic flight to compare the results in the wind tunnel and estimated values by the tools. In this report, the flight experiment with “S-520-RD1” is described. In early phase of the development, several rocket systems, flight trajectories and flight test bed configurations were explored. For the experiment, the attitude control of the flight test bed was required and it was achieved by the rhumb line device. The flight conditions inwhich actual combustion experiment was conducted were determined by the air data system specifically designed and integrated into the test bed. The duration time for the combustion test reached as long as 5.8 sec, with highest Mach No. of 5.8. The summery of the results of combustion test is also presented.

The ammonium dinitramide-based ionic liquid propellant (ADN-based ILP), which is a mixture of ADN, monomethylamine nitrate (MMAN), and urea, is a low toxic monopropellant with a higher performance than that of hydrazine. To clarify the combustion wave structure of ADN-based ILP, which has low volatility, we focused on the relationship between the phase state and temperature in ADN-based ILP combustion and on clarifying the gas-liquid phase reaction. The combustion still image and temperature distribution of ADN-based ILP were obtained by strand burning tests with a high-speed camera. As a result, two stages of the stable temperature region were found in the gas-liquid phase. The pressure dependences of temperature in the stable temperature region were compared with the vapor pressure curves of some chemical substances and with the decomposition temperatures of ADN, MMAN, and urea. Then, it was inferred that the thermal decompositions of ADN, MMAN, and urea, as well as the evaporation of urea had occurred at the first stage of the stable temperature region. Also, it was found that the liquid ammonium nitrate had been dissociated at the second stage of the stable temperature region. For a report on the existence of dissociation products of MMAN and urea vapor on the burning surface at 1.2 MPa, the dissociation of MMAN and evaporation of urea would occur at the first stage of the stable temperature region at 1.2 MPa. As stated above, the combustion wave structure of ADN-based ILP was developed at 1.2 MPa.

The Space Transportation System Committee of the Institute of Space and Astronautical Science of JAXA (ISAS/JAXA) continuously draws up the Medium- to long-term strategy for the research field of space transportation system in ISAS/JAXA since FY2018. The committee is also considering about an ISAS’ role in JAXA in cooperation with a JAXAs organization-wide activity to formulate strategy for the space transportation system field in JAXA. The strategy planning in ISAS/JAXA was begun with the three documents as the point of departure; ”Long-term Vision for Space Transportation System” established by the space policy committee of the Cabinet Office, ”ISAS’ Missions”, and ”Strategic Scenario over the Next Medium- to Long-term Programs for the research field of Space and Astronautical Science” drawn up by the ISAS. As current achievements, the committee have drawn up a strategic target and scenario for the space transportation system research field at the end of last fiscal year and is continuously revising it. Based on the scenario formulated, the committee identified three major research field to be tackled. They are ”reusable orbit transportation system” becoming the key to the activation of space development and utilization, ”inter-orbit and interplanetary transportation systems” effectively supporting deep space explorations, and ”small flying test bed system” promoting the advancement of space transportation system technology. The authors introduce the summary of the medium- to long-term strategy and some medium-term research activities toward the realization of it.

This paper describes the first revised version of the medium / long-term strategic scenario for the R & D in the ISAS’ space transportation system field. The strategic scenario has been studied by the Space Transportation System Committee of ISAS since FY2018. The committee had drawn up the third edition of the scenario by the middle of FY2021.

This paper describes the development status of a laser ignition system for a solid rocket motor. This system is being developed as a simple, lightweight, and small design with a high resistance to electrical disturbances and a high level of safety. The most notable advantage of this system is that its high level of safety can decrease the cost of launching rockets into space. A laser initiator and a laser safe-and-arm device (laser S/A), which are essential components of the proposed system, were developed. In particular, prototypes of the laser initiator and laser S/A for the ignition of an upper stage rocket motor were manufactured, and some environmental tests, which are required for space rocket devices, were conducted. In addition, the lowest laser energy that is needed to ignite the laser initiator was determined by changing the laser power and operating time of the laser S/A. Furthermore, a small rocket motor vacuum fire test was successfully conducted.

