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  2. Hiroyuki Ogawa

Hiroyuki Ogawa

  (小川 博之)

Profile Information

Affiliation
Professor, Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency
Degree
Doctor of Engineering(Mar, 1996, Nagoya University)
Contact information
ogawa.hiroyukijaxa.jp
J-GLOBAL ID
200901051344540154
researchmap Member ID
1000253790
External link

Research on advanced thermal control systems for future scientific satellites
　Based on the experience of scientific satellite projects, we analyze the current issues and future plans, and conduct research and development of advanced thermal control systems for future scientific satellites. The results of our research have been fed back to the thermal control system on board the X-ray astronomy satellite Hitomi, and are being considered for application to the next scientific satellite project.

Thermal control for scientific satellite projects
　In challenging projects that actively employ thermo-fluid devices, such as the Japan-Europe Mercury mission BepiColombo, which will be exposed to extreme environments that have never been experienced before, and the large X-ray telescope satellite Hitomi, new satellite development methods that have never been experienced before are required. In such challenging projects that actively employ thermo-fluid devices, conventional satellite development methods and their extensions cannot be applied. We are contributing to the success of the project from the viewpoint of heat by leading the new research and development with our academic knowledge of thermo-fluid mechanics, such as development of new materials that can withstand extreme environments, construction of thermal design and analysis methods, development of test facilities, and development of verification methods.

Application of thermo-fluid mechanics
　We are contributing to various space science project activities based on our academic knowledge of thermo-fluid and its related fields. In the research of reusable rockets, we are contributing to the solution of problems related to thermo-fluid such as engine flow, cryogenic tanks, and external flow. In the area of satellite propulsion, we have contributed to the improvement of thruster analysis technology by studying the chemical reaction flow inside hydrazine thrusters, and in the area of rocket propulsion, we have developed a method for analyzing the internal flow of solid rockets and contributed to the investigation of the causes of malfunctions in M-V rockets and SRB-A rockets. In the rocket propulsion system, he developed an internal flow analysis method for solid rockets and contributed to investigating the cause of the failure of the M-V rocket and SRB-A. He has also contributed to rocket research by working on rocket flight safety and radio frequency interference problems with rocket exhaust plumes. I have also conducted theoretical research on shock wave interference in high-speed electromagnetic fluids and propulsion systems using electromagnetic fluids.

Research Interests

 8

Research Areas

 4

Research History

 6
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Education

 1

Committee Memberships

 1

Awards

 1

Papers

 81
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Misc.

