研究者業績

川口 淳一郎

カワグチ ジュンイチロウ  (Jun'ichiro Kawaguchi)

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 宇宙飛翔工学研究系 特任教授 (シニアフェロー)
学位
工学博士

J-GLOBAL ID
200901015159678275
researchmap会員ID
0000023634

学歴

 1

論文

 278
  • Natsume Koichi, Saiki Takanao, Kawaguchi Jun'ichiro
    SPACE TECHNOLOGY JAPAN, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 3 11-18 2004年  
  • Takayuki Yamamoto, Jun'ichiro Kawaguchi
    Advances in the Astronautical Sciences 117 413-426 2004年  
    At the ascent path for the space plane with the air breathing engines and the large wings, the optimal steering law indicates sinusoidal behavior in contrast to a conventional law for rockets. The first primary result of this paper is the analytical description of this periodic behavior emerged from the optimal steering. It concludes that a conventional linear tangent law is applicable only to non-lift vehicles. Then the numerical analysis of the optimal control law with the state constraint is achieved by DCNLP method. In addition, the paper proposes two kind of the new guidance schemes to compensate the flight path error aroused by the disturbance. The first one is the way to approximate the optimal steering using the trigonometric function form and the second one is the way to get the control variable substituting the approximate expression of the dynamics derived from the result of the optimal steering into the equation of motion. These guidance schemes have only a few parameters to be determined and are easily obtained to keep maximizing the terminal horizontal speed by the consumption of a few percentage fuel margins.
  • Takayuki Yamamoto, Jun'ichiro Kawaguchi
    International Astronautical Federation - 55th International Astronautical Congress 2004 1 547-556 2004年  
    The optimal steering for the vehicles like spaceplane shows the different behavior from that of conventional rockets. This is because both thrust and lift force depend on the atmospheric dynamic pressure. This indicates that it is required the new guidance strategy for these kind of vehicles. The authors have proposed the trigonometrical functional form as the guidance strategy to approximate the numerical optimal steering directly. This method showed good performance, but the parameters have to be determined by solving the two-boundary value problem. And its computational load is too heavy to perform the task on the onboard computer. In this paper, we propose another new guidance strategy. This method is derived from some analytical assumptions and approximates the steering in simple linear and logarithmic function. Substituting this relation into the equation of motion, we can get the approximate description as for the vertical acceleration. Carrying out the numerical integration of this, the parameters are determined to satisfy the terminal boundary conditions. This procedure doesn't include the optimization process, but by sweeping the flight time to search parameters that the terminal horizontal speed is maximized, we can get the quasi-optimal solution. And its computational load is relatively light because only numerical integration has to be done to obtain the parameter. The simulation results show almost equivalent to that of the optimization method.
  • Jun'ichiro Kawaguchi, Hitoshi Kuninaka, Akira Fujiwara, Tono Uesugi
    European Space Agency, (Special Publication) ESA SP (542) 25-32 2003年11月  
    The MUSES-C was launched on May 9th of this year and was named 'Hayabusa'. It takes an aim at the world's first sample and return from a near Earth asteroid, 1998SF36 now renamed "Itokawa". The spacecraft is a kind of technology demonstrator with four key technologies. The paper presents a quick report on the initial operation of the ion engines aboard and will show how the attitude control has been performed incorporating the closed loop de-saturation function onboard. The paper also presents how much delta-V has been applied to the spacecraft as well as how the orbit determination under the low-thrust acceleration has been performed.
  • Takanao Saiki, Jun'ichiro Kawaguchi
    54th International Astronautical Congress of the International Astronautical Federation (IAF), the International Academy of Astronautics and the International Institute of Space Law 1 213-221 2003年  
