川勝 康弘

A study on the post-HINODE Solar Observation Mission has been started in the Solar physics community. One candidate of the mission targets on the observation of the Solar polar region from the orbit largely inclined with the ecliptic plane. Following the paper presented in the previous meeting, this paper focuses on the mission design option which uses the Solar electric propulsion. The design is refined and optimized based on the further analyses on the sensitivity to parameters and the additional consideration of practical constraints. The newly obtained mission baseline and its features are discussed as well.

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.

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.

Reported in this paper are the results of the orbit maneuver compensation in KAGUYA's Lunar transfer. Because of the uncoupled allocation of the attitude control thrusters, extra velocity increment (δv ) is induced whenever KAGUYA performs an orbit maneuver. Since the observed level of δv was unacceptable range from the point of maneuver accuracy requirement, it was compensated by means of deducting estimated δv from the orbit maneuver command. The δv estimation model was updated step-by-step during the Lunar transfer, which leaded to significant improvement of the orbit maneuver accuracy and resulted in the omission of the last trajectory correction maneuver. The method of the compensation and its results are introduced in detail.

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 study on the post- HINODE Solar Observation Mission has been started by members in the Solar physics community. One candidate of the mission targets on the observation of the high latitude region of the Sun, which requires the injection of the space observatory (spacecraft) into the orbit largely inclined with the ecliptic plane. Reported in this paper are the trajectory design results for this orbit transfer, which contains a sequential application of the Electric Propulsion Delta-V Earth Gravity Assist (EDVEGA) procedure.

The study on the post- HINODE Solar Observation Mission has been started by members in the Solar physics community. One candidate of the mission targets on the observation of the high latitude region of the Sun, which requires the injection of the space observatory (spacecraft) into the orbit largely inclined with the ecliptic plane. Reported in this paper are the trajectory design results for this orbit transfer, which contains a sequential application of the Electric Propulsion Delta-V Earth Gravity Assist (EDVEGA) procedure.

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.

KAGUYA (SELENE), launched on September 14, 2007, is Japan's first full-scale lunar explorer. The primary objectives of KAGUYA are the acquisition of scientific data related to the lunar origin and evolution and the development of technologies for future lunar exploration. A major challenge for the KAGUYA mission is reliable and robust lunar orbit insertion under various constraints. After intensive discussions and tradeoff studies, we ultimately selected a phasing lunar transfer orbit with two loops as the KAGUYA lunar transfer orbit. We developed a method for parameterized rocket trajectory by the celestial declination of the Moon to inject KAGUYA into the lunar transfer orbit and designed KAGUYA's phasing lunar transfer orbit and orbital maneuver plan to be robust against anomalies. This paper presents the KAGUYA's trajectory design and orbital maneuver plan, describes KAGUYA's orbital and attitude control system, and explains KAGUYA's flight results during lunar transfer orbit and lunar orbit injection. © 2008 by the American Institute of Aeronautics and Astronautics, Inc.

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.

In this paper, the reconfiguration of translunar trajectory in case of main engine anomaly is investigated. The objectives of the trajectory design are to reduce the excessive velocity at the Lunar encounter, as well as to reduce the total required δv to complete the sequence. 3-impulse Hohmann transfer based trajectory is adopted and possible trajectories are categorized under 2-body approximation. The solutions obtained are applied to more sophisticated models (3-body and 4-body approximation) and yields feasible trajectories.

In this paper, the reconfiguration of translunar trajectory in case of main engine anomaly is investigated. The objectives of the trajectory design are to reduce the excessive velocity at the Lunar encounter, as well as to reduce the total required Delta v to complete the sequence. 3-impulse Hohmann transfer based trajectory is adopted and possible trajectories are categorized under 2-body approximation. The solutions obtained are applied to more sophisticated models (3-body and 4-body approximation) and yields feasible trajectories.

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.

Reported in this paper is the initial phase operation of the attitude and orbit control system (AOCS) of Japanese satellite "AKARI". AKARI is the first Japanese satellite dedicated to the infrared astronomy. AKARI was successfully launched by M-V rocket from Kagoshima Space Center on February 22, 2006. Just after the launch, AKARI faced to a serious problem. It was found that something interferes with the fields of view of two (out of two) sun sensors. The two sensors were out of use in the subsequent AOCS operation. Fortunately, by using two star trackers and gyros, the attitude accuracy required for the scientific observation can be achieved without the sun sensors. However, as other many satellites, the sun sensors were to play important roles in AKARI's AOCS operation. Firstly, they were to be used for sun acquisition at the very initial phase. The anomaly prevented us from carrying out the pre-planned sequence. However, the crisis was overcome with the appropriate ground support operation. Secondly, for the observation in far infrared wavelength, the telescope and the scientific instruments of AKARI are stored in the cryostat and cooled by liquid Helium. To prevent the sun light inflow to the telescope, the attitude of AKARI against the sun is strictly constrained after the scientific observation starts. The sun sensors were to be used to watch the sun direction to keep the appropriate attitude against the sun even when AKARI falls into the safe mode. To save the cryostat from the sun light in the absence of the sun sensors, AOCS architecture was reconfigured so that the equivalent function is achieved by the remaining sensors. The experiences in this recovery operation are mainly reported in this paper. In addition, unexpected continuous orbit rising was observed in the operation. The cause of the phenomenon and the investigation process are also reported.

