基本情報
- 所属
- 国立研究開発法人宇宙航空研究開発機構 宇宙科学研究所 宇宙機応用工学研究系 教授
- 学位
- 博士（工学)(2000年3月 東京大学)
- 研究者番号
- 10342619
- J-GLOBAL ID
- 202101019944115931
- researchmap会員ID
- R000018454
- 外部リンク
主要な受賞
18論文
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Journal of Guidance, Control, and Dynamics 25-29 2025年3月11日 査読有り
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Acta Astronautica 226(P1) 772-781 2024年11月13日 査読有り
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Aerospace Science and Technology 155(P1) 2024年9月23日 査読有り
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Journal of Evolving Space Activities 2 2024年9月2日 査読有り
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Journal of Evolving Space Activities 2 2024年9月2日 査読有り
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Transactions of the Japan Society for Aeronautical and Space Sciences 67(1) 23-31 2024年1月4日 査読有り
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Transactions of the Japan Society for Aeronautical and Space Sciences 66(6) 199-208 2023年11月4日 査読有り
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The Journal of the Astronautical Sciences 69 1726-1743 2022年11月11日 査読有り
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JOURNAL OF SPACECRAFT AND ROCKETS 59(2) 651-659 2022年3月29日 査読有り
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JOURNAL OF GUIDANCE, CONTROL, AND DYNAMICS 46(4) 695-708 2022年2月1日 査読有り
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Journal of Guidance, Control, and Dynamics 45(2) 280-295 2022年2月1日 査読有り
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Transactions of the Society of Instrument and Control Engineers 58(3) 194-201 2022年
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 19(3) 334-343 2021年5月4日 査読有り
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Proceedings of the International Astronautical Congress, IAC C4 2021年In this paper, we introduce the 500N class bipropellant ceramic thruster for SLIM (Smart Lander for Investigating Moon). It has three main features. The first is a silicon nitride ceramic chamber. It is the only one used by MHI. Second, it has a wide operating range and provides stable performance. It can perform blowdown operation without requiring a high-pressure gas tank. The third activity is Pulse firing which is considered difficult in 500N class thrusters, is possible.
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AIAA Propulsion and Energy Forum, 2021 2021年SLIM (Smart Lander for Investigating Moon) is JAXA’s new lunar lander mission under development. The system firing test was conducted to acquire experimental data of transient behavior of propellant pressure in order to verify the requirement for the propulsion system design and the simulation model, using real propellants and the flight like model of SLIM propulsion system. We observed three remarkable phenomena, which are (1) wide frequency distribution in pressure fluctuation, (2) water hammer when the main engine’s valves are closed, and (3) cross talking. From the evaluation of test results and the combination with simulation, we understood the physical phenomena with real propellants and the flight like model. As a result, we concluded that the current design of SLIM propulsion system is compatible to all the requirements and the expected operational mode of SLIM.
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Journal of Guidance, Control, and Dynamics 44(4) 854-861 2020年12月1日 査読有り
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Acta Astronautica 176(6) 438-454 2020年11月1日 査読有り© 2020 IAA Onboard computation of a fuel-optimal trajectory is an indispensable technology for future lunar and planetary missions with pinpoint landings. This paper proposes a throttled explicit guidance (TEG) scheme under bounded constant thrust acceleration. TEG is capable of achieving fuel-optimal large diversions with good accuracy and can find optimal solutions. Thus far, the TEG algorithm is unique as it offers an explicit and simultaneous search method for the fuel-optimal thrust direction and thrust magnitude switching in predictor-corrector iterations. Fast numerical search is realized with a straightforward computation of seven final states (position, velocity, and the Hamiltonian) from seven unknowns (six adjoint variables for position and velocity and one final time). In addition, global convergence capability is enhanced by implementing the damped Newton's method. A number of simulations of large diversions show the excellent convergence of the TEG algorithm within at most 15 iterations from a cold start. The experimental results of the runtime measurement of the TEG algorithm support its real-time feasibility on a flight processor. These features of the TEG are suitable for onboard guidance of pinpoint landings.