    In formation flight missions, it is very important to control the relative positions of the satellites. Although there have been many researches on the relative position control for the formation flight, most of them assume the centralized architecture in which an administrator collects every the relative positions and velocities information about the formation. This control method is effective for small formations consisting of a few satellites. But, in case the formations consist of many satellites, the amount of the information that should be dealt with becomes huge and the satellites far away from the central control cannot directly communicate with it. Besides if the central control is down, the system loses the formation control functions thoroughly. So, in the large formation flight missions, each satellite should be informed of the proximity relative information by others and control for itself. This paper discusses the new guidance law based on the regional limited information to control and maintain the formation. In this case, the degree of the information sharing significantly influences the formation configuration. In this study, we investigate the influence of the degree of information sharing in formation flight. Copyright © 2003 by the International Astronautical Federation. All rights reserved.
  • M Yoshikawa, J Kawaguchi, H Yamakawa, T Kato, T Ichikawa, T Ohnishi, S Ishibashi
    SPACEFLIGHT MECHANICS 2003, PTS 1-3 114 2199-2216 2003年  査読有り
    The Japanese Mars explorer NOZOMI was launched in July 1998. It was planed to arrive at Mars in October 1999. But a problem occurred when it left from the earth to Mars and it will reach Mars at the beginning of 2004. NOZOMI has several issues in its orbit determination, such as the spin modulation, the solar radiation pressure, the small force related to the attitude change, and the solar conjunction. We tried to solve these issues by using range and Doppler data. Recently, however, much more difficult problem has occurred. That is that the range and/or Doppler data cannot be obtained for some periods because we cannot turn the high gain antenna toward the earth due to the attitude constraint. Therefore we considered the orbit determination by using the Delta-VLBI method. In this paper, we summarize the issues of the orbit of NOZOMI up to now and show our recent activities on Delta-VLBI observations for NOZOMI.
  • T Kubota, T Hashimoto, J Kawaguchi
    PROCEEDINGS OF THE 11TH INTERNATIONAL CONFERENCE ON ADVANCED ROBOTICS 2003, VOL 1-3 1221-1226 2003年  査読有り
    The MUSES-C mission is the world's first sample and return attempt from the near Earth asteroid. In deep space, it is hard to navigate and guide a spacecraft on a real-time basis remotely from the earth mainly due to the communication delay. So autonomy is required for final approach and landing to an unknown body. It is important to guide a spacecraft to the landing point without hitting rocks or big stones. In the final descent phase, cancellation of the horizontal speed relative to the surface of the landing site is essential. This paper describes image processing methods applied for MUSES-C mission. A global mapping method and an image based descent scheme are proposed and presented in detail. The effectiveness of the proposed methods is confirmed by graphical simulations.
  • 山本 高行, 稲葉 歩, 川口 淳一郎
    宇宙技術 2 35-44 2003年  
    本論文では,いわゆる空力上昇径路を飛行する機体の最適誘導則を新たに提案する.まずDCNLP法により最適解を示す.次に直接最適法であるSQP法により別の解を示す.後者の手法ではある直交関数で表現された操舵角を利用することにより,効率的にまた容易に実行することができる.本論文の主な結果は操舵則の解析的表現を示したことである.これは最適性の議論に関連するものである.これによ り従来の線形タンジェント則は揚力を発生しない機体のみに適用可能であることがはっきりと結論される.同時に最適誘導則は三角関数形式を従来の線形タンジェント則に加えることで得られることが結論づけられる.本論文で得られた結果はさらに数値的デモンストレーションによる誘導方策へと最適化プロセスを拡張している.線形化遷移運動が解析モデルによく一致しているため,本論文の結果 は実際的な正当性を示すことに成功している.機体パラメタがノミナル値から変化したり,パラメタ値に対する感度といった誘導計算例もまた示される.
  • Makoto Yoshikawa, Jun'ichiro Kawaguchi, Hiroshi Yamakawa, Takaji Kato, Tsutomu Ichikawa, Takafumi Ohnishi, Shiro Ishibashi
    Advances in the Astronautical Sciences 114(SUPPL.) 2197-2214 2003年  
    The Japanese Mars explorer NOZOMI was launched in July 1998. It was planed to arrive at Mars in October 1999. But a problem occurred when it left from the earth to Mars and it will reach Mars at the beginning of 2004. NOZOMI has several issues in its orbit determination, such as the spin modulation, the solar radiation pressure, the small force related to the attitude change, and the solar conjunction. We tried to solve these issues by using range and Doppler data. Recently, however, much more difficult problem has occurred. That is that the range and/or Doppler data cannot be obtained for some periods because we cannot turn the high gain antenna toward the earth due to the attitude constraint. Therefore we considered the orbit determination by using the Delta-VLBI method. In this paper, we summarize the issues of the orbit of NOZOMI up to now and show our recent activities on Delta-VLBI observations for NOZOMI.
