川口 淳一郎

The Hayabusa spacecraft rendezvoused with the asteroid Itokawa in 2005 after the powered flight in the deep space by the mu 10 cathode-less electron cyclotron resonance ion engines. Though the spacecraft was seriously damaged after the successful soft-landing and lift-off, the xenon cold gas jets from the ion engines rescued it. New attitude stabilization method using a single reaction wheel, the ion beam jets, and the photon pressure was established and enabled the homeward journey from April 2007 aiming the Earth return on 2010. The total accumulated operational time of the ion engines reaches 31,400 hours at the end of 2007. One of four thrusters achieved 13,400-hour space operation.

The Hayabusa spacecraft rendezvoused with the asteroid Itokawa in 2005 after the powered flight in the deep space by the mu 10 cathode-less electron cyclotron resonance ion engines. Though the spacecraft was seriously damaged after the successful soft-landing and lift-off, the xenon cold gas jets from the ion engines rescued it. New attitude stabilization method using a single reaction wheel; the ion beam jets, and the photon pressure was established and enabled the homeward journey from April 2007 aiming the Earth return on 2010. The total accumulated operational time of the ion engines reaches 31,400 hours at the end of 2007. One of four thrusters achieved 13,400-hour space operation.

The Hayabusa spacecraft rendezvoused with the asteroid Itokawa in 2005 after the powered flight in the deep space by the mu 10 cathode-less electron cyclotron resonance ion engines. Though the spacecraft was seriously damaged after the successful soft-landing and lift-off, the xenon cold gas jets from the ion engines rescued it. New attitude stabilization method using a single reaction wheel, the ion beam jets, and the photon pressure was established and enabled the homeward journey from April 2007 aiming the Earth return on 2010. The total accumulated operational time of the ion engines reaches 31,400 hours at the end of 2007. One of four thrusters achieved 13,400-hour space operation.

A team in JSPEC/JAXA has started a phase-A study and development for the high-speed reentry experiment called DASH-II. In the mission, a small capsule inserted into the geosynchronous orbit or a relevant orbit and conducts reentry flight at the super-orbital velocity after several orbital circulations. The main objective of the mission is to acquire and demonstrate the reentry/entry technologies associated with future planetary exploration missions and high-speed reentry from Moon and from farther small bodies such as Marco-Polo sample return. The present paper describes the DASH-II mission overview and recent development status. Copyright 2008 by the International Astronautical Federation. All rights reserved.

The Japanese spacecraft, HAYABUSA, was launched on May 9, 2003 and spent more than 2.5 years approaching the asteroid ITOKAWA. This spacecraft used 13.2 Ah lithium-ion secondary cells. To realize maximum performance of the battery for long flight operation, the state-of-charge (SOC) of the battery was maintained at ca. 65% during storage in case it is required for contingency operations. To maintain this SOC condition, the battery is charged once a week. After HAYABUSA touched down on ITOKAWA in December 2005, four of 11 lithium-ion secondary cells were over-discharged. However, the battery power was still indispensable for sealing the container with the asteroid sample. The seven remaining healthy cells were slowly recharged using minimum current. After its safety was confirmed, the lithium-ion secondary battery was used to transfer, latch, and successfully seal the sample container into the reentry capsule. The seven remaining cells have been slowlydischarged through the resistance of the monitoring circuit. The capacity by the discharge revealed that the storage performance of the lithium-ion secondary battery is predictable based on the ground-test results even after the storage for 5 years in space. Copyright © 2008 by the American Institute of Aeronautics and Astronautics, Inc.

The paper presents the attitude reorientation taking the advantage of solar radiation pressure without use of any fuel aboard. The strategy had been adopted to make Hayabusa spacecraft keep pointed toward the Sun for several months, while spinning. The paper adds the above mentioned results reported in Sedona this February showing another challenge of combining ion engines propulsion tactically balanced with the solar radiation torque with no spin motion. The operation has been performed since this March for a half year successfully. The flight results are presented with the estimated solar array panel diffusion coefficient and the ion engine's swirl torque.