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日本航空宇宙学会論文集 68(2) 89-95 2020年8月1日 査読有り<p>A Fault Detection, Isolation, and Recovery (FDIR) algorithm for attitude control systems is a key technology to increasing the reliability and survivability of spacecraft. Micro/nano interplanetary spacecraft, which are rapidly evolving in recent years, also require robust FDIR algorithms. However, the implementation of FDIR algorithms to these micro/nano spacecraft is difficult because of the limitations of their resources (power, mass, cost, and so on). This paper shows a strategy of how to construct a FDIR algorithm in the limited resources, taking examples from micro deep space probe PROCYON. The strategy focuses on function redundancies and multi-layer FDIR. These ideas are integrated to suit the situation of micro/nano interplanetary spacecraft and demonstrated in orbit by the PROCYON mission. The in-orbit results are discussed in detail to emphasize the effectiveness of the FDIR algorithm. </p>
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Aerospace 7(97) 97-97 2020年7月1日 査読有りJAXA’s ERG (Exploration of Energization and Radiation in Geospace) Spacecraft, which is nicknamed Arase, was launched on 20 December 2016. Arase is a spin-stabilized and Sun-oriented spacecraft. Its mission is to explore how relativistic electrons in the radiation belts are generated during space storms. Two different on-ground attitude determination algorithms are designed for the mission: A TRIAD-based algorithm that inherits from old missions and a filtering-based new algorithm. This paper, first, discusses the design of the filtering-based attitude determination algorithm, which is mainly based on an Unscented Kalman Filter (UKF), specifically designed for spinning spacecraft (SpinUKF). The SpinUKF uses a newly introduced set of attitude parameters (i.e., spin-parameters) to represent the three-axis attitude of the spacecraft and runs UKF for attitude estimation. The paper then presents the preliminary attitude estimation results for the spacecraft that are obtained after the launch. The results are presented along with the encountered challenges and suggested solutions for them. These preliminary attitude estimation results show that the expected accuracy of the fine attitude estimation for the spacecraft is less than 0.5°.
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DESIGN AND DEVELOPMENT OF FIBER OPTIC ROTATION SENSORS 113-124 2019年7月1日
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Proceedings of the International Astronautical Congress, IAC 2019- 2019年This paper focuses on an onboard method of computing a fuel-optimal trajectory for lunar and planetary pinpoint landings. We propose a throttled explicit guidance (TEG) scheme under a bounded thrust magnitude. The TEG algorithm is unique as it offers an explicit and simultaneous search method for the fuel-optimal thrust direction and magnitude switching in predictor-corrector iterations. The thrust direction is modeled exactly as an optimal solution whereas the thrust magnitude switching is obtained by evaluating a quadratically approximated thrust switching equation with its zeroth-order coefficient approximated to a constant value. These models are based on fuel-optimal control theory and enable a fast numerical search with a straightforward computation of seven final states (position, velocity, and the Hamiltonian) from seven unknowns (six adjoint variables for position and velocity and one final time). The Monte Carlo analysis shows an excellent convergence of the TEG algorithm to the optimal solutions within at most 22 iterations from a cold start. In addition, the zeroth-order coefficient of the thrust switching equation shows the best fuel optimality when it is taken around a nominal final mass of a lander. Nonetheless, it is remarkable that the fuel optimality is almost maintained in the order of only 0.1 % increase of the total fuel consumption for the worst case, even if the ambiguity exists on the value of the zeroth-order coefficient. This suggests that TEG does not necessarily require a precise estimate of the final mass or careful selection of the zeroth-order coefficient as a prerequisite to finding fuel optimal solutions. These results support TEG as being suitable for onboard guidance during pinpoint landings.
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Earth, Planets and Space 70(1) 102 2018年12月1日The exploration of energization and radiation in geospace (ERG) satellite, nicknamed "Arase," is the second satellite in a series of small scientific satellites created by the Institute of Space and Astronautical Science of the Japan Aerospace Exploration Agency. It was launched on December 20, 2016, by the Epsilon launch vehicle. The purpose of the ERG project is to investigate how high-energy (over MeV) electrons in the radiation belts surrounding Earth are generated and lost by monitoring the interactions between plasma waves and electrically charged particles. To measure these physical processes in situ, the ERG satellite traverses the heart of the radiation belts. The orbit of the ERG is highly elliptical and varies due to the perturbation force: the apogee altitude is approximately 32,200-32,300 km, and the perigee altitude is 340-440 km. In this study, we introduce the scientific background for this project and four major challenges that need to be addressed to effectively carry out this scientific mission with a small satellite: (1) dealing with harsh environmental conditions in orbit and electromagnetic compatibility issues, (2) spin attitude stabilization and avoiding excitation of the libration by flexible structures, (3) attaining an appropriate balance between the mission requirements and the limited resources of the small satellite, and (4) the adaptation and use of a flexible standardized bus. In this context, we describe the development process and the flight operations for the satellite, which is currently working as designed and obtaining excellent data in its mission.