The Japan Aerospace Exploration Agency, in partnership with academia and industry, are developing the Air Turbo Rocket for Innovative Unmanned Mission (ATRIUM) engine: an air turboramjet + rocket combine cycle propulsion system intended to replace conventional liquid rocket engines in Vertical Takeoff Vertical Landing applications, such as reusable sounding rockets. A subscale Flight Test Bed (FTB) vehicle is also being developed to demonstrate the ATRIUM engine in a flight environment. In this paper, the ATRIUM engine and FTB vehicle are introduced, and current progress in their development is summarized. Future test plans and practical applications are also discussed.

<p>A new space transportation system with an expendable solid-fuel booster and a reusable liquid-fuel orbiter is under consideration as part of activities in JAXA to construct a fully reusable space transportation system in the future. This paper shows this new system's conceptual study results, the system specifications, the new technology to be applied, the requirements to the subsystems, and the prospects.</p>

Copyright © 2016 by the International Astronautical Federation. All rights reserved. Although reusable launch vehicle's necessity and significance, being cost-effective, eco-friendly and reliable, have been recognized in a long time, practical system still has never been realized except the Space Shuttle. There are two main reasons in this. One reason is that reusable vehicle's recurring cost is high. The other reason is that reusable vehicle, especially that upper stage, have the problem of aerodynamic heating during re-entry. We are considering new upper stage reusable launch vehicle with solid rocket booster, which clear these problems concerning reusable launch vehicle. For the first problem on the recurring cost, the application of the auto inspection system which is cultivated in solid rocket motor's development and launch operation is being considered. That is expected to reduce the inspection cost drastically after the vehicle flight. For the second problem on the re-entry, challenging technologies are applied in the upper stage. Those are material and structure with heat tolerance and lightness, active-cooling system to share the hydrogen with the liquid propulsion system, advanced guidance and control system, and so on. On the other hand, to the lower stage or booster, application of solid rocket is considered. Since the challenging upper stage's size is expected to vary through the iteration of design cycles, the lower stage should be stable and flexible with the thrust level in development phase. Because solid motors of various sizes are developed in JAXA/ISAS since the first small solid motor started to be developed in 1954, those development method has been efficiently accumulated. Then this legacy's utilization for the new system is expected to be quite beneficial. On these technological backgrounds, this paper describes the system study for new upper-stage reusable launch vehicle with the solid rocket booster.

<p>As a replacement for hydrazine, ammonium-dinitramide-based ionic liquid propellant (ADN-based ILP) has been developed by JAXA and Carlit Holdings Co., Ltd. This propellant is made by mixing three solid powers: ADN, monomethylamine nitrate, and urea. The propellant's theoretical specific impulse is 1.2 times higher than that of hydrazine, and its density is 1.5 times higher at a certain composition. Although ionic liquids were believed to be non-flammable for a long time owing to their low-volatility, recently combustible ILs have been reported. The combustion mechanism of ILs is not yet understood. The objective of this paper is to understand the combustion wave structure of ADN-based ILP. The temperature distribution of the combustion wave in a strand burner test shows a region of constant temperature. This region would indicate boiling in a gas-liquid phase. Thus, the combustion wave structure consists of liquid, gas-liquid, and gas phases. The dependence of boiling point on pressure would identify chemical substances in the gas-liquid phase. The dependence of combustion and ignition characteristics on ADN content is also discussed. </p>

