川口 淳一郎

The Hayabusa spacecraft successfully recovered dust particles from the surface of near-Earth asteroid 25143 Itokawa. Synchrotron-radiation x-ray diffraction and transmission and scanning electron microscope analyses indicate that the mineralogy and mineral chemistry of the Itokawa dust particles are identical to those of thermally metamorphosed LL chondrites, consistent with spectroscopic observations made from Earth and by the Hayabusa spacecraft. Our results directly demonstrate that ordinary chondrites, the most abundant meteorites found on Earth, come from S-type asteroids. Mineral chemistry indicates that the majority of regolith surface particles suffered long-term thermal annealing and subsequent impact shock, suggesting that Itokawa is an asteroid made of reassembled pieces of the interior portions of a once larger asteroid.

Noble gas isotopes were measured in three rocky grains from asteroid Itokawa to elucidate a history of irradiation from cosmic rays and solar wind on its surface. Large amounts of solar helium (He), neon (Ne), and argon (Ar) trapped in various depths in the grains were observed, which can be explained by multiple implantations of solar wind particles into the grains, combined with preferential He loss caused by frictional wear of space-weathered rims on the grains. Short residence time of less than 8 million years was implied for the grains by an estimate on cosmic-ray-produced (21)Ne. Our results suggest that Itokawa is continuously losing its surface materials into space at a rate of tens of centimeters per million years. The lifetime of Itokawa should be much shorter than the age of our solar system.

Japan Exploration Agency (JAXA) launched a powered solar sail "Interplanetary Kite-craft Accelerated by. Radiation Of the Sun (IKAROS)" on May 21, 2010. One of the primal technologies demonstrated at IKAROS is the deployment of the sail whose diameter is 20m class. After the launch, the deployment operation was performed and successful expansion of the sail was confirmed. The deployment sequence in IKAROS consists of static first stage and dynamic second stage. In this paper, the flight data and observed dynamic motion during the first stage deployment are reported. These are compared with the results of numerical simulations using multi-particle model, and the accuracy and availability of this model is discussed. Copyright ©2011 by the International Astronautical Federation. All rights reserved.

工学実証を目的に2003年5月に打ち上げられた「はやぶさ」は,小惑星イトカワへの着陸を経て,2010年6月13日,オーストラリアに帰還し,試料カプセルを無事回収することに成功した.本報告は,その帰還,再突入における飛行運用を紹介する.

小惑星探査機「はやぶさ」は2010年6月13日にオーストラリアに帰還し,試料カプセルを無事回収することに成功した.これに先立ち2010年4月以降,合計5回の軌道制御により,地球をフライバイする軌道からオーストラリアに帰還する軌道へと精密誘導を実施した.本報告は,誘導計画立案に新たに導入したパラメータを紹介し,このパラメータを用いた計画立案手法について解説する.

The Hayabusa spacecraft aiming at technology demonstration for the world's first sample and return from an extra-terrestrial object was launched by the fifth M-V rocket from Uchinoura Space Center, JAXA on May 9, 2003. It went through several troubles and hardships during totally 7 years of interplanetary flight. It successfully returned to the earth and completed the powered- flight by the ion thruster in the beginning of 2010. After successive trajectory correction maneuvers for the reentry, the mother spacecraft, Hayabusa successfully released a small sample-return capsule with asteroid Itokawa sample contained in the sample canister aboard. The capsule entered the earth atmosphere in the desert of Australia on June 13, 2010, and was successfully recovered by June 15. This paper overviews the return operation of the Hayabusa mother spacecraft and reentry flight and recovery operation of the sample return capsule. And the paper also will provide how these TCM and EDL activities were performed and some associated information regarding the Hayabusa mission.