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Acta Astronautica 152 299-309 2018年11月1日© 2018 IAA This study proposes a solar sailing method for angular momentum control of the interplanetary micro-spacecraft PROCYON (PRoximate Object Close flYby with Optical Navigation). The method presents a simple and facile practical application of control during deep space missions. The developed method is designed to prevent angular momentum saturation in that it controls the direction of the angular momentum by using solar radiation pressure (SRP). The SRP distribution of the spacecraft is modeled as a flat and optically homogeneous plate at a shallow sun angle. The method is obtained by only selecting a single inertially fixed attitude with a bias-momentum state. The results of the numerical analysis indicate that PROCYON's angular momentum is effectively controlled in the desired directions, enabling the spacecraft to survive for at least one month without momentum-desaturation operations by the reaction control system and for two years with very limited fuel usage of less than 10 g. The flight data of PROCYON also indicate that the modeling error of PROCYON's SRP distribution is sufficiently small at a small sun angle (<10°) of the order of 10−9 Nm in terms of its standard deviation and enables the direction of the angular momentum around the target to be maintained.
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航空宇宙技術 17 35-43 2018年3月1日<p>SLIM (Smart Lander for Investigating Moon) is the Lunar Landing Demonstrator which is under development at ISAS/JAXA. SLIM demonstrates not only so-called Pin-Point Landing Technique to the lunar surface, but also demonstrates the design to make the explorer small and lightweight. Realizing the compact explorer is one of the key points to achieve the frequent lunar and planetary explorations. This paper summarizes the preliminary system design of SLIM, especially the way to reduce the size.</p>
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Advances in the Astronautical Sciences 162 1175-1193 2018年JAXA’s ERG (Exploration of Energization and Radiation in Geospace) Spacecraft, which is nicknamed Arase, was launched on 20 December 2016. Arase is a spin-stabilized and Sun-oriented spacecraft. Its mission is exploring how relativistic electrons in the radiation belts are generated during space storms. Two different on-ground attitude determination algorithms have been designed for the mission: a conventional straightforward algorithm that inherits from old missions and an advanced new algorithm. This paper discusses the design of the advanced attitude determination algorithm and presents the preliminary attitude estimation results for the spacecraft that were obtained after the launch. Results are presented along with the encountered challenges and suggested solutions.
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Proceedings of the International Astronautical Congress, IAC 2 856-866 2017年This paper proposes a guidance law that is suitable for the terminal phase of a precise lunar landing. During this phase, as a spacecraft continues its vertical descent for a few minutes until touchdown, such preparatory actions for landing as vertical braking, terrain relative navigation, position correction maneuver, and obstacle detection and avoidance must be taken continuously or simultaneously. The developed guidance law can produce a vertical descent trajectory with low calculation resources, where the fuel consumption and maneuver time for horizontal position correction are minimized. Moreover, the developed law is designed to output two indexes ( and ) that indicate the feasibility of vertical braking and horizontal position correction prior to trajectory computation, in order to prevent any divergence. The simulation results verify that the proposed law performs effectively in evaluating the feasibility of a trajectory based on the discriminants and , in addition to computing trajectories for minimizing fuel consumption and maneuver time when both discriminants are greater than or equal to zero.
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Advances in the Astronautical Sciences 160 2615-2629 2017年When quaternions are used for representing the attitude of a spinning spacecraft in an attitude estimation filter, several problems appear due to their rapid variations. These problems include numerical integration errors and violation of the linear approximations for the filter. In this study, we propose representing the attitude of a spinning spacecraft using a set of spin parameters. These parameters consist of the spin-axis orientation unit vector in the inertial frame and the spin phase angle. This representation is advantageous as the spin axis direction components in the inertial frame do not change rapidly and the phase angle changes with a constant rate in the absence of a torque. The attitude matrix and the kinematics equations are derived in terms of spin parameters. As the equations are highly nonlinear an Unscented Kalman Filter (UKF) is implemented to estimate the spacecraft's attitude in spin parameters. The estimation results are compared with those of a quaternion based UKF in different scenarios using the simulated data for JAXA's ERG spacecraft.