 380
  • 鎌田幸男, 川原康介, 坂井智彦, 水野貴秀, 峯杉賢治, 小川博之, 佐藤英一, 中野久松, 尼野理, 安達正樹, 村山直樹
    宇宙科学技術連合講演会講演集(CD-ROM), 50th, 2006  
  • 小川博之, 野中聡, 成尾芳博, 稲谷芳文
    日本航空宇宙学会年会講演会講演集, 37th, 2006  
  • 西田浩之, 小川博之, 船木一幸, 稲谷芳文
    流体力学講演会講演集, 38th, 2006  
  • 吉澤良典, 藤田和央, 小川博之, 稲谷芳文
    流体力学講演会講演集, 38th, 2006  
  • 西田浩之, 小川博之, 船木一幸, 稲谷芳文
    宇宙科学技術連合講演会講演集(CD-ROM), 50th, 2006  
  • 吉澤良典, 藤田和央, 小川博之, 稲谷芳文
    数値流体力学シンポジウム講演論文集(CD-ROM), 20th, 2006  
  • 西田浩之, 小川博之, 船木一幸, 稲谷芳文
    宇宙航空研究開発機構研究開発報告 JAXA-RR-, 5(05-014) 1-8, 2006  
    For the last few years, several deep space propulsion systems making use of the energy of the solar wind have been proposed and researched. Thrust production mechanisms of these propulsion systems, however, have not been understood well, hence they are now intensively studied in the United State and in Japan. In this paper, the thrust characteristics of the Magnetic Sail, which is considered as one of the simplest system in these propulsion systems, are quantitatively obtained by numerical simulations, which successfully revealed the momentum transferring process from the solar wind to the coil of the Magnetic Sail. Also, other propulsion systems making use of the solar wind are introduced in comparison with the Magnetic Sail, so as to reveal the unknown physics and the applicability to deep space missions.
  • 早川基, 小川博之, 笠羽康正, 曽根理嗣, 高島健, 松岡彩子, 向井利典, 山川宏
    宇宙科学シンポジウム, 2006/12/21-22、相模原, 2006, 2006  Peer-reviewed
  • 鎌田幸男, 川原康介, 坂井智彦, 水野貴秀, 峯杉賢治, 小川博之, 佐藤英一, 山川宏, 早川基, 笠羽康正, 向井利典, 中野久松, 石丸元, 尼野理, 安達正樹, BepiColombo
    2006/12/21-22,相模原, 2006,, 2006  Peer-reviewed
  • Kojima Hidenori, Minami Tsubasa, Funaki Ikko, Yamakawa Hiroshi, Fujita Kazuhisa, Ogawa Hiroyuki, Nishida Hiroyuki, Nakayama Yoshinori
    (5) 709-714, Aug 1, 2005  
    A magnetic sail is a unique propulsion system, which travels interplanetary space by capturing the energy of the solar wind. In order to simulate the interaction between the artificial magnetic field produced around a spacecraft and the solar wind, a scaled-down laboratory experiment was conducted in a space chamber. Preliminary results showed some strong interactions between the high-density (10(exp 19)/cu m) and high-velocity (17 km/s) plasma flow and an artificial magnetic field of about 1 T, hence the possibility of the magnetic sail simulator is provided; however, further improvement is required to realize a collision-less solar wind plasma flow in the laboratory.
  • 小松 敬治, 峯杉 賢治, 小川 博之
    宇宙科学シンポジウム, 5 145-148, Jan 6, 2005  
  • 杉田 寛之, 小川 博之, 大西 晃
    宇宙科学シンポジウム, 5 153-156, Jan 6, 2005  
  • 山川 宏, 船木 一幸, 小川 博之
    宇宙科学シンポジウム, 5 495-498, Jan 6, 2005  
  • 小川 博之, 船木 一幸
    ISAS Research Note 791, 1-91, 2005  
  • 小川博之
    平成16年度「サイレンと超音速飛行機実現のための実験 計算融合研究」「レーザー駆動管内加速装置 : 基礎物理の解明と実用展開」合同シンポジウム, 2005  
  • H. Nishida, H. Ogawa, I. Funaki, K. Fujita, H. Yamakawa, Y. Inatani
    41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 2005  Peer-reviewed
    A Magnetic Sail is a deep space propulsion system which captures the momentum of the solar wind by a large artificial magnetic field produced around a spacecraft. To verify the momentum transfer process from the solar wind to the spacecraft, we simulated the interaction between the solar wind and the artificial magnetic field of the Magnetic Sail using the magnetohydrodynamic model. The result showed the same plasma flow and magnetic field structure as those of the Earth. The change of the solar wind momentum results in a pressure distribution along the magnetopause, which is the boundary between the solar wind plasma and the magnetosphere. The pressure on the magnetopause is then transferred to the spacecraft through the Lorentz force between the induced current along the magnetopause and the current along the coil of the spacecraft. The simulation successfully demonstrated that the change of the momentum of the solar wind is transferred to the spacecraft via the Lorentz force. The drag coefficient (thrust coefficient) of the Magnetic Sail was estimated to be 5.0, and it became clear that the Magnetic Sail has weathercock stability. Copyright © 2005 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
  • 西田浩之, 小川博之, 船木一幸, 稲谷芳文
    宇宙科学技術連合講演会講演集(CD-ROM), 49th, 2005  
  • 永田靖典, 大津広敬, 小川博之, 船木一幸, 山川宏, 藤田和央
    宇宙科学技術連合講演会講演集(CD-ROM), 49th, 2005  
  • 南翼, 船木一幸, 小嶋秀典, 山川宏, 中山宜典, 小川博之
    スペース・プラズマ研究会, 2004, 2005  
  • 加藤裕之, 渡辺重哉, 橋本拓郎, 野中聡, 小川博之, 稲谷芳文
    宇宙航空研究開発機構研究開発報告 JAXA-RR-, 4(04-041) 1-12, 2005  