The Hayabusa spacecraft aiming at technology demonstration for the world's first sample and return from an extra-terrestrial object was launched by the fifth M-V rocket from Uchinoura Space Center, JAXA on May 9, 2003. It went through several troubles and hardships during totally 7 years of interplanetary flight. It successfully returned to the earth and completed the powered-flight by the ion thruster in the beginning of 2010. After successive trajectory correction maneuvers for the reentry, the mother spacecraft, Hayabusa successfully released a small sample-return capsule with asteroid Itokawa sample contained in the sample canister aboard. The capsule entered the earth atmosphere in the desert of Australia on June 13, 2010, and was successfully recovered by June 15. This paper overviews the return operation of the Hayabusa mother spacecraft and reentry flight and recovery operation of the sample return capsule. And the paper also will provide how these TCM and EDL activities were performed and some associated information regarding the Hayabusa mission.

Orbit determination is the process to estimate the position and velocity from a given set of observations of spacecraft. The common measurements used in orbit determination are range and range rate. For precise orbit determination, the acceleration of spacecraft by solar radiation pressure (SRP) must be taken into account. In order to incorporate the SRP acceleration, the attitude history and the SRP model - surface shape and optical surface feature of spacecraft - need to be known beforehand. Especially, in case of solar sail spacecraft, since their SRP effect is strong, coupling between attitude and orbital dynamics is more remarkable than ordinary ballistic missions. Therefore, by positively utilizing this coupling, it is possible to refine the attitude information based on the residual errors of orbit determination. In this study, we assume that the systematic errors in the orbit determination of the spacecraft are caused solely by the errors in attitude information, and propose an attitude error correction method which utilizes the coupling between attitude and orbital dynamics. In order to examine the validity of the method, the analysis of actual data is conducted. As actual data, flight data of IKAROS - Japanese interplanetary solar sail demonstration spacecraft - are used. The observations used in the orbit determination of IKAORS are range and range rate. In this paper, an estimator is designed whose input is the residuals of the orbit determination result - the difference between actual measurements and modelled observations of range and rage rate - and whose output is the amount of error correction in the attitude history. It is confirmed that orbit and attitude errors are both decreased through this method. This paper explains the theoretical background of this method and shows the result of analysis using the actual data. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

Japan Exploration Agency (JAXA) launched a powered solar sail "Interplanetary Kite-craft Accelerated by Radiation Of the Sun (IKAROS)" on May 21, 2010. One of the primal technologies demonstrated at IKAROS is the deployment of the sail whose diameter is 20m class. After the launch, IKAROS performed the deployment sequence and have confirmed that the membrane was successfully expanded. In this paper, the flight data and observed dynamic motion via deployment are reported. These are compared with the results of numerical simulations using multi-particle model, and the accuracy and availability of this model is discussed. Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc.

JAXA launched the world's first deep space solar sail demonstration spacecraft "IKAROS" on May 21, 2010. IKAROS was injected to an Earth-Venus trajectory to demonstrate several key technologies for solar sail utilizing the deep space flight environment. IKAROS succeeded in deploying a 20 m-span solar sail on June 9, and is now flying towards the Venus with the assist of solar photon acceleration. This paper describes the mission design, system design, solar sail deployment operation and current flight status of IKAROS. © 2011 Elsevier Ltd. All rights reserved.

In this paper, the attitude dynamics of IKAROS, which is spinning solar sail, is presented. Multi Particle Model (MPM) and First Mode Model of out-of-plane deformation (FMM) are introduced to analyze the out-of-plane oscillation mode of spinning solar sail. Considering the thruster configuration of IKAROS, the force on main body and membrane by thruster plume as well as reaction force by thruster are integrated into MPM. The attitude motion after sail deployment or reorientation using thrusters can be analyzed by MPM numerical simulations precisely. The out-of-plane oscillation of IKAROS is governed by three modes derived from FMM. FMM is simple and valid for the design of attitude controller. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

This note discusses the stability of collinear equilibrium points around a rotating system composed of two masses rigidly connected by a massless rod in the case, where the centripetal force outweighs the gravitational force. It is found that a stable region appears at L(1) when the ratio of gravitational to centripetal acceleration is less than 0.125, and that there is always no stable area at L(2) and L(3); the result is applied to the fast rotating Asteroid 2000EB(14). (C) 2009 Elsevier Inc. All rights reserved.