This paper describes the attitude dynamics and the control method of a spinning solar sail spacecraft. The solar sail considered here has no structure supporting membrane, therefore estimation of the effect of membrane flexibility is one of the problems to solve for the future validation flight. In this study, we established a dynamic model including membrane vibration to handle a coupled motion of a rigid spacecraft and a flexible membrane, and focus on the consideration of attitude control method. The result is confirmed with numerical simulation by use of Multi Particle Model(MPM).

The asteroid exploration spacecraft "HAYABUSA" took high resolution pictures of asteroid "ITOKAWA". ITOKAWA is a rubble-pile asteroid, whose shape and regolith size distribution are unique. These features are caused by composition of the Brazil-Nut effect and centrifugal force by spin. The Brazil-Nut effect is generally known as the phenomenon that causes larger particles to rise to the top of the shaken granular mixtures. In this study, the behaviors of various size particles on the ground or in the space are analyzed by numerical simulation using Multi-Particle model, which is model of a rubble-pile asteroid. At first, the behaviors of particles on the ground are simulated, and it is confirmed that Brazil-Nut effect occurs. Next, the behaviors of particles in the space are simulated. From this simulation results, the internal distribution of particles' upward trend is obtained. It is compared with the known information of ITOKAWA, and the inferred internal structure of rubble-pile asteroid is discussed.

This paper presents what the Hayabusa spacecraft experienced when it restarted the three axis stabilization. Due to the loss of two wheels aboard and also due to the loss of fuel, the primary axis of inertia of the spacecraft was tilted offset to the geometrical axis. The instability was observed unexpectedly since there was little thought about the instability associated with the single wheel system. The results can be applied for other spacecraft. The flight results are shown as well.

This paper describes the navigation and the guidance strategy of the small spacecraft for flyby using the images of the target. We derived navigation and guidance accuracies of this mission analytically, and confirmed them by numerical simulation. For high accurate flyby, the center of the target needs to be measured accurately. However the shade area on the target surface makes it difficult to know the target center position from the information of the center of brightness in image. This uncertainty affects the guidance accuracy from the target. In this paper, a method to calculate the target center from the tangent points of the sun is proposed. The validity of the method is examined in the experiment.

This paper presents what the Hayabusa spacecraft experienced when it restarted the three axis stabilization. Due to the loss of two wheels aboard and also due to the loss of fuel, the primary axis of inertia of the spacecraft was tilted offset to the geometrical axis. The instability was observed unexpectedly since there was little thought about the instability associated with the single wheel system. The results can be applied for other spacecraft. The flight results are shown as well.

The paper presents the attitude reorientation taking the advantage of solar radiation pressure without use of any fuel aboard. The strategy had been adopted to make Hayabusa spacecraft keep pointed toward the Sun for several months, while spinning. The paper adds the above mentioned results reported in Sedona this February showing another challenge of combining ion engines propulsion tactically balanced with the solar radiation torque with no spin motion. The operation has been performed since this March for a half year successfully. The flight results are presented with the estimated solar array panel diffusion coefficient and the ion engine's swirl torque.

This paper discusses dynamics of large membrane for achieving spinning solar sail-craft proposed by the Institute of Space and Astronautical Science (ISAS), Japan Aerospace Exploration Agency (JAXA). For comprehending the dynamics, firstly, two types of grand-scale experiments were conducted. One was an ice rink experiment. The other was a balloon experiment. In the both experiments, we succeeded in deploying 10m and 20m diameter membrane and observing the motion. Secondly, we constructed a numerical model using a multi-particle method, and numerical simulations were conducted. We drew a comparison between results of experiments and numerical simulations. The effectively of the multi-particle model was then indicated in this paper.

This paper describes the navigation and the guidance strategy of the small spacecraft for flyby using the images of the target. We derived navigation and guidance accuracies of this mission analytically, and confirmed them by numerical simulation. For high accurate flyby, the center of the target needs to be measured accurately. However the shade area on the target surface makes it difficult to know the target center position from the information of the center of brightness in image. This uncertainty affects the guidance accuracy from the target. In this paper, a method to calculate the target center from the tangent points of the sun is proposed. The validity of the method is examined in the experiment.