    For a vertical-landing rocket, one of the most important issues is the reliability of vertical soft landing using the opposing jet of the rocket engine. The interaction between engine plumes and the free stream becomes important in the design of an attitude control system in the powered deceleration phase of landing. Since this interacting region has highly nonlinear characteristics, extremely unsteady motion is observed. Measurements for this study were performed in the JAXA 2 x 2 m low-speed wind tunnel using PIV (particle image velocimetry), which improves our imagery and understanding of the complicated flow behavior. From the PIV measurements we obtained two component velocity fields of the opposing jet flow in a vertical model rocket landing, and derived the shape of the opposing jet, stagnation point, and intensity of turbulence. It was also found that the aspect of flow separation in the wake region was changed by the opposing jet.
  • Yamakawa H, Ogawa H, Sone Y, Hayakawa H, Kasaba Y, Takashima T, Mukai T, Tanaka T, Adachi M
    International Astronautical Federation - 56th International Astronautical Congress 2005, 2 778-783, 2005  Peer-reviewed
  • Hiroyuki Ogawa
    International Astronautical Federation - 56th International Astronautical Congress 2005, IAC-05-D2.4.08 5216-5225, 2005  
    A baseline system of a reusable sounding rocket and its studies conducted in ISAS (Institute of Space and Astronautical Science) / JAXA (Japan Aerospace Exploration Agency) are presented. The vehicle adopts a vertical take-off and vertical landing (VTVL) system to minimize ground support equipments (GSE) and turn-around time. An integrated propulsion system which consists of four 2-ton-class liquid-hydrogen / liquid-oxygen (LH2/LOX) expander cycle engines, a gaseous-hydrogen / gaseous-oxygen (GH2/GOX) reaction control system (RCS) and an auxiliary power unit (APU) are introduced in order to simplify ground operations. The vehicle is designed to meet the requirements of carrying 100kg payload on a round trip up to 100km and repeating the mission within 24 hours. The airframe shape is basically axisymmetric; wings are excluded or minimum. The pros and cons of two ways of descent, i.e., nose-forward and base-forward descents, are discussed according to the vehicle system requirements; the nose-forward descent is adopted. The airframe shape is optimized so that the drag is minimum in ascent and the lift-drag ratio is maximum in descent. The 'turn-over' maneuver, i.e., the change from the nose-forward descent attitude to the base-down landing attitude, is required for the nose-forward descent system.
  • M. Carter, A. Baryshev, M. Harman, B. Lazareff, J. Lamb, S. Navarro, D. John, A. L. Fontana, G. A. Ediss, C. Y. Tham, S. Withington, F. Tercero, R. Nesti, G. H. Tan, Y. Sekimoto, M. Matsunaga, H. Ogawa, S. Claude
    Proceedings of SPIE - The International Society for Optical Engineering, 5489(PART 2) 1074-1084, 2004  Peer-reviewed
    The ALMA telescope will be an interferometer of 64 antennas, which will be situated in the Atacama desert in Chile. Each antenna will have receivers that cover the frequencies 30 GHz to 970 GHZ. This frequency range is divided into 10 frequency bands. All of these receiver bands are fitted on a cartridge and cooled, with bands 1 and 2 at 15K and the other 8 are SIS receivers at a temperature of 4K. Each band has a dual polarization receiver. The optics has been designed so that the maximum of the optics is cooled to minimize the noise temperature increase to the receivers. The design of the optics will be shown for each frequency bands. Test results with the method of testing on a near field amplitude and phase measurement system will be given for the first 4 frequency bands to be used, which are bands 3 (84-116 GHz), 6 (211-275GHz), 7 (275-375 GHz and 9 (600-702 GHz). These measurements will be compared with physical optics calculations.
  • Yoshifumi Inatani, Yoshihiro Naruo, Nobuaki Ishii, Hiroyuki Ogawa, Satoshi Nonaka, Hiroshi Yamakawa
    International Astronautical Federation - 55th International Astronautical Congress 2004, 13 8502-8508, 2004  Peer-reviewed
    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 lessons were learned valuable for designing the fully reusable rocket vehicle.
  • ASAHI R, FUNAKI I, FUJITA K, YAMAKAWA H, OGAWA H, NONAKA S, SAWAI S, NAKAYAMA Y, OTSU H
    Papers. American Institute of Aeronautics and Astronautics, 2004  
  • 西田浩之, 小川博之, 船木一幸, 藤田和央, 山川宏, 野中聡, 稲谷芳文
    流体力学講演会講演集, 36th, 2004  
  • NONAKA Satoshi, OSAKO Yosuke, NISHIDA Toshiyuki, OGAWA Hiroyuki, INATANI Yoshifumi
    2004 170-171, 2004  
    For achieving a fully reusable rocket vertical as a future space transportation system, the conceptual designs of vehicle systems are presently being conducted in ISAS/JAXA. In this system design, aerodynamic design considerations are made on a vertical take-off and vertical landing vehicle. One of the considerable issues of a vertical lander is the effect of the interaction between a supersonic nozzle jet and a free-stream when the vehicle is decelerated by the main engine thrust in the landing phase. In order to investigate the influence of such counter-flow interaction in detail, wind tunnel tests were conducted in low speed wind tunnel in ISAS and ISTA/JAXA. The aerodynamic forces and pressure on the base surface were measured by using a scale model of the vehicle. The velocity distribution was measured by a particle image velocimetry (PIV) technique. The aerodynamic characteristics in the vertical landing phase are affected by not only the reduction of the base pressure but also the flow separation around the model side.