Subscale model tests and a numerical investigation are performed to predict and attenuate the acoustic level of the Epsilon launch vehicle at liftoff. Requirements for the subscale model test to predict a full-scale acoustic environment are investigated, and then the scale size of the model test is set at 1/42. The launch pads employed herein are designed to attenuate reflection of the Mach wave radiated from the free jet and the acoustic wave due to the jet's impinging on the flame deflector. When comparing the acoustic result taken at the fairing location with that at three different altitudes, 9D(sub e),14D(sub e), and 22D(sub e), where D(sub e)represents the nozzle exit diameter, the acoustic level observed at 14D(sub e)is the highest and exceeds the design requirement at all frequencies. It is revealed from the numerical analysis that the reflection of Mach waves radiating from the free jet within the flame path is the dominant acoustic source. A front cover is attached to the vertical flame path to shield the Mach wave radiation, and 3 dB attenuation in the overall sound pressure level is achieved. The design methods and the knowledge obtained in this study are valid for the design of the launch pad to attenuate liftoff acoustics.

The development of the Epsilon launch vehicle, Japan's next generation solid rocket launcher, has just moved to the final stretch for its first launch scheduled in the summer of 2013 to carry the planetary telescope satellite SPRINT-A. The JAXA appreciates the advantages of combined benefits of the standardized small satellites and the Epsilon's highly efficient launch system in order to increase the level of space activities. The primary purpose of Epsilon is to provide small satellites with a responsive launch that means "Small, Low cost, Fast and Reliable". The attention should be directed toward the innovative design concept of Epsilon, which aims at developing the next generation technologies such as the highly intelligent autonomous checkout system and the mobile launch control. Now that the full-scale development is about to be finished, the most important is what the next step should be beyond the Epsilon. This paper deals with the significance of the Epsilon launch vehicle and how it contributes to the possible evolution of future space transportation systems.

A new small solid launcher named Epsilon is currently under development in JAXA. The Epsilon launch vehicle is normally three stage rocket system and can be added an optional liquid propulsion system to the third stage for the missions requiring precision orbit insertion. The SRB-A motor boosting the H-IIA vehicle and the H-IIB vehicle will be shared as the first stage motor. Upper-stage motors are inherited from the fifth M-V launch vehicle, from the viewpoints of development cost reduction, performance increase, and advanced technology succession. The solid motor side jet (SMSJ) system, which is used for the roll control during the first stage powered flight and the three-axis control after the SRB-A burnout, will be newly developed based on the technology of the SMSJ for the M-V vehicle. A maiden flight of the first Epsilon is scheduled in the summer of 2013. A successive concept of the advanced propulsion technologies for next-gen Epsilon are also described in the present paper. There are many technical challenges, such as new propellants and mass reduction of nozzle liner, to be tackled with for the next couple of year. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

The Epsilon rocket, formerly called Advanced Solid Rocket (ASR) launcher, proceeded to the full development phase in August 2010 and its launch site was officially declared to be the Uchinoura Space Center (USC), the home of Japanese solid propellant rocket. The primary purpose of Epsilon is to provide small satellites with a responsive launch that means a low cost, user-friendly and ultimately efficient launch system. The slogan is "Small, Cheap, Fast and Reliable". This outcome is also a result of the excellent endeavors of those who devoted themselves to the next generation solid propellant rocket. However, this is not the final destination. Now that the development was approved, the most important is what the next step should be beyond Epsilon. This paper deals with the significance of the development of Epsilon launch vehicle and how it contributes to the possible evolution of future space transportation systems.

A micro solid propellant thruster for simple attitude control of a 10 kg class small spacecraft is currently under development. The prototype has 0.8 mm micro rocket elements, arrayed at a pitch of 1.2 mm on a 22 × 22 mm substance. Initially, solid propellants were used, obtaining only 20% ignition probability with a very long ignition delay (e.g. 1000 ms) as well as very high ignition energy. While observing the tested prototype sample, it became apparent that the main cause of these problems was a gap between the solid propellants and ignition heater. So, the thruster system was improved so that the propellants adhered to the heater. In addition, an ignition charge was used that starts to burn at 210°C, which was acetone having the form of slurry. As a result, a better ignition probability of over 80% in vacuum and half thrust of expectation were gained.