New thruster system using less toxic fuel, which vapor pressure is relatively low, has been developed for small or micro size spacecraft, which has less space or weight budget, etc. In this system, the tank store the fuel in liquid state, the thruster system ejects the fuel in gas state using vapor pressure to obtain thrust. This means high energy efficiency and the constant thrust level maintained right before fuel exhausted, compared to the widely used cold gas thruster. To achieve that the thruster system separate the fuel into gas phase from liquid phase, the tank has porous metal inside itself. The porous metal absorbs the liquid state fuel and only gas state fuel can pass through the porous metal. And having the porous metal inside the tank, the thruster system can reduce the effect of sloshing, and the heater control can be achieved easily because of the enhancement of the thermal conductivity. Furthermore, the surface area is drastically increased, so vaporization occurs everywhere in ejecting the fuel, which keeps the constant thrust level. This paper presents the experimental result of this gas-liquid equilibrium thruster. And it is shown that this tank system has the effectiveness mentioned above.

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 an 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. © 2007 Elsevier Ltd. All rights reserved.

The Japan Aerospace Exploration Agency (JAXA) will make the world's first solar power sail craft demonstrate for both its photon propulsion and thin film solar power generation during its interplanetary cruise. The spacecraft deploys and spans its membrane of 20 meters in diameter taking the advantage of the spin centrifugal force. The spacecraft weighs approximately 300kg, launched together with the agency's Venus Climate Orbiter, PLANET-C in June of 2010. This will be the first actual solar sail flying an interplanetary voyage.

Recent Space Exploration demands sometimes assume an apparent gravity capture around the Moon, using the solar tidal force effect accumulated while flying in the far boundary region of the Earth gravity field. However, when/where the capture occurs strongly depends on where the spacecraft departs from and approaches to the Moon. Therefore, such spontaneous capture is hardly applicable to the flexible and practical transportation systems to the Moon. The paper presents how the launch (departure) or swingby (arrival) conditions are relaxed taking the advantage of the artificial acceleration, electric propulsion means.

The Japan Aerospace Exploration Agency (JAXA) is currently studying the Hayabusa-follow-on missions to more primitive bodies. Hayabusa-2 is a replica to the Hayabusa spacecraft and will be ready for quick flight in less than four years. It will go to a C-type asteroid, while Hayabusa visited an S-type asteroid, Itokawa. And Hayabusa mk-II is much larger spacecraft that may go to an extinct comet, Wilson-Harrington. Both the spacecraft are propelled by the ion engines and the trajectory analysis results through an optimization process. The paper presents how those missions scenarios are built utilizing the ED VEGA strategy.

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 an NEO (near earth object) during a flyby. The weight of interceptor is less than 10kg. 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. (c) 2007 Elsevier Ltd. All rights reserved.

Solar sail is one of the promising propulsion systems for future deep space exploration missions as it does not require any fuel to acquire propulsive force. However, the attitude control system of the solar sail, which controls the direction of the sail and thus the propulsive force, has not been much studied, although this constitutes the essential part of the orbital control using solar sail. This paper discusses the attitude dynamics and the control method of a spinning type solar sail spacecraft. The spinning type solar sail has no rigid structure supporting its membrane. This type of mechanism has the advantage in its simple and lightweight structure, however, the attitude control is difficult due to the flexibility of the membrane. In this paper, we introduced a mathematical dynamics model including first vibration mode of the membrane which can handle coupled motion of a rigid spacecraft and a flexible membrane, and analytically developed a controller that can avoid unnecessary oscillatory motion. The performance of the controller and the effect of solar radiation pressure, which can deform the membrane of solar sail, on the controller were verified by numerical simulations using more precise multi-particle numerical model.

This paper deals with how the formation flight target is expressed as trivial in some appropriate coordinates, instead of an orbital target. There is attempted to show the regularization process to obtain a special coordinate through a Levi-Civita non-linear transform. The paper will show how and which the virtual target corresponds to the physical formation configuration.

