研究者業績
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 助教
- 研究者番号
- 30872444
- ORCID ID
https://orcid.org/0000-0003-1491-1940- J-GLOBAL ID
- 202001000326595612
- researchmap会員ID
- R000000445
研究キーワード
6経歴
4-
2023年8月 - 現在
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2022年12月 - 現在
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2015年10月 - 2022年11月
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2014年4月 - 2015年9月
学歴
3-
2018年9月 - 2021年9月
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2012年4月 - 2014年3月
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2008年4月 - 2012年3月
主要な委員歴
4主要な受賞
11主要な論文
19-
Astronomy & Astrophysics 682(A38) 2024年2月 査読有り筆頭著者最終著者責任著者
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Journal of Guidance, Control, and Dynamics 46(4) 695-708 2023年4月 査読有り筆頭著者責任著者
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Journal of Guidance, Control, and Dynamics 44(4) 854-861 2021年4月 査読有り筆頭著者責任著者
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Acta Astronautica 170 206-223 2020年5月 査読有り筆頭著者責任著者© 2019 IAA This paper investigates the launch capability of the SS-520 as a CubeSat launch vehicle. The SS-520 was developed by JAXA originally as a two-stage, spin-stabilized, solid-propellant sounding rocket. With less than 2.6 tons in total mass and 10 m in length, the SS-520-5 successfully launched a single 3U-sized CubeSat into orbit on February 3, 2018. The SS-520-5 obtained its capability as a CubeSat launch vehicle by installing a 3rd stage solid motor in addition to the RCS between the 1st and 2nd stages. However, its launch capability was limited due to its rocket system configuration. In order to pursue the SS-520's launch capability, two effective modifications from the SS-520-5 are proposed: thrust enhancement of the 1st stage motor and installation of an additional RCS between the 2nd and 3rd stages. The framework of launch capability analysis is established by a multi-objective genetic algorithm (MOGA), where its two objectives are selected as the altitudes of perigee and apogee. The analysis reveals that the two proposed modifications to the SS-520-5 work effectively but differently. The 10% increase of the 1st stage enhancement is particularly effective when the target altitude of perigee is low (e.g., 200 km), whereas the installment of the additional RCS with 30 kg increases accessibility to a much higher altitude of perigee, even to circular orbit reaching altitudes of 550 km for a 1U-sized CubeSat and 280 km for a 6U-sized CubeSat. The simultaneous application of both modifications would result in launch capability able to deliver a 10-kg payload. From a more general perspective, the results in this paper suggest that it is possible for a very small launch vehicle (VSLV) of the 3-ton class and 10 m in length to deliver a 10-kg-class payload into low Earth orbit.
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Acta Astronautica 152 299-309 2018年11月 査読有り筆頭著者責任著者© 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.
主要なMISC
75講演・口頭発表等
27-
Aerospace Engineering Department Seminar, San Diego State University, San Diego, The United States 2024年11月 招待有り
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Seminar, NASA Jet Propulsion Laboratory, Pasadena, The United States 2024年11月 招待有り
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The 75th International Astronautical Congress, Milan, Italy 2024年10月
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The 75th International Astronautical Congress, Milan, Italy 2024年10月
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The 75th International Astronautical Congress, Milan, Italy 2024年10月
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Joint research seminar (among Osaka University, Politecnico di Torino, and ISAS/JAXA), Osaka University 2024年9月
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Optical and Infrared Interferometry and Imaging IX 2024年8月28日 SPIE
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SICE Festival with Annual Conference 2024, Kochi 2024年8月
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The 34th Workshop on JAXA Astrodynamics and Flight Mechanics 2024年7月
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The 2024 European Control Conference, Stockholm 2024年6月
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29th International Symposium on Space Flight Dynamics, Darmstadt, Germany 2024年4月
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Seminar, Exoplanets and Habitability Group, ETH Zurich, Switzerland 2024年4月 招待有り
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33rd International Symposium on Space Technology and Science 2022年3月
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11th International ESA Conference on Guidance, Navigation & Control Systems 2021年6月
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Advances in the Astronautical Sciences 2020年© 2020, Univelt Inc. All rights reserved. ISAS/JAXA has successfully launched the micro-satellite “TRICOM-1R” by the world’s smallest orbit rocket “SS-520 No.5” from Uchinoura Space Center on February 3rd in 2018. ISAS modified the existing sounding rocket SS-520 adding a small 3rd-stage solid-motor and the attitude control system. It flies spinning for the attitude stabilization in the flight. Therefore, we devised the rhumb-line control system with a new scheme. This rhumb-line system has the high-performance functions; the high-preciseness, the high-maneuver rate and the suppression of the unnecessary nutation angle generated at the RCS injection. This paper reports the development of the G&C system and the flight results.