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SPACEFLIGHT MECHANICS 2017, PTS I - IV 160 2615-2629 2017年When quaternions are used for representing the attitude of a spinning spacecraft in an attitude estimation filter, several problems appear due to their rapid variations. These problems include numerical integration errors and violation of the linear approximations for the filter. In this study, we propose representing the attitude of a spinning spacecraft using a set of spin parameters. These parameters consist of the spin-axis orientation unit vector in the inertial frame and the spin phase angle. This representation is advantageous as the spin axis direction components in the inertial frame do not change rapidly and the phase angle changes with a constant rate in the absence of a torque. The attitude matrix and the kinematics equations are derived in terms of spin parameters. As the equations are highly nonlinear an Unscented Kalman Filter (UKF) is implemented to estimate the spacecraft's attitude in spin parameters. The estimation results are compared with those of a quaternion based UKF in different scenarios using the simulated data for JAXA's ERG spacecraft.
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Transactions of the Japan Society for Aeronautical and Space Sciences 60(3) 181-191 2017年© 2017 The Japan Society for Aeronautical and Space Sciences. This paper describes development strategies and on-orbit results of the attitude determination and control system (ADCS) for the world's first interplanetary micro-spacecraft, PROCYON, whose advanced mission objectives are optical navigation or an asteroid close flyby. Although earth-orbiting micro-satellites already have ADCSs for practical missions, these ADCSs cannot be used for interplanetary micro-spacecraft due to differences in the space environments of their orbits. To develop a new practical ADCS, four issues for practical interplanetary micro-spacecraft are discussed: initial Sun acquisition without magnetic components, angular momentum management using a new propulsion system, the robustness realized using a fault detection, isolation, and recovery (FDIR) system, and precise attitude control. These issues have not been demonstrated on orbit by interplanetary micro-spacecraft. In order to overcome these issues, the authors developed a reliable and precise ADCS, a FDIR system without magnetic components, and ground-based evaluation systems. The four issues were evaluated before launch using the developed ground-based evaluation systems. Furthermore, they were successfully demonstrated on orbit. The architectures and simulation and on-orbit results for the developed attitude control system are proposed in this paper.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 60(3) 181-191 2017年This paper describes development strategies and on-orbit results of the attitude determination and control system (ADCS) for the world's first interplanetary micro-spacecraft, PROCYON, whose advanced mission objectives are optical navigation or an asteroid close flyby. Although earth-orbiting micro-satellites already have ADCSs for practical missions, these ADCSs cannot be used for interplanetary micro-spacecraft due to differences in the space environments of their orbits. To develop a new practical ADCS, four issues for practical interplanetary micro-spacecraft are discussed: initial Sun acquisition without magnetic components, angular momentum management using a new propulsion system, the robustness realized using a fault detection, isolation, and recovery (FDIR) system, and precise attitude control. These issues have not been demonstrated on orbit by interplanetary micro-spacecraft. In order to overcome these issues, the authors developed a reliable and precise ADCS, a FDIR system without magnetic components, and ground-based evaluation systems. The four issues were evaluated before launch using the developed ground-based evaluation systems. Furthermore, they were successfully demonstrated on orbit. The architectures and simulation and on-orbit results for the developed attitude control system are proposed in this paper.
