川口 淳一郎

This paper presents a very high altitude, Sun-Synchronous orbit around the Earth for astronomy satellites. It includes a cart wheel orbit in Hill's motion. Considering the perturbation owing to the Earth gravity, the period of the in-plane motion and out-of-plane motion are different and trajectory becomes Lissajous. The results analytically obtained are examined and compared with those via numerical integration. In addition, it is shown that the fuels amount, which is inquired to keep nominal Cart Wheel orbit properties, decreases as the altitude increases. And, the conclusion is that the Cart Wheel orbit is a kind of sun-synchronous orbit and highly practical with affordable correction velocity increment.

In formation flight missions, it is very important to control and keep the relative positions of spacecrafts. Although there have been many researches on the relative position control, most of them assumes that each spacecraft uses the absolute information about its position and velocity. In this case, various optimizations with the absolute information are possible. However, in order to keep the formation strictly and to avoid the collisions, the feedback control with the relative information between the members is effective. The formations of birds, insects and fish and automatic control of cars are classified as this control system. In this control system, the information propagation structure of the system as well as the local control law of each member determines the control performance of the system. This information propagation structure means the adjacency relation of the formation, that is, "whose relative information can each spacecraft acquire". In this study, the information propagation structure of the formation is focused on and how the information propagation structure influences the behavior of the formation is investigated.

The paper proposes the new guidance strategy in the aerodynamic ascent path. This guidance form has only four parameters comprised of Linear and Logarithmic functions. And the paper presents the guidance results for the practical application cases with certain constraints. The results show the guidance is well performed and satisfies the terminal boundary conditions specified. This guidance strategy does not include the optimization process aboard but solves a simple two-by-two linear algebraic equations along with forward integration. This strategy does guarantee the robust and real-time solutions, excluding any optimization process, and it is concluded quite practical.

The paper proposes the new guidance strategy in the aerodynamic ascent path. This guidance form has only four parameters comprised of Linear and Logarithmic functions. And the paper presents the guidance results for the practical application cases with certain constraints. The results show the guidance is well performed and satisfies the terminal boundary conditions specified. This guidance strategy does not include the optimization process aboard but solves a simple two-by-two linear algebraic equations along with forward integration. This strategy does guarantee the robust and real-time solutions, excluding any optimization process, and it is concluded quite practical.

Discussed in this paper are the results of the mission analysis of near Earth asteroid flyby missions using miniature Asteroid Interceptors. The Interceptor is an autonomous self-contained interplanetary probe with 10kg mass which is now under development in ISAS/JAXA. It has the capability of navigating itself autonomously to flyby the target asteroid using optical navigation system. The image of the asteroid taken by the camera onboard at the closest approach is the main science output of the mission. Firstly discussed is the mission by a single Interceptor, which enables the minimum size interplanetary mission. The interceptor is launched as a piggy back mission on a geostationary mission, separated on a geostationary transfer orbit (GTO), kicked by a solid rocket motor, and injected into an orbit suitable for encountering the asteroid. It is shown that the utilization of the Earth synchronous orbit and the Earth swing-by drastically increase the number of the possible target asteroids, which enables the selection of more scientifically interesting target for a given opportunity. The second mission concept discussed is the multiple asteroids exploration with a single launch. A straightforward application of the single Interceptor mission, that is, the mission by several independent Interceptors is shown firstly, and an option to overcome the difficulty in performing critical operation of multiple spacecrafts simultaneously is also discussed. The list of the target asteroid candidates, detailed mission sequence and maneuver parameters are shown for the assumed example mission.

Japanese first Mars explorer NOZOMI, which was launched in July 1998, suffered several problems during the operation period of more than five years. It could have reached near Mars at the end of 2003, but it was not put into the orbit around Mars. Although NOZOMI was not able to execute its main mission, it provided us a lot of good experiences from the point of the orbit determination of spacecraft. One of the most difficult works was the orbit determination for the period without the telemetry. In this period, for the most of the time the high gain antenna did not point to the earth because of a constraint of the attitude. Therefore, the quality of the tracking data was not good, and for some period it was impossible to get the tracking data at all. Under such critical condition, we managed to get the solution of the orbit, and in a near-miraculous way, we were able to control NOZOMI and execute two earth swingbys successfully. Other issues related to the orbit determination are 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 the conventional way using range and Doppler data. However, we also tried the new method, that is the orbit determination by using the Delta-VLBI method (VLBI: Very Long Baseline Interferometry). In addition to this, we tried optical observations of NOZOMI at the earth swingbys. © 2005 Elsevier Ltd. All rights reserved.