This paper describes the attitude dynamics and the control method of a spinning solar sail spacecraft. The solar sail considered here has no structure supporting membrane, therefore estimation of the effect of membrane flexibility is one of the problems to solve for the future validation flight. In this study, we established a dynamic model including membrane vibration to handle a coupled motion of a rigid spacecraft and a flexible membrane, and focus on the consideration of attitude control method. The result is confirmed with numerical simulation by use of Multi Particle Model(MPM).

Usually, the formation flying associated with circular orbits is discussed through the well-known Hill's or C-W equations of motion. This paper dares to present and discuss the coordinates that may contain time-varying coefficients. The discussion presents how the controller's performance is affected by the selection of coordinates, and also looks at the special coordinate suitable for designating a target bin to which each spacecraft in the formation has only to be guided. It is revealed that the latter strategy may incorporate the J2 disturbance automatically.

Reusable launch vehicles are conceived to constitute the future space transportation system. If these vehicles use air-breathing propulsion and lift taking-off horizontally, the optimal steering for these vehicles exhibits completely different behavior from that in conventional rockets flight. In this paper, the new guidance strategy is proposed. This method derives from the optimality condition as for steering and an analysis concludes that the steering function takes the form comprised of Linear and Logarithmic terms, which include only four parameters. The parameter optimization of this method shows the acquired terminal horizontal velocity is almost same with that obtained by the direct numerical optimization. This supports the parameterized Liner Logarithmic steering law. And here is shown that there exists a simple linear relation between the terminal states and the parameters to be corrected. The relation easily makes the parameters determined to satisfy the terminal boundary conditions in real-time. The paper presents the guidance results for the practical application cases. The results show the guidance is well performed and satisfies the terminal boundary conditions specified. The strategy built and presented here does guarantee the robust solution in real-time excluding any optimization process, and it is found quite practical. © 2007 Elsevier Ltd. All rights reserved.

Reusable launch vehicles are conceived to constitute the future space transportation system. If these vehicles use air-breathing propulsion and lift taking-off horizontally, the optimal steering for these vehicles exhibits completely different behavior from that in conventional rockets flight. In this paper, the new guidance strategy is proposed. This method derives from the optimality condition as for steering and an analysis concludes that the steering function takes the form comprised of Linear and Logarithmic terms, which include only four parameters. The parameter optimization of this method shows the acquired terminal horizontal velocity is almost same with that obtained by the direct numerical optimization. This supports the parameterized Liner Logarithmic steering law. And here is shown that there exists a simple linear relation between the terminal states and the parameters to be corrected. The relation easily makes the parameters determined to satisfy the terminal boundary conditions in real-time. The paper presents the guidance results for the practical application cases. The results show the guidance is well performed and satisfies the terminal boundary conditions specified. The strategy built and presented here does guarantee the robust solution in real-time excluding any optimization process, and it is found quite practical. 2007 Elsevier Ltd. All rights reserved.

The electron cyclotron resonance ion engine has long life and high reliability because of electrodeless plasma generation in both the ion generator and the neutralizer. Four mu 10s, each generating a thrust of 8 mN, specific impulse of 3200 s, and consuming 350 W of electric power, propelled the Hayabusa asteroid explorer launched on May 2003. After vacuum exposure and several baking runs to reduce residual gas, the ion engine system established continuous acceleration. Electric propelled delta-V Earth gravity assist, a new orbit change scheme that uses electric propulsion with a high specific impulse was applied to change from a terrestrial orbit to an asteroid-based orbit. In 2005, Hayabusa, using solar electric propulsion, managed to successfully cover the solar distance between 0.86 and 1.7 AU. It rendezvoused with, landed on, and lifted off from the asteroid Itokawa. During the 2-year flight, the ion engine system generated a delta-V of 1400 m/s while consuming 22 kg of xenon propellant and operating for 25,800 h.