Hayabusa Sample Return Capsule (SRC) has accomplished reentry flight directly from the interplanetary-transfer orbit with velocity of about 12 km/s and successfully recovered on June 14, 2010. The SRC is considered to deploy parachute at the altitude of 5 km as planned, while the chute-trigger timer has been carefully set to cover contingency cases due to malfunctions of triggering devises taking account of dispersions of orbital and atmospheric density and errors in the aerodynamic coefficients. All the components separated at the chute-deployment were searched-out based on the landing position of the instrument module.

The Japan Aerospace Exploration Agency (JAXA) has been developing lithium-ion secondary cells for aerospace applications. Lithium-ion secondary cells were first used for the interplanetary spacecraft, HAYABUSA. For the long-term operations on the moon and interplanetary orbit, in-orbit performance of the lithium-ion battery demonstrated by HAYABUSA has been precisely examined. Furthermore, the regenerative fuel cell system is also one of the most important candidates to the future lunar and planetary missions where the higher energy density is required.

In this paper, we consider constellation of two orbiters around the Mars. We assume two orbiters, whose orbits are required to be orthogonal. After Mars orbit insertion, it will be needed to adjust their orbital planes appropriately for required configuration under perturbation. We discuss how to transfer S/C to the seperated orbits after simultaneous insertion, and how to maintain these orbits during the mission phase. We adopt the frozen orbit for one orbiter to fix the orbit axis, and utilize J2 perturbation on the argument of periapsis or the right ascension of ascending node for keeping the orthogonality between two orbital planes. Maneuver plan is also reported.

Preliminary feasibility assessment of a Martian nonstop sample return system has been conducted as a part of Mars Exploration with Landers and Orbiters (MELOS) which is currently under investigation in Japan Aerospace Exploration Agency. In the mission scenario, an atmospheric entry vehicle of aero-maneuver ability is flown into the Martian atmosphere, collects the Martian dust particles and atmospheric gases during the hypersonic atmospheric flight, exits the Martian atmosphere, and is inserted into a parking orbit from which a return system departs for the earth. In order to accomplish controlled flight and successful orbit insertion, aerocapture technologies are introduced into the vehicle guidance and control system. A conceptual design is presented as a result of the preliminary system analysis.

'Hayabusa' re-entry capsule was safely carried into the clean room of Sagamihara Planetary Sample Curation Facility in JAXA. After executing CT scanning, removal of heat shield, and surface cleaning of sample container, the sample container was enclosed into the clean chamber. After opening the sample container and residual gas sampling in the clean chamber, optical observation, sample recovery, sample separation for initial analysis is carried out. This curation work is continued with some selected member of Hayabusa Asteroidal Sample Preliminary Examination Team (HASPET) for several months. We report the 'Hayabusa' capsule recovery operation and current status of curation work for 'Hayabusa' sample.

IKAROS (Interplanetary Kite-craft Accelerated by Radiation Of the Sun) that was a demonstrator of the critical technologies of the solar power sail was launched on May 2010 from Tanegashima Space Center using H-IIA rocket. The solar power sail utilizes two kinds of forces; one by solar photon from sun and another from high performance ion engine. Minimum mission success was the deployment of the membrane and the generation of the electricity by means of the thin film solar array on the membrane within several weeks. Also, acceleration and navigation using the solar sail will be demonstrated within half year. This paper describes the structure, the space environmental evaluation of the thin film solar array and the initial results of the solar array installed on the IKAROS in the orbit. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

This paper discusses the transportation architecture for the round-trip exploration to Mars taking advantage of the combination of the "Deep Space Port" and "low-thrust propulsion." One of the merits of utilizing the Deep Space Port is the reusability of the interplanetary portion of the architecture. The main objective in this study is to verify the connectivity of the escape trajectory from Sun-Earth liberation point with the interplanetary trajectory. The paper divides the scenario into two segments; near Earth phase and interplanetary phase. Both ballistic trajectories and low-thrust trajectories are assessed. The low-thrust optimization is performed by DCNLP (Direct Collocation with Nonlinear Programming) which approximates the continuous problem with parameter optimization problem and provides precise results by simple formulation. This paper shows the connectivity of the escape trajectory from Sun-Earth liberation point with the interplanetary trajectory, and shows the usefulness for the round-trip exploration to Mars using Deep Space Port in combination with low thrust propulsion. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