The MUSES-C mission is the world's first sample and return attempt to/from the near Earth asteroid. In deep space, it is hard to navigate, guide, and control 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 on an unknown body. It is important to navigate and 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 various kinds of robotics technologies applied for MUSES-C mission. A global mapping method, an autonomous descent scheme, and a novel sample-collection method, and asteroid exploration robot are proposed and presented in detail. The validity and the effectiveness of the proposed methods are confirmed and evaluated by numerical simulations and some experiments.

今日の電気推進機関の技術進歩により，高比推力の推進機関が実用化されるようになり，さまざまな惑星探査が可能となった．そして今後も，火星や木星をはじめとする，深宇宙の惑星を探査する機会が増えていくと考えられている．宇宙航空研究開発機構・宇宙科学研究本部では，地球近傍ではなく地球引力圏界に，中継拠点として深宇宙港を建設しようという検討を行っている．本研究では，太陽－地球系L2点に深宇宙港を設置することを想定し，低推力推進機関を搭載した宇宙機によるL2点を発着地とした深宇宙往還システムを考え，その脱出軌道を論じる．また同時に，本軌道設計において地球と同期する回帰惑星間軌道上で軌道エネルギーを離心率の拡大によって蓄積する手法(Electric Delta-V Earth Gravity Assist(EDVEGA))が有効であることを示し，EDVEGA軌道への接続を考えた脱出軌道についても提案する． なお，本検討は，その対称性からL1点深宇宙港に対しても，同様に当てはめることができる．

HAYABUSA, which was launched in May 2003, is the first asteroid sample return mission in the world. It has arrived at its destination, Asteroid (25143) Itokawa, at the beginning of September 2005 successfully. HAYABUSA has an ion engine system (IBS) as the main thrusting system and it makes difficult to carry out the orbit determination because of the continuous low acceleration. We have tried many attempts to estimate the acceleration by IES. However, the accuracy of the orbit determination is not good enough when IBS is working, so we make ballistic period of at least three passes once in three weeks or so. We carry out accurate orbit determination at the ballistic period, and we propagate the orbit by using the telemetry information of IBS with some correction by estimating the acceleration by IBS. By this way we have carried out the orbital operation of HAYABUSA. We summarize here the status of the estimation of IBS acceleration. There are several important phases for the orbit operation, such as the launch, the earth swingby, the solar conjunction, and the asteroid arrival. The earth swingby was done on 19 May 2004, and it was successful. We were able to put the spacecraft into the orbit that approach to Asteroid Itokawa. In July of 2005, the solar conjunction occurred and HAYABUSA was opposite side of the Sun. We were able to go through the solar conjunction without any problem, and after that we carried out the optical navigation by using the optical camera on board. In this paper, we also summarize the results of our orbit determination at the launch, the earth swingby, and the solar conjunction. The first result of the optical navigation is mentioned briefly.

Solar sails are the spacecraft that are propelled by sunlight. The Institute of Space and Astronautical Science (ISAS) of Japan Aerospace Exploration Agency (JAXA) has studied the solar sails which are spinning and deployed by centrifugal force. One of the technological difficulties to be realized is how to design the spanned spinning solar sails to maintain the stability while those huge membranes unfurl with deformation and oscillation. In this paper, an attention is focused on the out-of-plane oscillation and an analysis is presented about the dynamics so that the characteristic parameters are identified as for the stability of those spinning solar sails.

Japan Aerospace Exploration Agency (JAXA) is currently studying on the "Solar Sail" propulsion for future deep space explorations. One of the key technologies to realize the solar sail is how light and how compact we can make the photon acceptance surface. JAXA has conducted extensive studies on utilizing centrifugal force to deploy the photon acceptance surface. The final objective is to realize the 7.5μmthickness and 50m diameter polyimide membrane, combined with thin flexible solar cells, as the photon acceptance surface that will be needed around the Jupiter orbit. In the August 9, 2004, JAXA has launched the S-310 sounding rocket, which tested two different shapes of membranes during the zerogravity flight. The first type of the membrane looks like a "clover-leaf", and another is like a "fan". These two membranes, both of them have 10m diameter, were unfolded sequentially during the zerogravity flight under the free spin condition, and their behavior was observed by onboard cameras. This paper focuses on the "clover-leaf" solar sail, which was fully deployed successfully, and introduces the S-310-34 experiments, and then shows the flight results and postflight evaluations.

