德留 真一郎

The Institute of Space and Astronautical Science (ISAS) of the Japan Aerospace Exploration Agency (JAXA) has been researching the Air Turbo Rocket for Innovative Unmanned Mission (ATRIUM) engine, which uses LOX and LH2 as propellants, since 2018. It can be used for both vertical take-off and landing (VTOL) reusable sounding rockets, such as the “New Sounding Rocket” proposed by the Institute of Space and Astronautical Science, and spaceplanes equipped with combined-cycle hypersonic air-breathing engines, as studied by the Japan Aerospace Exploration Agency. The ATRIUM engine combines an air-turbo engine with a rocket engine. It uses a gas generator (GG) cycle to deliver a high thrust-to-weight ratio suitable for VTOL, using only the air-turbo engine. The engine thrust control system includes a GG, a hot-gas flow-control valve (HGV), turbomachines, a secondary combustor (SC) that produces the resultant thrust, and a rocket engine. This system has two roles. The first role is to continuously switch between air-turbo-only mode, air-turbo-and-rocket mixed mode, and rocket-only mode as the vehicle ascends from the ground to space. The second role is to enable deep throttling using only the air-turbo engine during landing. These components, except for the rocket engine, are essential for a small-scale VTOL technology demonstrator vehicle powered by the air-turbo engine part of the ATRIUM. In this presentation, we report on progress in researching and developing the thrust control system for the ATRIUM prototype air-turbo engine for low-altitude VTOL flight demonstration.

This study aims to assess thermal-resistant and insulation systems for the secondary combustor of the Air-Turbo Rocket for Innovative Unmanned Mission (ATRIUM) engine. We try to be capable of multiple reuses by employing an "ablation cooling method." This is a passive cooling approach utilizing carbon fiber reinforced plastics (CFRP), which facilitates weight reduction and minimal load on the engine system. The heat resistance and thermal insulation characteristics of the ablator material were evaluated under the ATRIUM engine environment using a subscale test apparatus at Teikyo University. In this experiment, measurements of heat flux and total temperature were conducted in both free-jet and All-around style configurations and the ablator test was conducted by free-jet style. Based on these results, the heat transfer coefficient was derived and compared with those of the CRAFT-Phase.1 and subscale models estimated using Bartz's equation. As a result, the heating environment in the free-jet configuration is considered appropriate as a condition for evaluating the ablator. Moreover, we found the relationships of the surface temperatures and reusability by the ablator test.

This paper presents the concept, technical features, strategic significance, and current development status of a solid propellant kick stage for deep space exploration under development at the Institute of Space and Astronautical Science (ISAS), JAXA. The kick stage injects a spacecraft into a highly elliptical Earth orbit after a flagship small Japanese rocket places it into low Earth orbit. Integrated with the spacecraft as part of the payload, the kick stage accelerates the combined stack after separation from the launch vehicle and releases the spacecraft upon reaching the target orbit. By providing additional velocity after orbital insertion, the system expands the capability of this class of small launch vehicles and enables mission profiles requiring higher characteristic energy for deep space departure. The development addresses a key objective: increasing the feasibility and launch frequency of deep space missions using small launch vehicles. ISAS aims to establish a mission architecture in which its flagship small rocket supports more frequent and flexible scientific exploration. Within this framework, the kick stage functions as an enabling infrastructure technology that broadens mission opportunities. It also advances solid rocket propulsion technology and supports further development of upper-stage solid motors. The main propulsion is provided by a high-performance solid motor carrying 800 kg of propellant and achieving a mass ratio of approximately 0.91, representing a high level of transport capability in its class. The kick stage employs a decoupled architecture without electrical signal interfaces to either the launch vehicle or the spacecraft. This design minimizes integration complexity and enhances mission adaptability. A laser ignition system initiates both motor firing and separation events. By delivering ignition energy optically, the system remains inherently resistant to lightning, static electricity, and stray electrical currents, improving operational safety and simplifying ground handling. In addition, a spacecraft adapter incorporating a CFRP lattice structure manufactured via the VaRTM process reduces structural mass while maintaining the required stiffness. Although initially planned for application to the DESTINY+ mission, the kick stage is no longer slated for that mission due to a change in launch vehicle. Development nevertheless continues, recognizing its importance as an enabling technology for future small scale deep space missions. At present, detailed design of the major components has been completed, and preparations are underway for the system-level Critical Design Review (CDR). The kick stage aims to serve as a common foundational technology supporting future deep space exploration missions.