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2016 AIAA Guidance, Navigation, and Control Conference 2016年The spin-axis tilt, which is also known as dynamic imbalance or coning error, is one of the most significant bias errors deteriorating the attitude determination accuracy for spinning spacecrafts. Although it is a common practical issue for spin spacecraft missions, estimation algorithms for the dynamic imbalance have not been studied and issued well. This paper proposes a simple algorithm for spin-axis tilt estimation. The algorithm is based on the Singular Value Decomposition (SVD) and makes use of the attitude rates estimated by an Unscented Kalman Filter (UKF), along with the star scanner measurements. Its accuracy is demonstrated using the models for the Exploration of Energization and Radiation in Geospace (ERG) spacecraft. Other bias errors’ effects on the estimation accuracy are examined and the results are compared with a straightforward averaging approach for dynamic imbalance calculation.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 14(30) Pt_7-Pt_14 2016年<p>In this study, a crater detection method for a moon-landing system with low computational resources is proposed. The proposed method is applied to the Smart Lander for Investigating Moon (SLIM), which aims for a pin-point landing on the moon. According to this plan, surface images of the moon will be captured by a camera mounted on the space probe, and the craters are to be detected from the images. Based on the positional relationship between detected craters, the method estimates the exact flight position of the space probe. Because the computational resources of SLIM are limited, rapid and accurate crater detection must be performed using fixed-point arithmetic on a field-programmable gate array (FPGA). This study proposes a crater detection method that uses principal component analysis (PCA). The computational processing for crater detection by PCA is performed by product-sum operations, which are suitable for fixed-point arithmetic. Moreover, this method is capable of parallel processing; hence high-speed processing is expected. This study not only introduces a crater detection method using PCA but also evaluates the properties of this method.</p>
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日本航空宇宙学会論文集 63(6) 257-264 2015年 査読有りIn this paper is presented a microgravity experiment system utilizing a high altitude balloon. The feature is a double shell structure of a vehicle that is dropped off from the balloon and a microgravity experiment section that is attached to the inside of the vehicle with a liner slider. Control with cold gas jet thrusters of relative position of the experiment section to the vehicle and attitude of the vehicle maintains fine microgravity environment. The design strategy of the vehicle is explained, mainly referring to differences from the authors' previous design. The result of the flight experiment is also shown to evaluate the characteristics of the presented system.
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INTERNATIONAL JOURNAL OF MICROGRAVITY SCIENCE AND APPLICATION 32(2) 2015年 査読有り
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航空宇宙技術 14 59-65 2015年In this paper, direction control of balloon gondola with only untwisting motor is proposed. Typically a reaction wheel and another actuator for unloading the reaction wheel are in use to control the attitude (or direction) of the gondola. Although this method can get high accuracy control performance, two actuators spend many resources of the gondola. The proposed method uses only untwisting motor installed above the gondola to rotate. This method can not realize such high accuracy control performance but realize direction control with the most simple configuration. The proposed method is applied to prototype GAPS (General Anti-Particle Spectrometer) balloon experiment in 2012. This paper shows control design for this experiment and the results of the proposed method.
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SPACE SCIENCE REVIEWS 184(1-4) 259-274 2014年11月 査読有りHISAKI (SPRINT-A) satellite is an earth-orbiting Extreme UltraViolet (EUV) spectroscopic mission and launched on 14 Sep. 2013 by the launch vehicle Epsilon-1. Extreme ultraviolet spectroscope (EXCEED) onboard the satellite will investigate plasma dynamics in Jupiter's inner magnetosphere and atmospheric escape from Venus and Mars. EUV spectroscopy is useful to measure electron density and temperature and ion composition in plasma environment. EXCEED also has an advantage to measure spatial distribution of plasmas around the planets. To measure radial plasma distribution in the Jovian inner magnetosphere and plasma emissions from ionosphere, exosphere and tail separately (for Venus and Mars), the pointing accuracy of the spectroscope should be smaller than spatial structures of interest (20 arc-seconds). For satellites in the low earth orbit (LEO), the pointing displacement is generally caused by change of alignment between the satellite bus module and the telescope due to the changing thermal inputs from the Sun and Earth. The HISAKI satellite is designed to compensate the displacement by tracking the target with using a Field-Of-View (FOV) guiding camera. Initial checkout of the attitude control for the EXCEED observation shows that pointing accuracy kept within 2 arc-seconds in a case of "track mode" which is used for Jupiter observation. For observations of Mercury, Venus, Mars, and Saturn, the entire disk will be guided inside slit to observe plasma around the planets. Since the FOV camera does not capture the disk in this case, the satellite uses a star tracker (STT) to hold the attitude ("hold mode"). Pointing accuracy during this mode has been 20-25 arc-seconds. It has been confirmed that the attitude control works well as designed.