Solar sails are the spacecraft that are propelled by sunlight. The Institute of Space and Astronautical Science (ISAS) of Japan Aerospace Exploration Agency(JAXA) has studied the solar sails which are spinning and deployed by centrifugal force. One of the technological difficulties to be realized is how to design the spanned spinning solar sails to maintain the stability while those huge membranes unfurl with deformation and oscillation. In this paper, an attention is focused on the out-of-plane oscillation and an analysis is presented about the dynamics so that the characteristic parameters are identified as for the stability of those spinning solar sails. Copyright © 2005 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Institute of Space and Astronautical Science (ISAS) of Japan Aerospace Exploration Agency (JAXA) is currently planning the missions that the small probe 'intercepter' flybys near-Earth objects. In general, it is impossible to determine the relative orbit during fly-by only with optical information. And usually the optical navigation needs combined with the radio navigation that should provide the relative velocity vector information. This paper first discusses the new integrated guidance and navigation strategy, provided the relative velocity vector between the probe and the object is obtained. And next, paper presents how to obtain full orbit property using the camera and velocity collection information for the guidance. Finally, this paper shows the outline of experiment to verify this strategy.

This paper shows the dynamic properties interpreted about a falling cat motion. In the IAC in Vancouver last year, a new interpretation on non-holonomic turns was presented from coning effect point of view. While the last year's synthesis infers even the kinematics effect independent of dynamics may still drive the non-holonomic turn. However, that motion is, in this paper, proved not fully true and shown partly dependent on the inertia properties. This paper presents new interpretation findings which show the resulted non-holonomic turn may become reverse dependent on the moment of inertia ratio. This Fukuoka conference paper corrects the motion structure interpretation, in which a very interesting combination of kinematics and dynamics is found to govern the motion. And a numerical example is given for the spacecraft attitude maneuver, whose time history is explicitly described. A special feedforward control law is derived and applied to the maneuver. The results show the control strategy established well functions and enables the reorientation to be accomplished only via internal torque. What is presented does provide a comprehensive strategy widely applicable to spacecraft and space robots.

In formation flight missions, it is very important to control the relative positions of satellites. Although there have been many researches on the relative position control for formation flight, most of them assumes the centralized architecture in which an administrator collects every relative position and velocity information about the formation. However, in case the formation consists of many satellites, it takes long time to collect the information of all the satellites and to distribute the command to every satellite. Therefore, the centralized architecture is unsuitable for the control system that requires a rapid response. Beside, if the central control is down, the system loses the formation control functions thoroughly. One strategy that overcomes these flaws is to make each satellite be controlled by using the information on its nearby satellites. This paper discusses the guidance law based on the regional limited information to maintain the formation. In this case, the information structure of the system greatly influences the control law and the control performance. In this study, we discuss how the control laws are designed, first. And how the information structure influence the configuration of the formation is focused on. Copyright © 2004 by The American Astronautical Society.

In formation flying missions, the spacecrafts should keep particular formation, so it is very important to control the relative positions between them. Although there have been many researches on the relative position control, most of them assumes that each satellite use the absolute information about its position and velocity. However, in order to keep the formation strictly and to avoid the collisions, the feedback control with the relative information between the members is effective. In such formation maintenance control, 1) the local control law (control gain, etc.) of each spacecraft and 2) "whose relative information can each spacecraft acquire" (we call this "the information propagation structure") determine the behavior of the formation. Especially, the information propagation structure significantly influences the behavior of the formation. However there are few researches about the information propagation structure. In this study, the information propagation structure of the formation is focused on and how the information propagation structure influences the behavior of the formation is investigated. First, we express the information propagation structure of the formation by a matrix and derive the transfer function matrix of the whole system. Next, from this transfer function matrix, we clarify the relation between the information propagation structure and the stability, the steady-state error and the reactivity of the formation and derive the indexes of these control performances. Finally the method for constructing the information propagation structure by optimizing these indexes is shown.

As the authors reported, the optimal steering for the ascending vehicles like a spaceplane, which has the air breathing engines and the large wings, shows oscillatory behavior in contrast to a conventional law for rockets. In the case of more complicated and proper equations of motion, it is difficult to solve two-point boundary value problem on the on-board computer. This paper firstly shows 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. Next, 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 newly proposed guidance scheme in this paper using the approximate expression of the dynamics derived from the result of the optimal steering. Adopting this method, the guidance scheme with only a few parameters is easily obtained by substituting the approximate expression of the dynamics into the equation of motion and can keep maximizing the terminal horizontal speed by the consumption of a few percentage fuel margins.