A nitrous oxide (N2O)/ ethanol propulsion system is distinguished as the liquid propulsion with non-toxic, user-friendly, and storable bipropellant. The current target of the present study is to build a quickresponse and maneuverable main engine of a sounding-rocket like flying test bed which will be applied to the hypersonic air-breathing propulsion researches in the near future. The application to the spacecraft propulsion is also considered due to its compatibility in low-temperature operation environment. Two series of static firing tests were performed with 700 N class gas generator models so far. Current test results showed that valuable design data were collected and operational procedure was verified. Potential of application of composite materials to the combustion chamber was also examined from the chamber wall heat flux data obtained and the result of firing test using a thick SFRP chamber.

Nitrous Oxide (N2O) / ethanol propulsion system is distinguished as the liquid propulsion with non-toxic, user-friendly, and storable bipropellant. The current target of the present study is to build a quick-response and maneuverable main engine of a sounding-rocket like flying test bed which will be applied to the hypersonic air-breathing propulsion researches in the near future. The application to the spacecraft propulsion is also considered due to its compatibility in low-temperature operation environment. Two series of static firing tests were performed with 700 N class gas generator models so far. Current test results showed that valuable design data were collected and operational procedure was verified. Potential of application of composite materials to the combustion chamber was also examined from the chamber wall heat flux data obtained and the result of firing test using a thick SFRP chamber.

Modified combined cycle engine models were tested in subsonic and transonic regions to improve the aerodynamic performance and overcome the problems which were found in the previous tests with the ejector jet prototype model. Three inlet geometries and two cowl options were tested and all models successfully reduced the inlet drag comparing to that of the prototype. In a high subsonic or transonic °ow, the current models showed the possibility of the thrust augmentation by the ejector-jet cycle. One dimensional aerodynamic estimations were also carried out and the results were compared with the data evaluated from the pitot pressure measurements. The mass °ow was in good agreement and, with a slight adjustment of the parameter, the total pressure prediction method showed reasonable match with the experimental data.

A fully reusable rocket vehicle is proposed to demonstrate good operability characteristics both on the ground and in flight. The proposed vehicle is to be used as a sounding rocket and has the capabilities of ballistic flight, returning to the launch site, and landing vertically making use of clustered liquid hydrogen rocket engines. Before initiating the development of this type of reusable rocket, a small test vehicle with a liquid hydrogen rocket engine was built and flight-tested. A demonstration of vertical landing and exercise of turnaround operation for repeated flights are the major objectives of the test vehicle. Three series of flight tests were performed in 1999, 2001 and 2003, and the flight test operation provided repeated flight environment and many valuable lessons were learned for designing the fully reusable rocket vehicle. © 2005 Published by Lister Science.

The erosive burning effects in a small test motor loaded with practical composite propellants have been investigated in detail by using an X-ray absorption diagnostics to measure the propellant local regression. The motor was specially designed to have two propellant slabs and was called as DSM (Double Slab Motor). Significant erosive burning was forced to occur in the motor by varying combustion pressure and mass flux in the port. A new simple correlative equation is derived, which predicts accurately not only pressure history but also local regression of burning surface. The equation is characteristic in the non-dimensionalized threshold mass flux and includes the terms to consider both effects of pressure and of motor scale.

The prototype of a solid propellant rocket array thruster for simple attitude control of a 10 kg class micro-spacecraft was completed and tested. The prototype has 10×10 φ0.8 mm solid propellant micro-rockets arrayed at a pitch of 1.2 mm on a 20×22 mm substrate. To realize such a dense array of micro-rockets, each ignition heater is powered from the backside of the thruster through an electrical feedthrough which passes along a propellant cylinder wall. Boron/potassium nitrate propellant (NAB) is used with/without lead rhodanide/potassium chlorate/nitrocellulose ignition aid (RK). Impulse thrust was measured by a pendulum method in air. Ignition required electric power of at least 3–4 W with RK and 4–6 W without RK. Measured impulse thrusts were from 2×10⁻⁵ Ns to 3×10⁻⁴ Ns after the calculation of compensation for air dumping.