  • 高野忠, 川口淳一郎, 高橋忠幸, 中谷一郎, 橋本樹明, 村上浩, 山川宏
    電子情報通信学会技術研究報告. SANE, 宇宙・航行エレクトロニクス 102(172) 43-50 2002年6月21日  査読有り
    科学衛星ミッションは宇宙研究・観測を目的にしており、宇宙科学研究所の理学・工学両グループおよび外部の共同研究者の緊密な協力の下に進められる。理学グループがその企画をし、工学グループが担当する衛星製作やロケット打上げにも深く関与していくので、典型的ボトムアップ型プロジェクトと言える。本稿では工学実験衛星(MUSES)のミッションも含めて、宇宙科学研究所で進められるミッションを例にして、歴史の概観、ミッション内容さらには研究・開発の仕方について述べる。特にミッションを支える工学技術については、個々のミッションに対応して新しく採用したものと共に、共通的なものを重点的に説明する。
  • Tetsuo Yoshimitsu, Takashi Kubota, Ichiro Nakatani, Jun'ichiro Kawaguchi
    Advances in the Astronautical Sciences 108 491-501 2001年  
    The authors have developed a small robotic lander named "MINERVA" (MIcro/Nano Experimental Robot Vehicle for Asteroid) and proposed it as an optional payload of ISAS's MUSES-C spacecraft which will conduct the sample return mission from newly discovered new-Earth asteroid 1998SF36. This paper describes the mobility system of the developed robot as well as the surface exploration strategy how to explore and navigate along the asteroid surface. Also the scientific observation functions onboard the robot are detailed.
  • T Hashimoto, T Kubota, J Kawaguchi, M Uo, K Baba, T Yamashita
    SPACEFLIGHT MECHANICS 2001, VOL 108, PTS 1 AND 2 108 469-480 2001年  査読有り
    This paper presents an autonomous descent and touch-down scheme of the asteroid sample and return spacecraft, MUSES-C. The spacecraft uses some optical sensors. such as a navigation camera (ONC-W1). a laser altimeter (LIDAR), a short range laser sensor (LRF), and an artificial landmark (TM) which is released at about 100m altitude. Navigation system contains image processing, integration of visual and range information, and Kalman filtering. To realize "time of arrival" guidance, the descending plan is uploaded to the spacecraft, considering the asteroid motion. Six degree-of-freedom control is performed by RCS and reaction wheels (RW). In this paper, after brief explanation of MUSES-C navigation, guidance, and control (NGC) system and descent and touch-down scenario, the navigation scheme mainly focused on image processing, descent guidance scheme, and six degree-of-freedom thruster control are described. To verify the performance of the proposed scheme, computer simulations including Graphical Asteroid Simulator are performed.
  • 山川 宏, 川口 淳一郎
    計測と制御 = Journal of the Society of Instrument and Control Engineers 39(9) 559-563 2000年9月10日  
  • 藤原 顕, 安部 正真, 長谷川 直, 島田 孝典, 小野瀬 直美, 矢野 創, 樋口 健, 沢井 秀次郎, 川口 淳一郎, 高木 周, 高木 靖彦, 高山 和喜, 野中 聡, 岡野 康一, 三輪 治代美, 奥平 俊暁, 矢島 暁
    JASMA : Journal of the Japan Society of Microgravity Application 17(3) 178-182 2000年7月31日  
  • J Kawaguchi, T Hashimoto, T Misu, S Sawai
    ACTA ASTRONAUTICA 44(5-6) 267-280 1999年3月  査読有り
    An impending demand for exploring the small bodies such as the comets and the asteroids envisioned the Japanese MUSES-C mission to the near Earth asteroid Nereus, An autonomous optical guidance and navigation strategy around the asteroid is discussed in this paper. Four major new schemes are dealt with hers: They are (1) Aligned intercept guidance, (2) Strategic building of the flight phases, (3) Image processing of line-of-sight shift information instead of characteristic point tracking, and (4) Stability and accuracy analysis associated with the guidance and navigation strategies developed here. Some comprehensive numerical illustrations are also given to support them. 1999 Elsevier Science Ltd. All rights reserved.
  • Hiroshi Yamakawa, Hirobumi Saito, JuN'Ichiro Kawaguchi, Yasunori Kobayashi, Hajime Hayakawa, Toshinori Mukai
    Acta Astronautica 45(4-9) 187-195 1999年  
    This paper shows the recent results of ISAS Mercury orbiter mission study conducted by ISAS Mercury Exploration Working Group. Two options are under study 1) A Spacecraft which utilizes Solar Electric Propulsion (SEP) as a primary propulsion system for interplanetary transfer phase as well as Mercury orbit insertion phase and 2) Conventional chemical propulsion spacecraft. © 1999 Elsevier Science Ltd. All rights reserved.