本発表では、宇宙航空研究開発機構（JAXA）で開発中の複合サイクルエンジンの技術実証を目的として設計されるロケット試験機の概念研究についてご紹介します。この技術は、高速ポイント・ツー・ポイント輸送機、ひいてはスペースプレーンシステムの実現に寄与するとともに、それらの競争力向上にもつながるものです。

The Institute of Space and Astronautical Science (ISAS) of the Japan Aerospace Exploration Agency (JAXA) conducts a deep space exploration mission named Demonstration and Experiment of Space Technology for INterplanetary voYage with Phaethon fLyby and dUst Science (DESTINY+). The mission requires a high-performance, compact solid kick stage with a high mass ratio and high system safety. The kick stage employes a newly developed laser ignition system to meet these requirements. We designed a laser unit for upper stages (LUUS), two types of laser-initiated pyrotechnic devices for solid motor ignition and a separation device actuator (the LID and LCTG) for the kick stage system. Optical fiber paths connecting the LUUS to LID/LCTG enables a continuity check by optical frequency domain reflectometry (OFDR). We successfully conducted continuity checks with OFDR and ran laser ignition tests to validate the design in simulating assembly- and launchsite operations.

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.

Solid propulsion systems offer the potential to transform deep-space mission design by enabling operations in extreme environments beyond the asteroid belt. These systems support larger space probes, reduce power consumption, and lower mission costs, making them particularly suited for the exploration of distant, large gravity bodies, such as Jupiter. Since FY2021, the authors have been developing technologies for a solid rocket motor (SRM) tailored for orbiting and exploring large gravity bodies beyond the asteroid belt. The initial phase has focused on gathering experimental data on the long-term stability, reliable ignition, and steady combustion of SRMs under simulated space conditions, including low temperatures, high vacuum, and intense radiation. This paper presents the results of a conceptual spacecraft system study, preliminary SRM design considerations, and the current progress of the experimental research aimed at optimizing the SRM for deep-space exploration.

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

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.

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.

宇宙科学研究所（ISAS）の宇宙輸送系専門委員会において策定されている宇宙科学・探査分野における宇宙輸送系の中長期ミッションシナリオでは，2040年頃の達成を目指すミッションを「多様な宇宙科学の世界をカバーする軌道間輸送ネットワークを構築する」ことと設定している． ミッション達成に向けて，地上から地球周回軌道へ高頻度に大量の物資や人員を輸送する宇宙往還機を実現するためには，往還飛行において「極超音速飛行」を避けることはできない．本稿では，主に宇宙往還機の実現と競争力向上に資する極超音速エンジンの技術実証を中心に，極超音速飛行に係る技術課題に実証的に取り組むための飛行実験機の目的と検討例について紹介する．

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 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 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.

<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>

<p>The development of enhanced propulsion system for the next Epsilon rocket was progressed. The development of Enhanced Epsilon is mainly the renewal of the second stage, and also includes each subsystem's improvement. The second stage motor M-35 was newly designed and manufactured. In order to verify the design, the static firing test of the second motor M-35 under the condition of vacuum ambient was conducted in 2015. The JAXA successfully launched the first Enhanced Epsilon launch vehicle. All solid propulsion systems for the Enhanced Epsilon launch vehicle showed a very good behavior during the flight</p>