  • FUJITA Kazuhisa, FUNAKI Ikko, OGAWA Hiroyuki, YAMAKAWA Hiroshi
    2004 612-613, 2004  
    A numerical analysis of the high-speed plasma flow and the electromagnetic field around a magnetic dipole was performed to investigate acceleration processes of the Magnetoplasma Sail (MPS). In an attempt to take into account deviations of the flow characteristics from the ideal magnetohydrodynamic (MHD) predictions, the particle-in-cell (PIC) model was used with a reduced dimension of the flow and the electromagnetic field. The induction current and the induced electromagnetic field around the magnetic dipole were computed, and the acceleration force acting on the magnetic dipole was roughly estimated. Important scaling parameters in an interaction between the rarefied high-speed plasma and the weak magnetic field are also discussed.
  • NISHIDA Hiroyki, FUNAKI Ikkoh, FUJITA Kazuhisa, OGAWA Hiroyuki, NONAKA Satoshi, NAKAYAMA Yoshinori, OTSU Hirotaka
    2004 614-615, 2004  
    Magnetic sail and Magnetic plasma sail are propulsion systems that make use of the solar wind. These propulsion systems create a large magnetic field around a spacecraft and the magnetic field captures the energy of the solar wind. These propulsion systems are suited for deep space missions because it is estimated to achieve high thrust and efficiency. But there are some problems about these propulsion systems. The process of force transfering from the solar wind to the spacecraft is not understood in detail, thus the metods of estimating thrust vector and controlling thrust vector are not established. We simulated the interaction between the solar wind and the magnetic field of the spacecraft numerically, and verified the method of estimating thrust vector. Additionally we researched the methods of controlling thrust vector.
  • Mukai Toshinori, Hayakawa Hajime, Yamakawa Hiroshi, Ogawa Hiroyuki, Kasaba Yasumasa, Okada Tatsuaki, Suisei Tansa Working Group
    Abstracts for fall meeting of the Japanese Society for Planetary Science, 2004 108-108, 2004  
  • 船木一幸, 藤田和央, 山川宏, 小川博之, 野中聡, 朝日龍介, 中山宜典
    宇宙航空研究開発機構特別資料 JAXA-SP-, JAXA-SP-03-001(03-001) 68-71, 2004  
    A magnetic sail is the way to propel a spacecraft by the solar wind in the interplanetary space. Although original concept of the magnetic sail depends solely on very large magnetic field generated by using such device as superconductors coil, in 2001, Winglee et al. proposed an efficient method to realize a huge magnetic field around a spacecraft with an assistance of plasma emission. From their theoretical analysis of what they call as mini-magnetospheric plasma propulsion (M2P2), it was shown that if a dense plasma were exhausted near the center of the dipole magnetic field, the magnetic field can be expanded far away from the spacecraft, thus the energy of the solar wind can be captured by this huge magnetic field in spite of very low-density solar wind. Based on the idea of such plasma sail, firstly studied were deep space missions targeting at some outer satellites like Jupiter. Plasma sail has great advantage against other electric propulsion systems because of its ability to produce larger thrust to power ratio. However, the thrust formula shown by Winglee et al. is doubtful in some respects. Therefore, an analysis model is discussed that can describe a process of the magnetic field inflation accompanied by plasma emission from the spacecraft.
  • H. Ogawa, S. Nonaka, Y. Inatani
    34th AIAA Fluid Dynamics Conference and Exhibit, AIAA-2004-2538, 2004  
    A baseline system of a sub-orbital reusable rocket and its aerodynamic studies conducted in ISAS (Institute of Space and Astronautical Science) / JAXA (Japan Aerospace Exploration Agency) are presented. The vehicle adopts a vertical take-off and vertical landing (VTVL) system to minimize ground support equipments (GSE) and turn-around time. An integrated propulsion system which consists of four 2-ton-class liquid-hydrogen / liquid-oxygen (LH2/LOX) expander cycle engines, a gaseous-hydrogen / gaseous-oxygen (GH2/GOX) reaction control system (RCS) and an auxiliary power unit (APU) are introduced in order to simplify ground operations and reduce turn-around time. The vehicle is designed to meet the requirements of carrying 100kg payload on a round trip to 120km and repeating the mission within 24 hours. Aerodynamic design considerations are made on the vehicle. The airframe shape is basically axisymmetric; wings are excluded or minimum. The pros and cons of two ways of descent, i.e., nose-forward and base-forward descents, are discussed according to the vehicle system requirements. Since in