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.

Reported in this paper is the initial phase operation of the attitude and orbit control system (AOCS) of Japanese satellite "AKARI". AKARI is the first Japanese satellite dedicated to the infrared astronomy. AKARI was successfully launched by M-V rocket from Kagoshima Space Center on February 22, 2006. Just after the launch, AKARI faced to a serious problem. It was found that something interferes with the fields of view of two (out of two) sun sensors. The two sensors were out of use in the subsequent AOCS operation. Fortunately, by using two star trackers and gyros, the attitude accuracy required for the scientific observation can be achieved without the sun sensors. However, as other many satellites, the sun sensors were to play important roles in AKARI's AOCS operation. Firstly, they were to be used for sun acquisition at the very initial phase. The anomaly prevented us from carrying out the pre-planned sequence. However, the crisis was overcome with the appropriate ground support operation. Secondly, for the observation in far infrared wavelength, the telescope. and the scientific instruments of AKARI are stored in the cryostat and cooled by liquid Helium. To prevent the sun light inflow to the telescope, the attitude of AKARI against the sun is strictly constrained after the scientific observation starts. The sun sensors were to be used to watch the sun direction to keep the appropriate attitude against the sun even when AKARI falls into the safe mode. To save the cryostat from the sun light in the absence of the sun sensors, AOCS architecture was reconfigured so that the equivalent function is achieved by the remaining sensors. The experiences in this recovery operation are mainly reported in this paper. In addition, unexpected continuous orbit rising was observed in the operation. The cause of the phenomenon and the investigation process are also reported.

In this paper, the concept of Orbit Transfer Vehicle for Deep Space Exploration (Deep Space OTV) is proposed, and its effectiveness and feasibility are discussed. Basic concept is the separation of two functions required for the deep space exploration, the transportation to the destination, and the exploration at the destination. Deep Space OTV is a spacecraft specialized for the transportation to the deep space destination. It is an expendable spacecraft propelled by solar electric propulsion. The payload of Deep Space OTV is Explorer, which is a spacecraft specialized for the exploration at the deep space destination. The effectiveness of the concept is discussed qualitatively, focused on the merits of the separations of two functions. The feasibility of Deep Space OTV is discussed based on the conceptual design of the spacecraft and its applicability to deep space missions. Several deep space missions are modeled and the payload capacity of Deep Space OTV is estimated. The missions include Asteroid rendezvous, Mars orbiter, Lunar lander, and so on. © 2007 Elsevier Ltd. All rights reserved.

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.

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.

We introduce a Japanese plan of infrared(z-band:0.9μm) space astrometry(JASMINE-project). JASMINE is the satellite (Japan Astrometry Satellite Mission for INfrared Exploration) which will measure distances and apparent motions of stars around the center of the Milky Way with yet unprecedented precision. It will measure parallaxes, positions with the accuracy of 10 micro-arcsec and proper motions with the accuracy of ∼ 4micro-arcsec/year for stars brighter than z=14mag. JASMINE can observe about ten million stars belonging to the bulge components of our Galaxy, which are hidden by the interstellar dust extinction in optical bands. Number of stars with ω/π < 0.1 in the direction of the Galactic central bulge is about 1000 times larger than those observed in optical bands, where π is a parallax and σ is an error of the parallax. With the completely new quot;map of the bulge in the Milky Way", it is expected that many new exciting scientific results will be obtained in various fields of astronomy. Presently. JASMINE is in a development phase, with a target launch date around 2015. We adopt the following instrument design of JASMINE in order to get the accurate positions of many stars. A 3-mirrors optical system(moclified Korsch system)with a primary mirror of ∼ 0.85m is one of the candidate for the optical system. On the astro-focal plane, we put dozens of new type of CCDs for z-band to get a wide field of view. The accurate measurements of the astrometric parameters requires the instrument line-of-sight highly stability and the opto-mechanical highly stability of the payload in the JASMINE spacecraft. The consideration of overall system(bus) design is now going on in cooperation with Japan Aerospace Exploration Agency(JAXA).

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.