The cathode-less electron cyclotron resonance ion engines, μ10, propelled the Hayabusa asteroid explorer, launched in May 2003, which is focused on demonstrating the technology needed for a sample return from an asteroid, using electric propulsion, optical navigation, material sampling in a zero gravity field, and direct re-entry from a heliocentric orbit. It rendezvoused with the asteroid Itokawa after a two year deep space flight with a delta-V of 1.4 km/s, 22 kg of xenon propellant consumption and 25800 hours of total accumulated operational time of all the four ion engines added up. Though it succeeded in landing on the asteroid on November 2005, the spacecraft was seriously damaged. This delayed the Earth return in 2010 from the original plan in 2007. Reconstruction on the operational scheme using remaining functions and newly uploaded control logic made Hayabusa leave for Earth in April 2007. After a coasting period of 2008, the ion propulsion was reignited in February 2009. Although most of the neutralizers were degraded and unable to be used by fall of 2009, a combination of an ion source and its neighboring neutralizer has been successfully operated for the last 3230 hours including a series of final trajectory correction maneuvers. Before reentry, the total accumulated operational time reached 39637 hours consuming a total of 47 kg Xenon propellant. Total duration of powered spaceflight is 25590 hours which provided a delta-V of 2.2 km/s and a total impulse of 1 MN·s, approximately. Finally, the spacecraft returned to Earth. Its reentry capsule, which may contain samples from asteroid Itokawa, was retrieved from the Australian outback according to plan. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

Japan Aerospace Exploration Agency (JAXA) has developed the small demonstration solar sail spacecraft IKAROS (Interplanetary Kite-craft Accelerated by Radiation of the Sun), which will be launched in mid 2010. The main objective of this spacecraft is to deploy the 20m class sail membrane, and demonstrate the acceleration of a spacecraft by the solar radiation pressure (SRP) by means of that sail. It is important to model the SRP force adequately for the objective of navigation, especially for interplanetary spacecraft. In order to improve the model of the SRP torque induced by the sail membrane, the IKAROS project team plans to estimate the SRP torque parameters in orbit. In this paper, we present the approach to obtain the parameters needed for constructing the photon torque model through the analysis of the attitude dynamics.

JAXA launched the world's first deep-space solar sail demonstration spacecraft "IKAROS" on May 21, 2010. IKAROS was injected to an Earth-Venus trajectory to demonstrate several key technologies for solar sail utilizing the deep space flight environment. IKAROS succeeded in deploying a 20m-span solar sail on June 9, and is now flying toward Venus with the assist of solar photon acceleration. This paper describes the mission design, system design, solar sail deployment operation and current flight status of IKAROS. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

This paper presents new orbital synthesis results to achieve ballistic and short period out-of ecliptic trajectories, instead of using electric propulsion or solar sail acceleration. The strategy developed utilizes a Jovian gravity first, followed by polar very high speed gravity assists by Earth or Venus. So far, the use of very high speed gravity assists has been conceived not practically useful. However, this paper presents those still effectively contribute to amending the trajectories periods, and to acquiring small sized out-of-ecliptic ballistic trajectories. The biggest advantage of this strategy is to reduce propellant mass carried drastically.

This paper presents new orbital synthesis results to achieve ballistic and short period out-of ecliptic trajectories, instead of using electric propulsion or solar sail acceleration. The strategy developed utilizes a Jovian gravity first, followed by polar very high speed gravity assists by Earth or Venus. So far, the use of very high speed gravity assists has been conceived not practically useful. However, this paper presents those still effectively contribute to amending the trajectories periods, and to acquiring small sized out-of-ecliptic ballistic trajectories. The biggest advantage of this strategy is to reduce propellant mass carried drastically.