Although a number of erosive burning correlations of solid propellants have been proposed, there is no decisive one. Therefore, a simple correlation such as Dickinson's one is preferred from practical point of view. In this study, another simple correlation expressed in terms of the ratio of mass flow rate to mass burning rate is examined based on the experiments using sub-scale double slab motors. As a result, the present correlation is shown to have the advantage that the threshold value beyond which erosive burning occurs can be treated as a constant value.

Restricting partially and temporally the initial propellant burning surfaces with thin HTPB inhibitors realized a new method of controlling the chamber pressure rise rate at ignition. The merits of the method are to be manufactured easily, to hardly affect the whole motor thrust pattern, and low in cost. The ground firing tests with sub-scale motors have verified the method to be reliable and flexible. The principal concept of a "staged ignition system" is to replace the ignitor main charge by an "ignition cavity" perforated at the fwd.-end of the motor propellant grain. This 'ignition cavity' generates the combustion gas required for igniting the remaining portion of the grain. It has been employed in the designs of two ISAS' solid motors now under development and has contributed to simplification of the ignition systems.

Combustion products flowing in the free port of a solid-rocket motor are responsible for the erosive burning effect. To improve the knowledge of the mechanism governing erosive burning, a nonintrusive X-ray diagnostic system is utilized with the objective of acquiring burning rate data with high temporal and spatial resolutions. The motor has a rectangular cross section, and it is loaded with two solid-propellant slabs parallel to each other. The X-ray source is placed above the motor. The X-ray beam passes through the aluminum windows, which are inserted in the top side of the motor case, and the component transmitted through the slabs is collected by three sensors inserted in the case on the opposite side of the windows. By means of an experimentally determined calibration curve of the attenuation of the X-ray beam through various thicknesses of an unburned propellant, the local time history of the propellant thickness can be deduced and the erosive burning can be evaluated. Results demonstrate a diagnostic reliability and provide data for improving model development and validation of the erosive burning characteristics of the tested solid propellant.