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Proceedings of the International Astronautical Congress, IAC 2019年Copyright © 2019 by the International Astronautical Federation (IAF). All rights reserved. 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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PROMOTE THE PROGRESS OF THE PACIFIC-BASIN REGION THROUGH SPACE INNOVATION 2019年 UNIVELT INCOn February 3, 2018 at the JAXA Uchinoura Space Center, JAXA experimented SS-520 No. 5 launch with a 3U sized cube sat called TRICOM-1R aboard. After liftoff, flight of SS-520 No. 5 proceeded normally. Around 7 minutes 30 seconds into flight, TRICOM-1R separated and was inserted into its target orbit. And the launcher became the world's smallest class satellite launcher. SS-520 launch vehicle is one of sounding rockets operated in JAXA/ISAS, and originally two stage rocket. In this experiment, to make this vehicle put a satellite into orbit, the third stage motor is added. And this sounding rocket has four tail fins for spin stabilization, but usually don't have an attitude control system during the flight. But in this mission, it is needed to control its attitude to ignite second and third motor toward horizontal after first stage bum-out. The gas jet system is installed into between the first stage and the second stage of the vehicle as a unique active attitude control system. The gas jet system can control the spin axis direction and the spin rate of the vehicle during the coasting fight. Because of this constraint, the apogee altitude after the burn out of the first stage motor almost correspond with the perigee altitude of the elliptical orbit. In this mission, the sounding rocket-based Nano launcher is planned to put TRICOM-1R into the elliptical orbit. Its targeted apogee altitude is about 1,800 km and its perigee altitude is about 180 km. Because the perigee altitude is relatively low, the orbit life is very short. One of the mission requirements is to make the vehicle an orbit insertion with more than 30 days orbital lifetime. The vehicle error or the environment error deeply affect the achieved trajectory. These errors must be small enough to put TRICOM-1R into orbit. This paper discusses about the trajectory design on how to manage the sounding rocket into a satellite launching vehicle, the effect of the orbital distribution depending on the various errors, the flight safety analysis, and finally flight performance evaluation.
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SPACEFLIGHT MECHANICS 2019, VOL 168, PTS I-IV 2019年 UNIVELT INCIn order to deeply understand orbital disturbances, the flight data of the PROCYON, which is the 50kg-class interplanetary micro-spacecraft was analyzed. In the telemetry data, we found two unexpected behaviors of angular momentum in Z-axis as compared with the accurate solar radiation pressure model. In order to clarify the causes of the angular momentum anomalies, several small disturbances like thermal radiation pressure, deformation of the structure, and interplanetary magnetic field effect, which are usually ignored are discussed in this study. The thermal radiation and deformation of the structure can explain the over-large Z-axis anomaly. The interplanetary magnetic field effect is correlated with the sudden change of Z-axis torque anomaly in several cases, but the cause of the anomaly is not completely revealed yet.