We propose a novel strategy for precise landing of a space probe on an asteroid surface. In Hayabusa mission, a Japanese asteroid sample-return spacecraft launched in 2003, a small marker was dropped onto the ground and utilized as a landmark for landing. In this paper, we propose two scenarios to drop multiple markers one by one, with a view to more precise landing. Preliminary Monte Carlo simulation of guidance using multiple markers with assumption of some control errors are presented. It was indicated that more than one markers allow us to increase the landing accuracy compared to the conventional method using a single marker.

In this paper, we investigated how to establish and maintain orthogonal constellation of two orbiters around a planet, taking Mars as an example. We assumed a mission where two spacecraft (S/C) will be injected into the Mars orbit simultaneously. One S/C (MOA) has a low altitude orbit, while the other (MOB) has a highly elliptical orbit. It is assumed that MOB should look down upon MOA's orbital plane at MOB's apoapsis; i.e., the two orbits should be preferably orthogonal. Such constellation can be actually required in scientific missions featuring both global and in-situ observation of Mars atmosphere, and it can also be extended to application in other celestial bodies. How to keep the orthogonal constellation under perturbations or other restrictions, and how to establish it with small fuel consumption is the main topic of this paper. © 2010 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

This paper describes how to establish and maintain an orthogonal constellation by two orbiters. We assume the constellation where one orbiter looks down upon the orbital plane of another from the apoapsis, or the normal vector of one orbiter and the eccentricity vector of another are parallel. There have been demands and requests for such orthogonal constellation composed by an in-situ observation orbiter and a remote-sensing orbiter in studies of planetary atmosphere. Since their orbits can be perturbed by several external forces, they must be designed carefully so as to keep orthogonality. In this paper, we introduced the "Orthogonality Index" by using orbital elements of two orbiters, which provides us a quantitative index for orthogonality of the two orbits. By differentiating the index with respect to time, we derived several conditions to keep orthogonality via a bottom-up approach. They were well coincident with the top-down results derived in our former approach. Four solutions for the inclination of the remote-sensing orbiter were derived, which is independent from orbit shapes or central celestial bodies. If we assume planets with a relatively large J2 term such as the earth or the mars, the orbital elements for the in-situ orbiter can also be calculated easily. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

Japan has launched many interplanetary spacecraft for exploration of solar system bodies including Mars. Now we are planning the next Mars mission in the late 2010's. This paper describes the preliminary mission analysis and orbit design for this plan. The combined exploration by several spacecraft requires complicated and careful consideration, different from those for single-probe missions. Mission plans to realize required configuration by a single launch and simple simulation results are reported.

The Japan Aerospace Exploration Agency (JAXA) will make the world's first solar power sail craft demonstration of photon propulsion and thin film solar power generation during its interplanetary cruise by IKAROS (Interplanetary Kite-craft Accelerated by Radiation Of the Sun). The spacecraft deploys and spans a membrane of 20 meters in diameter taking the advantage of the spin centrifugal force. The spacecraft weighs approximately 310kg, launched together with the agency's Venus Climate Orbiter, AKATSUKI in May 2010. This will be the first actual solar sail flying an interplanetary voyage.

This paper introduces new attitude control system for solar sail which leverages solar radiation pressure and achieves completely fuel-free and oscillation-free attitude control of flexible spinning solar sail. Novel attitude control device was developed, which is a thin film-type device and can electrically control its optical parameters such as reflectivity to generate unbalance of the solar radiation pressure applied to the edge of the sail. By using this device, minute and continuous control torque can be applied to the sail so that very stable and fuel-free attitude control of large and flexible membrane is realized. The control system was implemented as an optional attitude control system for small solar power sail demonstrator IKAROS. On-orbit attitude control experiments were conducted and the performance of the controller was successfully verified compared with the ground-based analytical performance estimation. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