<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>

資料番号: PA1510016000

This paper describes the acoustic measurement of the sub-scale booster and launchpad. The 1:42 scale solid propellant booster was settled over the launch pad model on a flat plate. Designed to deflect the hot and high speed jet plume, the launchpad model was expected to mitigate acoustic impact fed back toward the vehicle. The launchpad model plays a role in attenuating the sound due to the impingement of the plume and the deflector. To investigate the acoustic field with a different booster height, acoustic measurement was carried out. The measurement involved the conventional acoustic measurement and the sound source localization. The conventional measurement employed the near-field microphones around the booster model and the far-field microphones on an arc centered on the base of the launchpad. In the sound source localization, a phased array microphone system was settled to focus the deflector exit. The obtained acoustic data helped revise the design of the launchpad model.

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.

イプシロンロケット二段階開発の最初のステップでは,本質的な低コスト化と即応性の向上を目指す革新的機体システム技術の開発に重きを置いている.推進系の開発においては,H-IIAやM-Vの開発で培われた技術を最大限活用することによって,期間,コスト,リスクを抑え,革新的機体システム技術の早期実証及び近い将来の小型衛星打上げの要求に応える.第1段モータには基幹ロケットのSRB-Aモータを共用し,第2段,第3段にはM-V-5号機の第3段モータ,キックモータをほぼそのまま流用してM-Vをしのぐ輸送効率を達成する.推進系の新しい開発課題は,多様なミッションへの対応能力を高めるPBSの小型液体推進系,そして第1段推力飛行中のロール制御と同コースティング中の3軸制御を担うSMSJ装置である.2013年度の初飛行を目指すイプシロンの推進系開発は,2011年度内に詳細設計を完了して初号機製造に進む見通しである.

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.

イプシロンロケットの目的は,小型衛星に対して即応性豊かな打ち上げシステム,すなわち自在性と機動性に富みユーザーフレンドリな輸送手段を構築,宇宙への敷居を下げて宇宙科学や宇宙利用の裾野を拡大することにある.一方,これを輸送系の視点でみると,打ち上げシステムの革新というひと言に尽きる.すなわち,今後のロケット開発にあたっては,射場設備と運用はもとより,製造プロセスから搭載系に至るまで,およそロケットの打ち上げに必要な設備や運用をとことんコンパクトで身軽なものにしていこう,それが未来への扉を開く鍵であるという理念である.イプシロンロケットでは,このような壮大なビジョンを実現する第一歩として,ロケットのインテリジェント化やモバイル管制などの超革新技術を開拓,これを世界に先駆けて実証するために,初号機を2013年度に打ち上げる計画である.

無毒で常温貯蔵可能な液体推進剤として亜酸化窒素(N_2O)/エタノールの組合せに着目し,それによる扱い易い液体推進系の実証研究を進めている.当面の目標として大気吸い込み式極超音速推進系の飛行試験に用いる加速用ロケットエンジンへの適用を目指しているが,その低温環境順応性を活かす衛星・探査機搭載推進系への応用も視野に入れている.これまでに,推力700N級の要素試験供試体を用いた燃焼試験を2シリーズ行って,エンジン噴射器設計のための有用なデータと運用特性を取得してきた.併せて,水冷式燃焼器による燃焼器壁面熱流束分布の測定や厚肉のシリカ繊維強化プラスチックSFRP製燃焼器を用いた燃焼試験によって燃焼器への耐熱複合材料適用の可能性も探っている.

The erosive burning effects in a small test motor loaded with highly aluminized 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 for several combustion pressure levels and ranges of mass flux in the port. A new simple correlative equation, which includes the effects of mass flux, mass burning rate, combustion pressure and motor scale, was derived from the results of DSM firing tests. The parameters of correlative equation were finally determined by taking into account of the experimental results of cylindrical test motors. In addition, the correlative equation was applied to practical full-scale motors, and the predicted pressure-time curves agreed well with experiments. As the result, a pragmatic correlative equation applicable to grain design application was obtained.