the base-forward descent the downrange requirement is not met, the nose-forward descent is adopted. The airframe shape is optimized combined with the aerodynamic force calculation so that the drag is minimum in ascent and the lift-drag ratio is maximum in descent. The 'turn-over' maneuver, i.e., the change from the nose-forward descent attitude to the base-down landing attitude, is required for the nose-forward descent system. The interaction between engine plumes and a free-stream become important for design of an attitude control system in the phase of deceleration using engines, and the interaction between an engine jet plume and a ground become important for design of a thermal protection system for landing gears just before the touch-down as well as the take-off. © 2004 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
  • Kamata Yukio, Kawahara Kousuke, Mizuno Takahide, Ogawa Hiroyuki, Yamakawa Hiroshi, Hayakawa Hajime, Mukai Toshifumi
    Proceedings of the Society Conference of IEICE, 2003(1) 71-71, Sep 10, 2003  
  • Hiroyuki Ogawa
    33rd International Conference on Environmental Systems, SAE-2003-01-2689, Jul, 2003  Lead author
  • Yamada Tetsuya, Ogawa Hiroyuki, Nonaka Satoshi, Inatani Yoshifumi, Nakakita Kazuyuki, Yamazaki Takashi
    The Institute of Space and Astronautical Science report. S.P. : Aerodynamics, Thermophysics, Thermal Protection, Flight System Analysis and Design of Asteroid Sample Return Capsule, 17(17) 133-144, Mar, 2003  
    In order to study the onset of the ablation gas-induced boundary layer transition during the reentry phase, heat flux measurement through infrared thermography has been conducted on the capsule-shape body with gas injection from the porous material surface in the shock wind tunnel. In the present simulation experiment, Reynolds number and the ratio of the gas injection mass flow rate to the main mass flow is parametrically changed as similarity law parameters taking account of those in the flight environment; not a few assumptions are applied without verification; the effect of temperature of the boundary layer and the wall etc. At the local Reynolds number of 4 x 10(exp 4) and the ratio of gas injection to main flow is about 2 percent, the heat flux enhancement at the skirt part of the capsule model has been observed and it is considered to be an evidence of the boundary layer transition form laminar to turbulence.
  • 小川 博之, 平木 講儒, 富沢 利夫
    宇宙科学研究所報告. 特集: M-V型ロケット（1号機から4号機まで）, 47(47) 543-550, Mar, 2003  
    M-Vの打ち上げに際し，射点の周囲5箇所における音響計測（音圧レベル）を行っている．これまでに得られている結果を報告する．資料番号: SA0200152000
  • 山川 宏, 小川 博之, 藤田 和央
    宇宙科学シンポジウム, 3 385-390, Jan 9, 2003  
  • 高橋 久里子, 横田 力男, 小川 博之
    宇宙科学シンポジウム, 3 97-100, Jan 9, 2003  
  • FUNAKI I, ASAHI R, YAMAKAWA H, FUJITA K, OGAWA H, NONAKA S, SAWAI S, KUNINAKA H, OTSU H
    Papers. American Institute of Aeronautics and Astronautics, 2003  
  • 野中聡, 小川博之, 大迫庸介, 本郷素行, 稲谷芳文
    宇宙航行の力学シンポジウム 平成14年度, 2003  
  • 小川博之, 野中聡, 成尾芳博, 稲谷芳文
    宇宙科学技術連合講演会講演集(CD-ROM), 47th, 2003  
  • 野中聡, 小川博之, 稲谷芳文
    宇宙科学技術連合講演会講演集(CD-ROM), 47th, 2003  
  • 船木一幸, 山川宏, 藤田和央, 小川博之, 野中聡, 沢井秀次郎, 国中均, 大津広敬
    宇宙輸送シンポジウム 平成14年度, 2003  
  • 船木一幸, 山川宏, 小川博之, 藤田和央, 野中聡, 国中均, 大津広敬
    航空原動機・宇宙推進講演会講演集, 43rd, 2003  
  • 朝日龍介, 船木一幸, 南翼, 山川宏, 藤田和央, 小川博之
    宇宙科学技術連合講演会講演集(CD-ROM), 47th, 2003  
  • Hiroyuki Ogawa, Hiroshi Yamakawa, Yasunori Kobayashi, Mitsuhiko Nakano
    32rd International Conference on Environmental Systems, 2002-ICES-236, Jul, 2002  Peer-reviewedLead author
  • Ikkoh Funaki, Hiroyuki Ogawa, Teruo Kato, Takashi Abe, Kazuhisa Fujita, Satoshi Nonaka
    33rd Plasmadynamics and Lasers Conference, 2002  
    Some of solid rocket motor plumes are reported to cause a telecommunication black-out. To clarify the mechanism of this microwave-plume interaction, attenuations of the microwave signals were measured in three ground firing tests for Japanese solid rocket motors, M25, M14, and SRB-A. As a microwave diagnostics, multifrequency microwave technique was employed (S-band, 2.4 GHz, C-Band, 5.6 GHz, and X-Band, 8.4GHz), by which both the electron density and the electron collision frequency were simultaneously determined using theoretical attenuation by a plasma slab model as 3x1016 m-3 and 7x1010 Hz during the effective firing period of the motors. Although this successful determination of plume plasma properties indicated the cause of the telecommunication black-out is plasma, however, near the end of the firing, large attenuations and departure from the theoretical curve will imply another possible attenuation mechanism in addition to the plasma effect depending on the firing period. © 2002 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
  • 山川宏, 藤田和央, 小川博之, 野中聡, 国中均, 船木一幸
    宇宙科学技術連合講演会講演集, 46th(Pt.4), 2002  
  • Hiroyuki Ogawa, Hiroshi Yamakawa, Ken Goto
    ESA-SP-521, 109-115, 2002  Lead author
  • 小川博之, 小林康徳, 中野充彦, 大西晃, 太刀川純孝, 松藤幸男, 長島和宏
    宇宙航行の力学シンポジウム 平成13年度, 2002  