  • 川口 淳一郎
    計測と制御 = Journal of the Society of Instrument and Control Engineers 36(9) 655-663 1997年9月  
  • Yasuhiro Morita, Jun'ichiro Kawaguchi, Tatsuaki Hashimoto, Takashi Nakajima, Kenichi Baba, Hiroshi Terada
    Advances in the Astronautical Sciences 96 853-863 1997年  
    The paper reveals the essential feature of the attitude control of the lunar penetrator system and evaluates its fundamental performance. The spinning LUNAR-A mother spacecraft, orbiting a low lunar elliptic orbit, is to release penetrator modules one by one, which penetrate into the moon's surface carrying scientific instruments. This final phase of the journey is featured by a rhumb line controlled attitude maneuver followed by an active nutation control, to ensure the proper impact point attitude. The accuracy of the control will play a key role in the mission as it directly affects the level of the impact load. Although the maneuvering strategy itself cannot be considered special, a relatively high spinning rate of the module makes the problem absolutely different. The level of fluctuation in response time delay of the actuating system has significant influence on the control accuracy: as small as 1 msec of error leads to approximately 0.7 degree of directional dispersion, almost half the required accuracy. Thus a special autonomous delay compensating algorithm has been developed while the active nutation control is also expected to enhance the control capability of the system. The performance of the entire attitude control system has been finally established through a flight test via an ISAS' sounding rocket in January 1997.
  • Hiroshi Yamakawa, Jun'ichiro Kawaguchi, Kuninori Uesugi, Hiroki Matsuo
    Acta Astronautica 39(1-4) 133-142 1996年  査読有り
    Multiple Mercury swingby sequence is applied to Mercury orbiter concept to provide sufficient payload mass. On the other hand, consecutive flybys may become the mission objective itself as was realized by the Mercury flyby mission U.S.Mariner 10 in 1974 - 1975 which included triple consecutive flybys. This paper focuses on this Mariner 10 type flyby mission and investigates the use of an Solar Electric Propulsion (SEP) in the Mercury-Mercury transfer phase, taking the advantage of the solar power availability during the interplanetary cruising inside Earth orbit. The use of SEP upgrades the resultant spacecraft mass as well as increases scientific observation opportunity. As an illustration, a small spacecraft design example is also presented. Copyright © 1997 Elsevier Science Ltd.
  • Jun'ichiro Kawaguchi, Yasuhiro Morita, Tatsuaki Hashimoto, Takashi Kubota, Hiroshi Yamakawa, Hirobumi Saito
    Space Technology 15(5) 277-284 1995年9月  査読有り
    To determine the origin of asteroids and furthermore the solar system, a sample return mission is now planned. This paper presents a mission scenario and the spacecraft design. Some new technologies which must be developed to achieve the mission under strict weight constraint are also described, for example, sampler which must be adaptable to any case of the asteroid surface state, electric propulsion system which is essential to reduce fuel, autonomous navigation of the spacecraft using optical camera, and design of capsule for Earth direct re-entry. © 1995 Elsevier Science Ltd.
  • Journal of Guidance, Control, and Dynamics 18(3) 605-610 1995年5月  
  • Junichiro Kawaguchi, Hiroshi Yamakawa, Tono Uesugi, Hiroki Matsuo
    Acta Astronautica 35(9-11) 633-642 1995年  査読有り
    ISAS (the Institute of Space and Astronautical Science, Japan) is currently planning to launch the LUNAR-A spacecraft to the Moon in 1997 and the PLANET-B spacecraft toward Mars in 1998. Since these two spacecraft have been facing mass budget hurdles, ISAS have been studying how to make good use of lunar and solar gravity effects in order to increase the scientific payload as much as possible. In the LUNAR-A mission, the current orbital sequence uses one lunar swingby via which the spacecraft can be thrown toward the SOI (sphere of influence) boundary for the purpose of acquiring solar gravity assist. This sequence enables the approach velocity to the Moon to be diminished drastically. In the PLANET-B mission, use of lunar and solar gravity assist can help in boosting the increase in velocity and saving the amount of fuel. The sequence discussed here involves two lunar swingbys to accelerate spacecraft enough to exceed the escape velocity. This paper focuses its attention on how such gravity assist trajectories are designed and stresses the significance of such utilization in both missions. © 1995, All rights reserved.