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Proceedings of the International Astronautical Congress, IAC 2018年Copyright © 2018 by the International Astronautical Federation (IAF). All rights reserved. This paper investigates the launch capability of the SS-520 as a CubeSat launch vehicle. The SS-520 was developed by JAXA originally as a two-stage, spin-stabilized, solid-propellant sounding rocket. With less than 2.6 tons in total mass and 10 meters in length, the SS-520-5 successfully launched a single 3U-sized CubeSat into orbit on February 3, 2018. The SS-520-5 obtained its capability as a CubeSat launch vehicle by installing a 3 rd stage solid motor in addition to the RCS between the 1st and 2nd stages. However, its launch capability was limited (in target altitudes of perigee and apogee at 180 km and 1800 km, respectively) due to its rocket system configuration. In order to pursue the SS-520's launch capability, two effective modifications from the SS-520-5 are proposed: thrust enhancement of the 1st stage motor and installation of an additional RCS between the 2nd and 3rd stages. Furthermore, the framework of launch capability analysis is established by a multi-objective genetic algorithm (MOGA), where its two objectives are selected as the altitudes of perigee and apogee. The problem maintains its simplicity through the selection of only eight design variables of the six acceleration directions and two coasting durations. The analysis reveals that the two proposed modifications to the SS-520-5 work effectively but differently. The 10% increase of the 1st stage enhancement is particularly effective when the target altitude of perigee is low (e.g., 200 km), whereas the installment of the additional RCS with 30 kg increases accessibility to a much higher altitude of perigee, even to circular orbit reaching altitudes of 550 km for a 1U-sized CubeSat and 280 km for a 6U-sized CubeSat. Each modified configuration with the 1st stage enhancement and additional RCS installment enables carrying a payload about twice as heavy as that of the SS-520-5. The application of both modifications would result in launch capability able to deliver a 10-kg payload. From a more general perspective, the results in this paper suggest that it is possible for a very small launch vehicle of the 3-ton class and 10 meters in length to deliver a 10-kg-class payload into low Earth orbit.
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Advances in the Astronautical Sciences 2018年© 2018 Univelt Inc. All rights reserved. On February 3, 2018 at the JAXA Uchinoura Space Center, JAXA experimented SS-520 No. 5 launch with a 3U sized cube sat called TRICOM-1R aboard. After liftoff, flight of SS-520 No. 5 proceeded normally. Around 7 minutes 30 seconds into flight, TRICOM-1R separated and was inserted into its target orbit. And the launcher became the world’s smallest class satellite launcher. SS-520 launch vehicle is one of sounding rockets operated in JAXA/ISAS, and originally two-stage rocket. In this experiment, to make this vehicle put a satellite into orbit, the third stage motor is added. And this sounding rocket has four tail fins for spin stabilization, but usually don’t have an attitude control system during the flight. But in this mission, it is needed to control its attitude to ignite second and third motor toward horizontal after first stage burn-out. The gas jet system is installed into between the first stage and the second stage of the vehicle as a unique active attitude control system. The gas jet system can control the spin axis direction and the spin rate of the vehicle during the coasting fight. Because of this constraint, the apogee altitude after the burn out of the first stage motor almost correspond with the perigee altitude of the elliptical orbit. In this mission, the sounding rocket-based Nano launcher is planned to put TRICOM-1R into the elliptical orbit. Its targeted apogee altitude is about 1,800 km and its perigee altitude is about 180 km. Because the perigee altitude is relatively low, the orbit life is very short. One of the mission requirements is to make the vehicle an orbit insertion with more than 30 days orbital lifetime. The vehicle error or the environment error deeply affect the achieved trajectory. These errors must be small enough to put TRICOM-1R into orbit. This paper discusses about the trajectory design on how to manage the sounding rocket into a satellite launching vehicle, the effect of the orbital distribution depending on the various errors, the flight safety analysis, and finally flight performance evaluation.