In this paper, the attitude motion and attitude control strategy of spinning solar sail are discussed. As the spinning type solar sail does not have any rigid structure to support its membrane, the impulsive torque by the RCS can introduce oscillatory motion of the membrane. Thus, an "oscillation free" attitude controller is needed, which takes into account the flexibility of the membrane and avoid unnecessary oscillatory motion. First, the dynamics model and numerical model were introduced, and the validity of these models and dominant out-of-plane membrane vibration mode is examined by membrane vibration experiment and comparison between both models. Then, based on the analysis of the dynamics of torque-free motion, it was shown that a spinning solar sail has three oscillation modes of nutation, one of which is equal to the spinning rate of the spacecraft. The dominancy of each nutation mode was analytically and numerically discussed. Then, we discussed the spin axis maneuver control using conventional RCS. It was analytically shown that continual impulsive torque synchronizing the spin rate can excite nutation velocity and that a controller is needed to damp the nutation while controlling the spin axis at the same time. The authors proposed new controller named Flex-RLC and improved one. Their effectiveness was verified by numerical simulations using precise multi-particle numerical model which can express higher order oscillatory motion of the flexible membrane, and it was found that the proposed method can control the attitude of spinning solar sail while drastically reduces the nutation velocity compared with conventional control logic. So, it can 983 be said that the proposed method is promising fast and stable controller for spinning solar sail.

Japan Aerospace Exploration Agency (JAXA) has developed the small demonstration solar sail spacecraft IKAROS (Interplanetary Kite-craft Accelerated by Radiation of the Sun), which will be launched in mid 2010. The main objective of this spacecraft is to deploy the 20m class sail membrane, and demonstrate the acceleration of a spacecraft by the solar radiation pressure (SRP) by means of that sail. It is important to model the SRP force adequately for the objective of navigation, especially for interplanetary spacecraft. In order to improve the model of the SRP torque induced by the sail membrane, the MAROS project team plans to estimate the SRP torque parameters in orbit. In this paper, we present the approach to obtain the parameters needed for constructing the photon torque model through the analysis of the attitude dynamics.

The solar sailing spacecraft is one of the promising propulsion systems for the future deep space exploration mission. Japan Aerospace Exploration Agency (JAXA) has been studying the spin solar sail spacecraft which has a squared-shape type solar sail. One of the most significant objective to control the satellite orbit of the spacecraft is to estimate the thrust force induced by the photon, namely to establish the acceleration model before the launch. In a view point to use this model in orbit, the calibration of the acceleration model and evaluation of the dynamics on orbit are also important issue. This paper presents the way to construct the acceleration model of the Solar Radiation Pressure (SRP) on ground, and the calibration and evaluation strategy for this model by using the on-orbit data.

Marco Polo is a mission to return a sample from a near-Earth object of primitive type (class C or D). It is foreseen as a collaborative effort between the Japan Aerospace Exploration Agency (JAXA) and the European Space Agency (ESA). Marco Polo is currently in a Phase-A study (status as of summer 2009). This paper focuses on the scientific requirements provided to the industrial study consortia in Europe as well as the possible mission scenario at the target object in order to achieve the overall mission science objectives. The main scientific reasons for going to a near-Earth object are to...

Reported in this paper are the results of a mission analysis conducted for an asteroid exploration mission. Following the results obtained from HAYABUSA, a Japanese asteroid explorer, the Japan Aerospace Exploration Agency has started studying the possibility of the next asteroid exploration mission. The mission studied gives priority to the early retrieval of a sample from an asteroid with a primitive composition. Therefore, the design of the spacecraft follows that of the HAYABUSA, basically as it is, and the spacecraft is planned to be launched early in the next decade. The objective of the mission analysis is to design a mission sequence which has a launch window early in the next decade, is feasible utilizing a HAYABUSA-type spacecraft, and whose target asteroid complies with the scientific objectives. The results include the selection of the target asteroid, the design of the nominal mission sequence, and the alternative sequence to overcome the drawbacks of the nominal sequence.

The Hayabusa spacecraft launched in 2003 is now completing its round trip to a near Earth asteroid Itokawa, to which it accessed in 2005. Hayabusa suffered from a fuel leak and eruption incidents after its successful descent, touch-down and lift-off at the end of November in 2005. Hayabusa lost two reaction wheels among three aboard and the chemical propulsion. The only means left for Hayabusa are the ion engines and xenon gas reserved for them as well as a single reaction wheel. Hayabusa project team devised the use of xenon gas for cold gas propulsion, and also developed the new attitude control strategy taking the advantage of solar radiation pressure. The spacecraft started the delta-V maneuver using the ion engines from 2007 and will continue it by next February. It is scheduled for Hayabusa to reenter into the atmosphere and land in middle of Australia in June of 2010.