Institute of Space and Astronautical Science (ISAS) of Japan Aerospace Exploration Agency (JAXA) is currently planning the missions that the small probe 'interceptor' flybys near-Earth objects. Interceptor is very small probe. An interceptor observes spectrum, takes close images, and determines mass of a NEO (near earth object) during a flyby. The weight of interceptor is less than 10 kg. This paper shows three types of missions. In general, it is impossible to determine the relative orbit during flyby only with optical information. Thus the optical navigation needs to be combined with the radio navigation that should provide the relative velocity vector information. In this paper, the integrated guidance and navigation strategy of interceptor is proposed. The interceptor needs the thruster for the attitude and orbit control. This paper introduces the development of the gas-thrust equilibrium thruster for small satellites.

Japanese first Mars explorer NOZOMI, which was launched in July 1998, suffered several problems during the operation period of more than five years. It could have reached near Mars at the end of 2003, but it was not put into the orbit around Mars. Although NOZOMI was not able to execute its main mission, it provided us a lot of good experiences from the point of the orbit determination of spacecraft. One of the most difficult works was the orbit determination for the period without the telemetry. In this period, for the most of the time the high gain antenna did not point to the earth because of a constraint of the attitude. Therefore the quality of the tracking data was not good, and for some period it was impossible to get the tracking data at all. Under such critical condition, we managed to get the solution of the orbit, and in a near-miraculous way, we were able to control NOZOMI and execute two earth swingbys successfully. Other issues related to the orbit determination are 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 the conventional way using range and Doppler data. However, we also tried the new method, that is the orbit determination by using the Delta-VLBI method (VLBI: Very Long Baseline Interferometry). In addition to this, we tried optical observations of NOZOMI at the earth swingbys.

The MUSES-C spacecraft was launched on May 9, 2003 and has been propelled by microwave discharge ion engines. The mission duration is four years in total and most of it is the cruising phase driven by ion propulsion. After one-year acceleration it succeeded the Earth swing-by on May 19, 2004 and still on the way to the asteroid "ITOKAWA". Active thrust vector control contributes to keep the reaction wheels within an appropriate rotation rate. The hydrazine thruster firing tunes off the ion engines before it and restart them again after it. Depending on the operation of ion engines the heater power is adjusted. These operations are autonomously controlled on the spacecraft without supervisions from Earth. A ground operation system that integrates all the process required for ion engine operations such as orbit synthesis, orbit determination, operation scheduling, command generation and telemetry data analysis has been developed. Automation of on-board and ground operations saves the operation cost and time, and makes the operation stable and reliable.

ISAS/JAXA is studying a deployment method using centrifugal force for solar sail mission. In this paper, the clover type sail is investigated. The deployment sequence consists of two stages. In order to analyze the motion of the dynamic deployment, S-310 flight experiment is conducted. In this experiment, the clover type sail of 10 m diameter is deployed dynamically. Numerical simulations by multi-particle model are also conducted to analyze the complicated motion. The experiment and simulation results are compared with each other to validate the analytical model. On the other hand, we schedule to conduct a balloon experiment. The clover type sail of 20 m diameter is deployed statically. The mechanisms for first stage and second stage deployments are introduced.

As one means of propulsion for the future deep space explorers, Japan Aerospace Exploration Agency (JAXA) is conducting research on the "Solar Sail", which is driven by the momentum of photons from the sunlight. Among several candidates of deployment and shape-forming strategies of the sail, JAXA has been focusing on the utilization of centrifugal force. However, it is very difficult to conduct the deployment experiment of the large membrane structure due to aerodynamic drag and gravity. Then, In August in 2004, we performed the experiment to deploy two types of membrane structures made by polyimide film with a sounding rocket. During the ballistic flight, two types of membranes are deployed in turn. We call the two sails "Clover-type sail" and "Fan-type sail", respectively. The cameras on the rocket recorded the images of the membrane and various sensors measured the behavior of the membrane. The data are transmitted to the ground as the telemetry. We succeeded in deploying the membrane whose diameter is 10 meters in space for the first time in the world. This paper deals with the experiment with the rocket and the result of the Clover-type sail.

The paper presents how the formation is controlled to the intended shape by a decentralized control. The formation behaviour is dealt via a z-transformation method and a uniformly convergent strategy that each spacecraft performs is proposed. Not only one-dimensional but tow-dimensional examples are shown. Since the strategy is highly flexible, it is applicable to a variety of the formation flying space missions.

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.

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.

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.

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.

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．

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.

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。

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.

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.

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.

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.

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.

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.

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.

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.

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.