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SPACE TELESCOPES AND INSTRUMENTATION 2014: OPTICAL, INFRARED, AND MILLIMETER WAVE 9143 2014年SPICA (Space Infrared Telescope for Cosmology and Astrophysics) is an astronomical mission optimized for mid-and far-infrared astronomy with a 3-m class telescope which is cryogenically cooled to be less than 6 K. The SPICA mechanical cooling system is indispensable for the mission but, generates micro-vibrations which could affect to the pointing stability performances. Activities to be undertaken during a risk mitigation phase (RMP) include consolidation of micro-vibration control design for the satellite, as well as a number of breadboarding activities centered on technologies that are critical to the success of the mission. This paper presents the RMP activity results on the micro-vibration control design.
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SPACE TELESCOPES AND INSTRUMENTATION 2012: OPTICAL, INFRARED, AND MILLIMETER WAVE 8442 2012年We present the current status of the development of the SPICA Coronagraph Instrument (SCI). SPICA is a next-generation 3-meter class infrared telescope, which will be launched in 2022. SCI is high-contrast imaging, spectroscopic instrument mainly for direct detection and spectroscopy of extra-solar planets in the near-to-mid infrared wavelengths to characterize their atmospheres, physical parameters and evolutionary scenarios. SCI is now under the international review process. In this paper, we present a science case of SCI. The main targets of SCI, not only for direct imaging but also for spectroscopy, are young to matured giant planets. We will also show that some of known exoplanets by ground-based direct detection are good targets for SCI, and a number of direct detection planets that are suitable for SCI will be significantly increased in the next decade. Second, a general design of SCI and a key technology including a new high-throughput binary mask coronagraph, will be presented. Furthermore, we will show that SCI is potentially capable of achieving 10(-6) contrast by a PSF subtraction method, even with a telescope pointing error. This contrast enhancement will be important to characterize low-mass and cool planets.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 10(28) Pf_15-Pf_19 2012年We are in progress to develop a system for automatic operation of a satellite in order to reduce human load at satellite steady operation phase. The ground station for small satellite REIMEI (INDEX : INnovative-technology Demonstration EXperiment) is used as a test bench for verification of the proposed method. In our new automatic operation system, a scheduler software as a substitutive operator manages all the operations through a unified procedure, including sending command, receiving telemetry, and driving antenna in accordance with an operation time line which is prepared before the operation pass. The scheduler also performs diagnostics of satellite anomaly based upon the received telemetry data and status of the ground station. In case that some anomaly of the satellite is detected, the scheduler initiates an emergency schedule that was prepared depending on the emergency level. The automatic operation system is nearly completed for downlink operations of the data recorder that account for 75% of REIMEI steady operation. This approach is very effective to reduce psychological and physical load of operators.
MISC
41主要な講演・口頭発表等
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Japan Geoscience Union Meeting (JpGU) 2025 2025年3月25日
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IEEE Aerospace Conference 2025年3月5日
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35th AAS/AIAA Space Flight Mechanics Meeting 2025年1月20日
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AIAA Scitech Forum 2025 2025年1月6日
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31st International Display Workshops (IDW'24) 2024年12月4日 招待有り
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16th International Space Conference of Pacific-basin Societies (ISCOPS) 2024年11月20日
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16th International Space Conference of Pacific-basin Societies (ISCOPS) 2024年11月20日
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35th International Photovoltaic Science and Engineering Conference (PVSEC-35) 2024年11月12日
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35th International Photovoltaic Science and Engineering Conference (PVSEC-35) 2024年11月12日 招待有り
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75th International Astronautical Congress (IAC), 2024年10月18日
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75th International Astronautical Congress (IAC), 2024年10月18日
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Asia Oceania Geosciences Society (AOGS) 2024 2024年6月26日
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Asia Oceania Geosciences Society (AOGS) 2024 2024年6月25日 招待有り
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The 55th Lunar and Planetary Science Conference (LPSC) 2024年3月14日
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The 55th Lunar and Planetary Science Conference (LPSC) 2024年3月14日
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The 55th Lunar and Planetary Science Conference (LPSC) 2024年3月12日
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34th International Symposium on Space Technology and Science(ISTS) 2023年6月5日
共同研究・競争的資金等の研究課題
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日本学術振興会 科学研究費助成事業 2013年5月 - 2018年3月
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日本学術振興会 科学研究費助成事業 2011年 - 2013年
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日本学術振興会 科学研究費助成事業 2004年 - 2008年
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日本学術振興会 科学研究費助成事業 2005年 - 2006年
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日本学術振興会 科学研究費助成事業 2003年 - 2004年