In this paper, the concept of Orbit Transfer Vehicle for Deep Space Exploration (Deep Space OTV) is proposed, and its effectiveness and feasibility are discussed. Basic concept is the separation of two functions required for the deep space exploration, the transportation to the destination, and the exploration at the destination. Deep Space OTV is a spacecraft specialized for the transportation to the deep space destination. It is an expendable spacecraft propelled by solar electric propulsion. The payload of Deep Space OTV is Explorer, which is a spacecraft specialized for the exploration at the deep space destination. The effectiveness of the concept is discussed qualitatively, focused on the merits of the separations of two functions. The feasibility of Deep Space OTV is discussed based on the conceptual design of the spacecraft and its applicability to deep space missions. Several deep space missions are modeled and the payload capacity of Deep Space OTV is estimated. The missions include asteroid rendezvous, Mars orbiter, lunar lander, and so on.

Sample return missions to multiple near-Earth asteroids are investigated assuming solar electric propulsion as propulsive means. The aim of the mission is to collect the soil samples from multiple asteroids and to return them to the Earth. The objective of the trajectory design is not only to maximize the number of the target asteroids and the payload mass of the spacecraft, but also to make the flight time up to the Earth return as short as possible for quick science return. Two options are studied for the orbital sequence. The first scenario introduces Earth gravity-assist after each asteroid rendezvous in order to return the reentry capsule as fast as possible as well as to reflect the result of the sample analysis for the following mission scenario. The other trajectory option is designed without Earth gravity assists.

In this paper, translunar trajectory of the simple sequence is investigated. The trajectory discussed in this paper is two-burn ballistic trajectory from the low earth parking orbit to the low lunar orbit. The problem is practical in that the geometric relation from the launch site to the moon is fully considered. The paper includes three subjects. The first is the structure of the problem and the solution space. The problems is defined and the solutions are grouped by the properties of lunar transfer sequence. The second is the characteristics of the optimal solutions. The topics discussed are, the transition of the required velocity increment by the launch date, and the difference by the property of the lunar transfer sequence. The third is the sensitivity analysis of a number of items to the deviation of the parameters from the optimal solution.

We have been studying a large membrane space structure named "Furoshiki Satellite," as a promising candidate of a future space system for those missions requiring large area in space such as solar power generation, a large communication antenna, or a large radiator. This membrane is folded in a very small volume during launch and is deployed and controlled by a set of several satellites at its corners or using centrifugal force generated by rotating the central satellite. It is expected that such a structure will reduce the weight per area of the space structure and, if the control technology is properly applied, it can be efficiently folded during launch and easily deployed after release. This paper shows the concept of Furoshiki Satellite, its applications, and its dynamics on orbit and how to control it. A nano-satellite project to demonstrate the concept of Furoshiki Satellite will also be described briefly. © 2001 International Astronautical Federation. Published by Elsevier Science L td.

This paper discusses on minimum-fuel trajectories for interplanetary explorers. The explorers are accelerated by solar electric propulsion(SEP) system with minimum fuel consumption. SEP has higher specific impulse than chemical propulsion, but long firing time is necessary because of low thrust. According to optimal control theory, the optimal thruster control is bang-bang control. Thus the problem is to find the switching times and the direction of thrust vector. A new transformation for numerical optimization is introduced to obtain the trajectories with high accuracy. The results show that a coasting arc saves the fuel and the fuel consumption depends on only the specific impulse of the propulsion system. © 2000 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

This paper focuses on the three topics related to the trajectory design of SELENE. The first is the orbit maneuver of the orbiter. The altitude of the orbiter is 100km and the orbit is strongly perturbed by high order term of the gravity potential. In order to satisfy the mission requirements, ten maneuvers are scheduled during one year mission. The second topic is the orbit design of relay satellite. The relay satellite has no propulsion system and has no orbit maneuver capability. The orbit of the relay satellite is perturbed mainly by earth's gravity and the shape of the orbit changes through the one year mission. The initial orbit is selected carefully to meet the mission requirements through the mission considering the effect of perturbation. The third topic is the trajectory design of the landing mission. The navigation error in the landing phase is expected to be large value. Main reason of the error is orbit determination error and long duration of inertial navigation. The landing trajectory is designed to permit this navigation error and assure the safe landing.

The orbital transfer operation utilizing the OTV (orbital transfer vehicle). equipped with a tether is proposed. There have been many ideas of the transportation using the tether since 1960's. Most of them simply apply the gravity gradient effect to change the orbit of the payload attached at the tip of the tether. However, it is clear that a larger orbital change can be achieved by utilizing a librating or rotating tether. In this paper, the basic model of the mission using the rotating tether is introduced. The methods to rotate the tether are explained and their effectivenesses are verified by numerical simulations. At the same time, the fuel consumption for typical orbital transfer mission is studied. The results indicate that this type of OTV can reduce the fuel consumption by 45% maximum compared with conventional nontethered OTV system.