  • Jun'ichiro Kawaguchi, Masafumi Kimura, Hiroshi Yamakawa, Tono Uesugi, Hiroki Matsuo, Robert W. Farquhar
    Advances in the Astronautical Sciences 85(pt 2) 1651-1664 1993年  査読有り
    PLANET-B is a Mars orbiter mission currently under fabrication in ISAS (Institute of Space and Astronautical Science, Japan), whose launch is scheduled in either 1996 or 1998 as a backup. ISAS is planning to make it fly boosted by means of multiple lunar swingbys and its design and results are presented here. This paper presents an universal design chart for the interplanetary spacecraft that may utilize such lunar gravity assist, based on which swingby point is determined. As a verify practical illustration of this scheme, this provides with rigorous trajectory example that is currently a baseline for PLANET-B spacecraft.
  • Hiroshi Yamakawa, Jun'ichiro Kawaguchi, NObuaki Ishii, Hiroki Matsuo
    Advances in the Astronautical Sciences 85(pt 1) 397-416 1993年  査読有り
    Approaching from outside the sphere of influence, a particle may attain low relative velocity was a celestial body and even rotate around it temporarily, without utilizing any other effects than gravitational force. This mechanism is called gravitational capture (i.e. ballistic capture) and it has been considered to be one of the mechanisms which can explain the origin of planetary satellites. In astrodynamics field, lunar gravitational capture is applied to earth-moon transfer trajectory along with positive use of solar perturbation, paying attention to small delta-V required at lunar orbit insertion. In this paper, from the viewpoint of local two-body energy variation, the mechanism of gravitational capture is analytically investigated. Furthermore, a systematic design method for earth-moon transfer trajectory followed by gravitational capture is established.
  • 播磨 浩一, 川口 淳一郎, 中谷 一郎
    日本ロボット学会誌 10(5) 621-631 1992年9月15日  
    This paper describes a retraction control scheme for a space manipulator after grasping a floating object in space. Many control methods for a pre-retracting phase have been proposed by other authors. However those for a retracting phase have been very few, and the essential problem which an object is moving had not been treated so far. Firstly we divide a retracting process into four phases, and clearly show the problems to be resolved. Secondly we derive a new kinematic equation which relates angular velocities of joints with the integral of force and moment at an end-effector in space. Thirdly a force control method is proposed by using that important equation, and the possibility to grasp within the admissible error is proved effective: This control method corresponds to a velocity control for an end-effector feedbacking the integral of force and moment considering the derived kinematic equation. Fourthly we apply that control method to a retraction control, and demonstrate the stability of the proposed control scheme. Finally the effectiveness of the proposed retraction control method is shown by computer simulations.
  • Hiroshi Yamakawa, Jun'ichiro Kawaguchi, Nobuaki Ishii, Hiroki Matsuo
    Advances in the Astronautical Sciences 79(pt 2) 1113-1132 1992年  査読有り
    Gravitational capture is a mechanism by which an object from outside the sphere of influence can orbit around a celestial body temporarily, without any other effects such as atmospheric drag. In this paper, gravitational capture conditions are extensively sought laying emphasis on lunar capture portion using backward time integration mainly in the earth-moon-S/C three-body system. Perilune velocity band satisfying the gravitational capture conditions is found to be constituted of three sub-bands mainly corresponding to its capture direction in the earth-moon fixed rotating frame. An analysis of geocentric orbit with solar effect linking the earth and lunar gravitational capture orbit is also performed for construction of earth-moon transfer orbit. Transfer orbits of various types are designed in the sun-earth-moon-S/C four-body system, which indicate the feasibility of gravitational capture with reasonable ΔV and flight time.

MISC

 218

共同研究・競争的資金等の研究課題

 8