Books and Other Publications

 1

Presentations

 33
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Professional Memberships

 5

Research Projects

 9
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Industrial Property Rights

 6
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Academic Activities

 1

● 指導学生等の数

 6
  • Fiscal Year
    2018年度(FY2018)
    Doctoral program
    1
  • Fiscal Year
    2019年度(FY2019)
    Doctoral program
    2
    Master’s program
    1
    JSPS Research Fellowship (Young Scientists)
    1
  • Fiscal Year
    2020年度(FY2020)
    Doctoral program
    1
    Master’s program
    1
    JSPS Research Fellowship (Young Scientists)
    1
  • Fiscal Year
    2018年度(FY2018)
    Doctoral program
    1
  • Fiscal Year
    2019年度(FY2019)
    Doctoral program
    2
    Master’s program
    1
    JSPS Research Fellowship (Young Scientists)
    1
  • Fiscal Year
    2020年度(FY2020)
    Doctoral program
    1
    Master’s program
    1
    JSPS Research Fellowship (Young Scientists)
    1

● 専任大学名

 2
  • Affiliation (university)
    東京大学（University of Tokyo)
  • Affiliation (university)
    東京大学（University of Tokyo)

● 所属する所内委員会

 6
  • ISAS Committee
    研究所会議
  • ISAS Committee
    プログラム会議
  • ISAS Committee
    信頼性品質会議
  • ISAS Committee
    環境・安全管理統括委員会
  • ISAS Committee
    ISASニュース編集小委員会
  • ISAS Committee
    宇宙科学プログラム技術委員会
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