Asteroid probe Hayabusa was launched in May 2003 aiming for sample return from asteroid Itokawa. In the course of the mission, Hayabusa is suffering from failure of attitude control devices such as two of three reaction wheels (RWs) and chemical thrusters. As a means of precaution for an additional problem, more specifically, the last RW trouble an attitude control method which requires no RW should be considered. Ion thrusters can be utilized for attitude control by changing the direction of the thrust by two axis gimbals. In this paper, an attitude control method using only an ion thruster is proposed and the effectiveness is confirmed by numerical simulations.

Dust particles which are observed as zodiacal light exist around the Earth's orbit and can create a characteristic distribution by the Earth's gravity. This paper investigates the distribution by numerical simulation on the assumption that dust particles trapped in horseshoe orbits contribute the distribution, because the orbit has the following characteristics. The velocity of a trapped dust varies by its location on the horseshoe orbit and the dust will transit with time to wider horseshoe orbit. And this paper also considers the possibility that these particles are trapped or not and reveals that the possibility depends on the dust diameter.

This paper presents new orbital synthesis results to achieve ballistic and short period out-of ecliptic trajectories, instead of using electric propulsion or solar sail acceleration. The strategy developed utilizes a Jovian gravity first, followed by polar very high speed gravity assists by Earth or Venus. So far, the use of very high speed gravity assists has been conceived not practically useful. However, this paper presents those still effectively contribute to amending the trajectories periods, and to acquiring small sized out-of-ecliptic ballistic trajectories. The biggest advantage of this strategy is to reduce propellant mass carried drastically.

The Hayabusa spacecraft embarked on its return trajectory to Earth, after its touchdown on the asteroid Itokawa. During the cruise phase Sun-pointing mode, the spin-axis of the Hayabusa performs a coning motion under the solar radiation pressure effects. The effect mainly originates from the diffuse reflection of the solar radiation pressure on the solar array panels. In the simple analysis of this coning motion, however, the diffuse reflection coefficient is inconsistent in comparison to the typical diffusive parameter of the solar array panel. This discrepancy must be clarified by constructing a more accurate model of the solar radiation pressure in order to be able to dissolve the uncertainty during the return phase of the Hayabusa spacecraft. The accurate model should also be able to support the extremely precise navigation during the return phase to the Earth. This paper presents the precise model of the solar radiation pressure of the Hayabusa spacecraft and the estimation method for obtaining the optical parameters of the solar radiation pressure model.

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 315kg, launched together with the agency's Venus Climate Orbiter, PLANET-C in 2010. This will be the first actual solar sail flying an interplanetary voyage.

Recent advances in the characterization of small body surfaces with stereophotoclinometry are & discussed. The principal data output is an ensemble of landmark maps (L-maps), high-resolution topography/albedo maps of varying resolution that tile the surface of the body. Because they can have a resolution comparable to the best images, and can be located on a global reference frame to high accuracy, L-maps provide a significant improvement in discriminatory power for studies of small bodies, ranging from regolith processes to interior structure. These techniques are now being used to map larger bodies such as the Moon and Mercury. L-maps are combined to produce a standard global topography model (GTM) with about 1.57 million vectors and having a wide variety of applications. They can also be combined to produce high-resolution topography maps that describe local areas with much greater detail than the GTM. When combined with nominal predictions from other data sources and available data from other instruments such as LIDAR or RADAR, solutions for the spacecraft position and camera pointing are the most accurate available. Examples are drawn from studies of Phobos, Eros, and Itokawa, including surface characterization, gravity analysis, spacecraft navigation, and incorporation of LIDAR or RADAR data. This work has important implications for potential future missions such as Deep Interior and the level of navigation and science that can be achieved. © The Meteoritical Society, 2008.