地上静止状態から音速程度までの低速飛翔環境で作動可能なエジェクタ式空気吸い込みロケットにおいては、混合ダクト出口で超音速混合流を形成することが推力増強効果の点でより有利とされている。しかし、未だその作動特性に関する一般性のある理論や経験則は確立されていない。本研究では、低亜音速までの飛翔環境も勘案した上で、超音速混合流を形成するエジェクタ式空気吸い込みロケットの地上静止状態における作動特性を明らかにすることを主目的として、固体ロケットと円筒ダクトを組み合わせて作動特性実験を行うとともに、内部流評価ために新しい解析法を導入して実験結果を詳細に検討している。そして、広い条件にわたる作動性能の予測法を提案し、基本的設計指針を示している。 第1章は序論で、空気吸い込み式ロケットに関する従来の研究動向と問題点を総括し、本研究の目的と意義を述べている。 第2章では、実験装置と方法を説明している。本研究のために試作された小型固体ロケットモータのノズル出口部を円筒型混合ダクト入口に挿入し、混合ダクト出口背圧が、吸い込まれる二次流側総圧(Pts)以下の地上静止状態で実験を行っている。実験供試体としては、ノズルスロート面積(Atp)を一定として開口比の異なる三種類のロケットノズル、および断面積(Ad)の異なる二種類の円筒型混合ダクトを長さを変えて用いている。また一回の実験で広い主流側総圧範囲の計測データを取得するため、ロケットモータ燃焼室圧力(主流側総圧:Ptp)を連続的に降下させている。取得される測定データは、ロケット燃焼室圧力、大気圧、ダクト壁圧およびダクト壁面近傍におけるガス温度である。 第3章では、安定した超音速混合流を形成する条件において、「二次流閉塞モード」と「亜音速混合モード」の二つの作動形態があることを確認している。二次流閉塞モードは従来から知られている作動形態である。一方亜音速混合モードは、二次流が亜音速のまま混合して、ある程度混合が進行した位置で流れが閉塞するもので、従来報告例のない、著者が見い出した作動形態である。本研究では、広い条件にわたって亜音速混合モードとなったことを指摘している。 第4章では、総圧比(Ptp/Pts)、混合ダクト断面積比(Ad/Atp)およびロケットノズル開口比(Aep/Atp)が作動性能に及ぼす影響を系統的に把握するために、内部流の一次元解析法を提案している。この解析法では、混合ダクト出口位置において一様な混合流が形成されていると仮定している。また新たに、主流と二次流が未混合のまま静圧の等しい流れ(Compound-compressible flow以下CCFと略す)を形成すると仮定した、静圧平衡位置が想定されている。CCFの振舞いについては、Compound-flow indicator を導入して判別できることを指摘している。 第5章では、超音速混合流の形成条件に関する解析結果より求めた超音速混合流維持に必要な最小総圧比が、実験値と一致することを示している。また、圧力測定データを用い、対応する混合ダクト出口位置状態と静圧平衡位置状態を推算している。混合ダクト出口位置静圧P2の解析結果は壁圧測定値に良く一致し、また静圧平衡位置の解析結果は、実験で確認された現象をある程度捉えている。特に、静圧平衡位置における二次流側マッハ数M1sの解析結果から、各作動形態に対応する次のような特徴を得ている。すなわち、二次流閉塞モードではPtp/Ptsの値に依らずM1s＝1.1であり、亜音速混合モードではPtp/Pts、Ad/AtpおよびAep/Atpの値に依らずM1s0.52である。二次流閉塞モードについては、本研究の解析法でM1s＝1を仮定することにより,作動性能を予測できることを示している。亜音速混合モードについても、経験的に得られるM1sの値を仮定することにより作動性能予測が可能なこと、またM1sを一定とみなすと、流量率比で代表される作動性能は、Ptp/PtsとAd/Atpの比をCとするとき、ほぼ一義的にCのみに依存することを指摘している。また、Cが一定の条件下では、混合ダクト内部に形成される流れ場は、大まかに混合ダクト直径Dを基準とした寸法で発達すると推察し、実験結果によってこれを実証している。さらに、推力増強効果に関する考察結果に基づいて、より大推力・高比推力を達成する超音速混合流形成型エジェクタ式空気吸い込みロケットの基本的な設計指針を示している。 第6章は結論で、本研究の成果が要約されている。補遺には本文で使用される数式等の詳細が記されている。 以上要するに本論文は、超音速混合流を形成するエジェクタ式空気吸い込みロケットの地上静止状態における作動特性実験ならびに混合ダクト内部流の一次元解析を行い、同型ロケットに関して作動特性の新しい予測法と基本的設計指針を与えるものであり、航空宇宙工学上寄与するところが大きい。よって本論文は博士(工学)の学位請求論文として合格と認められる。

The SAS’ nozzle two-phase flow computer program has recently been revised to achieve a higher .accuracy in calculating thrust characteristic and behavior of particle streamlines. Alumina particles had been gathered in lots of motor firings over a wide range in scale. The particle size distribution characteristics were obtained and the resulted particle sizing equation in terms of the mass averaged particle diameter D43 is quite close to that employed in the improved SPP, i.e., D43 is a unique function of motor throat diameter. Subscale motors containing highly aluminized propellant were fired, whose nozzles were designed to have significant amount of particle impingement and then severe erosion at the exit lip. The nozzle eroded locations were used to verify the computational accuracy of impingement location prediction. It has been shown that the improved two-phase flow program can well predict the particle impingement location by assuming an appropriate particle diameter DpIMP, which is proportional to D43. The proportional coefficient is rather universal.