Discussed in this paper is a lunar approach strategy tolerant of a lunar orbit injection (LOI) failure. LOI is one of the most critical events for a lunar orbiting mission. If LOI is not performed, the spacecraft flies by the moon, and in the worst case, it escapes not only from the moon but also from the earth, which leads to mission failure. The proposed strategy is to design a trajectory so as to re-encounter the moon even when LOI is not performed. It provides an opportunity for mission recovery even in the case of an unexpected fly-by. A trajectory design procedure is introduced and an example of the designed trajectory is shown.

The Moon is recognized as an important destination for space science and exploration. The National Aeronautics and Space Administration (NASA) has presented a new policy for space exploration that includes human expedition to the lunar surface. The other agencies that include the Japan Aerospace Exploration Agency (JAXA) have also planned to send spacecraft to the moon. To put the spacecraft into orbit around the moon, various methods are conceived. For example, one of them is that the special spacecraft, such as an orbital transfer vehicle (OTV), is used in each transfer segment. It is the very important thing in cost that OTV between Low Earth Orbit (LEO) and Lunar Low Orbit (LLO) is reused. To promote lunar exploration efficiently, reusable OTV between LEO and LLO is required. OTV has been studied at JAXA, which is based on further development of H-II Transfer Vehicle (HTV). The OTV system is roughly composed of the payload segment, the propulsion segment and the fuel cartridge. The payload segment is separated from OTV body on Lunar Transfer Orbit (LTO), and is transported to LLO. The propulsion segment and the empty fuel cartridge are returned to LEO through Earth Transfer Orbit. (ETO). This paper aims at improving the payload ratio of reusable OTV. In this research, we focus on optimization problems of return trajectories with small fuel consumption. Flight time in LTO should be short generally. When payload is empty, flight time isn't restricted. So, multi-impulse flight using electrical propulsion is used in ETO. Aero-assisted flight using aero-brake is also utilized near the Earth. In ETO, we discuss ballistic flight using chemical propulsion, multi-impulse flight using electrical propulsion and aero-assisted flight using aero-brake. Optimal return trajectories of OTV between the Earth and the Moon are proposed. This research can be also extended to the transfer vehicles between the Earth and the Lagrange point of Sun-Earth system. The L2 point in the Sun-Earth system may be utilized for a space port for the outer planets exploration in the future. In this paper, we show the concept of the OTV system and the result of the optimal return trajectories for OTV.

The proposed balloon experiment for deploying large-sized membrane, which is intended to be shown in this paper, is postponed to September, 2008. We are now designing a new structure and mechanism of a spacecraft which is to be validated at the balloon experiment. We proposed "Multi-step deployment" of membrane in this paper, which reduces the risk of complexity of the deploying system compared to conventional system. In addition, we analyzed the feasibility of attitude control required for orbital maneuver with conventional chemical thruster and positively solved. The amount of propellant is also calculated and found to be feasible. The system we discussed in this paper is to be equipped with a newly designed spacecraft planned in 2010s. Copyright IAF/IAA. All rights reserved.

ISAS/JAXA is studying a deployment method using centrifugal force for solar sail mission. The larger sail is required to be deployed statically as for actual spacecraft. We scheduled to deploy the square sail of diameter 20m using a high altitude balloon. In this paper, the mechanisms for static first and second stage deployments are proposed and developed experiment system is introduced. And the operation and results are shown in detail.

宇宙探査の中でも天体への着陸ミッションは目標地点への高精度な航法・誘導が非常に重要である。本研究では2005年に探査機はやぶさが小惑星イトカワに着陸した際に用いられた航法アルゴリズムを基に、事前観測で得られた対象天体モデルを利用し、探査機が撮像した画像を入力として自律的に自己位置を推定するアルゴリズムを考案し、シミュレーションにより精度、ロバスト性等の検証を行った。

ISAS/JAXA is studying a deployment method using centrifugal force for solar sail mission. The larger sail is required to be deployed statically as for actual spacecraft. We scheduled to deploy the square sail of diameter 20m using a high altitude balloon. In this paper, the mechanisms for static first and second stage deployments are proposed and developed experiment system is introduced. And the operation and results are shown in detail.