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Proceedings of the International Astronautical Congress, IAC 2018年© 2016 Institute of Electrical and Electronics Engineers Inc.. All rights reserved. JAXA has successfully launched the SS-520 No.5 with micro-satellite 'TASUKI' on February 3rd 2018 at Kagoshima Space Center at Uchinoura in Japan. The base-line of the SS-520 sounding rocket is a two-stage rocket which has a capability for launching an 80kg payload to a maximum altitude of about 1000 km. and spun by 4 tails for attitude stability. Enhanced SS-520 No.5 is a three-stage rocket for the smallest-class launch system in the world, which has the orbit injection capability of a micro-satellite of a few kilograms by adding a high-performance third solid motor and advanced rhumb-line control system. Total length of the rocket is about 9.6 meters, the gross weight is 2.6 metric tons, and the reference diameter is 0.52 meters. The `TASUKI` has some experimental purposes for 'store & forward' communication on orbit and earth observation by some commercial cameras and others. The key points of this launch was to newly develop the rhumb-line control system, compact and high performance avionics, some lightweight structures, and the third motor made of CFRP. The rhumb-line control system established an attitude maneuver of about 70 degrees to inject the 'TASUKI' into the orbit of perigee altitude 180km and apogee altitude 2000km. This rhumb-line control system has some high performance functions. It has an angular momentum control function with high attitude maneuver rate, and the suppression function of nutation angle generated by the disturbance of RCS thruster injection during high spin rate of about 1.6Hz. We performed a Motion Table (M/T) Test 'Real-time Simulation Test' with flight models of the avionics for verification of the rhumb-line control design and the soft-wear in the loop test for verification of the flight soft-wear. An active nutation control (ANC) function is also equipped for the reduction of the residual nutation angle after the rhumb-line control. We show the outline of the rocket system and developments, especially the rhumb-line control system with the compact avionics system. Finally flight results are showed and we show one of the future enhanced ideas of SS-520 No.5 type launcher for 10 kilograms class satellite.
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Proceedings of the International Astronautical Congress, IAC 2017年© 2017 International Astronautical Federation IAF. All rights reserved. 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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Proceedings of the International Astronautical Congress, IAC 2017年Copyright © 2017 by the International Astronautical Federation (IAF). All rights reserved. On January 2017, JAXA/ISAS launched the Nano launcher based on the sounding rocket. Unfortunately, the rocket was not able to put a satellite into orbit. This paper discusses the trajectory design regarding how to manage the sounding rocket as a satellite-launching vehicle. At JAXA/ISAS, there are three types of sounding rockets. Two are single-stage rockets called S-310 and S-520, and one is a two-stage rocket called SS-520. These sounding rockets have tail fins for spin stabilization, but usually lack an attitude control system during flight. When attitude control is required to achieve the mission requirements, a gas jet system is available as an optional device. The gas jet system can control the vehicle's spin axis direction and spin rate during coasting fight. To enable the sounding rocket to put a satellite into orbit, a third-stage motor is added to the SS-520 two-stage sounding rocket. The gas jet system is a unique and active attitude control system installed between the first and second stages of the vehicle. Given this constraint, the apogee altitude after burnout of the first-stage motor almost corresponds with the perigee altitude of the elliptical orbit. In this mission, the sounding rocket based Nano launcher is planned to put a 3U-size CubeSat into elliptical orbit. Its targeted apogee altitude is about 1,800 km and its perigee altitude is about 180 km. As the perigee altitude is a relatively low altitude, the orbit life is very short. Thus, any vehicle or environment errors significantly affect the achieved trajectory. Such errors must be small enough to put a CubeSat into orbit. This paper also discusses the effect of orbital distribution depending on various errors.
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The 25th International Symposium on Space Flight Dynamics, 2015年
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The International Symposium on Antennas and Propagation 2013年
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Tsinghua University IAF-SUAC International Student Workshop 2013年
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UN/Japan Nano-Satellite Symposium 2012年
担当経験のある科目(授業)
2-
2024年9月サステイナブル工学先端特論：月惑星着陸探査の最前線 (東京工科大学大学院)
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2024年4月Applied Control in Astronautics (School of Aeronautics and Astronautics, Purdue University)
Works(作品等)
7共同研究・競争的資金等の研究課題
2-
日本学術振興会 科学研究費助成事業 若手研究 2023年4月 - 2026年3月
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宇宙科学研究所 宇宙科学の技術フロントローディング 2021年4月 - 2026年3月
主要な学術貢献活動
5主要な社会貢献活動
20メディア報道
2-
Universe Today Space and astronomy news 2023年11月 インターネットメディア
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Phys.org Astronomy & Space 2023年11月 インターネットメディア