Reported in this paper is the result of the mission analysis of the asteroid explorer mission. Following the results of HAYABUSA, the Japanese asteroid explorer, JAXA has started the study of the next asteroid exploration mission. The mission now under study gives priority on "early" achievement of the sample return from an asteroid with primitive composition. Therefore, the design of the spacecraft follows that of HAYABUSA basically as it is, and the spacecraft is planned to be launched in early 2010s. The objective of the mission analysis is to design a mission sequence, which has launch window in early 2010s, which is feasible by a HAYABUSA-type spacecraft, and whose target asteroid complies with the science objective. The result includes the selection of the target asteroid, the design of nominal mission sequence, and some back up plans.

The MUSES-C was launched on May 9th of this year and was named 'Hayabusa'. It was aimed to be 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 on board. 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. (c) 2005 Elsevier Ltd. All rights reserved.

The spacecraft Hayabusa, which was launched in 2003, arrived at its destination, asteroid (25143) Itokawa in September 2005. The appearance of Itokawa, a small S-type near Earth asteroids, was totally unexpected. The surface is covered with a lot of boulders and there are only a few craters on it. It looks like a contact binary asteroid. The surface composition is quite similar to LL-chondrite. The estimated density is 1.9 0.13 (g/cm ), so the macro-porosity is about 40%. This means that Itokawa is a rubble pile object. In Itokawa, we may see such things that are very close to building blocks of asteroids. In this paper, we review the mission and the first scientific results. © 2007 International Astronomical Union. 3

Results of the Hayabusa mission to image and sample the asteroid Itokawa are summarized. © 2007 International Astronomical Union.

During the interval from September through early December 2005, the Hayabusa spacecraft was in close proximity to near-Earth asteroid 25143 Itokawa, and a variety of data were taken on its shape, mass, and surface topography as well as its mineralogic and elemental abundances. The asteroid's orthogonal axes are 535, 294, and 209 meters, the mass is 3.51 × 10 kilograms, and the estimated bulk density is 1.9 ± 0.13 grams per cubic centimeter. The correspondence between the smooth areas on the surface (Muses Sea and Sagamihara) and the gravitationally low regions suggests mass movement and an effective resurfacing process by impact jolting. Itokawa is considered to be a rubble-pile body because of its low bulk density, high porosity, boulder-rich appearance, and shape. The existence of very large boulders and pillars suggests an early collisional breakup of a preexisting parent asteroid followed by a re-agglomeration into a rubble-pile object. 10

After global observations of asteroid 25143 Itokawa by the Hayabusa spacecraft, we selected the smooth terrain of the Muses Sea for two touchdowns carried out on 19 and 25 November 2005 UTC for the first asteroid sample collection with an impact sampling mechanism. Here, we report initial findings about geological features, surface condition, regolith grain size, compositional variation, and constraints on the physical properties of this site by using both scientific and housekeeping data during the descent sequence of the first touchdown. Close-up images revealed the first touchdown site as a regolith field densely filled with size-sorted, millimeter- to centimeter-sized grains.

The MUSES-C mission is the world first sample and return attempt to or 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 to 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. An autonomous descent scheme and a novel visual navigation system are proposed and presented in detail. The validity and the effectiveness of the proposed methods are confirmed and evaluated by numerical simulations and flight results.

The Japanese asteroid exploration spacecraft Hayabusa autonomously performed touchdown two times in November 2005. The autonomous guidance and navigation capability is installed aboard the spacecraft. The GNC system collects the laser altimeter, laser range finders and navigation cameras information aboard and is designed to estimate where the spacecraft is and to decide the path correction maneuvers. The programmed function includes the image processing designed to detect an artificial target marker location to approach and cancel the relative velocity. A terrain alignment maneuver is also accomplished by both altitude and attitude control. This paper presents how autonomous guidance and navigation was performed in Hayabusa mission.

The dynamical environment about and on Asteroid 25143 Itokawa is studied using the shape and rotation state model estimated during the close proximity phase of the Hayabusa mission to that asteroid. We first discuss the general gravitational properties of the shape model assuming a constant density. Next we discuss the actual dynamical environment about this body, both on the surface and in orbit, and consider the orbital dynamics of a Hayabusa-like spacecraft. Then we detail one of the approaches used to estimate the mass of the body, using optical and lidar imaging, during the close proximity phase.

The electron cyclotron resonance ion engine, "μ10,"" has a long life and high reliability because of electrodeless plasma generation in both the ion generator and the neutralizer. Four μ10, each generating a thrust of 8 mN, specific impulse of 3,200 seconds, and consuming 350 W of electric power, propel the "HAYABUSA" asteroid explorer that was launched on May 2003. After vacuum exposure and several runs of bailing to reduce residual gas, the ion engine system established continuous acceleration. Delta-V Earth Gravity Assist, a new orbit change scheme that uses electric propulsion with a high specific impulse was applied to change from a terrestrial orbit to an asteroid-based orbit. In 2005, HAYABUSA, using solar electric propulsion, managed to successfully cover the distance between 0.86 AU and 1.7 AU in the solar system, as well as rendezvous with, land on, and lift off from the asteroid Itokawa. During the 3-year flight, the ion engine system generated a delta-V of 1,400 m/s while consuming 22 kg of xenon propellant and operating for 25,900 hours.

A large scale system including multi-elements that turn on and off their electric circuits such as heaters has some high risk of running short of the resource, that is power in this case, and may sometimes go down. The power consumed may exceed the capacity of the power supply, when a lot of the elemental heaters are switched on at the same instance. On average, even in that case, the power consumed may be enough low for the power to be supplied with adequate margin. If a very sophisticated server is built and concentrates allocating resource to the elements keeping the peak power consumed in the whole system lowered by gathering information from them, this shutdown may be circumvented. However, such specific server, even though built, is very much tuned to a specific system structure and not flexible, while the system had better be easily reconfigured and be robust for small mull functions that could occur. The paper here presents a novel multi-agent heater controller that maintains the temperature at each element, while the whole power consumed is kept constant, without any specific server in the system.

At the Institute of Space and Astronautical Science(ISAS),Japan Aerospace Exploration Agency(JAXA), Spinning Solar Sail-craft is proposed and examined the possibility. In August 2006, a membrane of 20[m] in diameter was deployed statically by using a flying balloon. The size of membrane is the largest in the world, furthermore, to deploy statically is the first achievement in the world. This paper shows the experiment apparatus and the results. Additionally, numerical simulation is used to analyze the behavior of the membrane. The deployment of membrane is divided first stage and second stage. At first stage deployment, the membrane deploys dynamically because the accident occurs. However, the deployment is successful. At second stage deployment, the membrane deploys statically and finally forms square-shape. The numerical simulation accords with the behavior of membrane. The membrane has the inclination by gravity, so the experiment can examine the out-of-motion of the membrane. And the membrane is affected by air drag, the contortion of the membrane is also examined. As a conclusion, to deploy a membrane of 20[m] in diameter statically is the first achievement in the world, and to analyze the behavior of the membrane is meaningful for the realization of Solar Sail-craft.

This paper shows the dynamic properties interpreted about a falling cat motion. A new interpretation on non-holonomic turns was presented from coning effect point of view. It 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 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.

Japan Aerospace Exploration Agency (JAXA) is currently studying "Solar Sail" populsion for future deep space explorations. One of the key technologies to realize a solar sail is how light and how compact we can make the photon acceptance surface. JAXA has conducted extensive studies on utlizing centrifugal force to deploy the photon acceptance surface. The final objective is to realize a 7.5μm-thickness and 50m diameter polyimide membrane, combined with a thin flexible solar cells, as the photon acceptance surface that will be needed around the Jupiter orbit. In August 9, 2004, JAXA launched the S-310 sounding rocket, which tested two different shapes of membranes during the zero-gravity flight. The first type of membrane looked like a "clover-leaf", and another is like a "fan". These two membranes, both of them having 10m diameter, were unfolded sequentially during the zero-gravity 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. © 2006 Lister Science. All rights reserved.