川勝 康弘

AKATSUKI, the Japanese Venus explorer, once failed to inject itself into an orbit around Venus in 2010. AKATSUKI is now on its way to re-encounter Venus in 2015. However, due to a malfunction in the propulsion system, AKATSUKI can be only injected into the orbit much higher than that originally planned. It causes a couple of issues to be considered in its orbit design around Venus. One of which is long eclipse, which is the main topic of this paper. This paper introduces an orbit design strategy to avoid the long eclipse. An example of design result is shown as well. © 2013 2013 California Institute of Technology.

The trajectory design of the interplanetary mission "Demonstration and Experiment of Space Technology for INterplanetary voYage, DESTINY" is discussed. The trajectory optimization of low-thrust spacecraft pose a difficult design challenge. Moreover, DESTINY mission requires many constraints for the orbital design. In advance of optimizing the whole transfer mission, we investigated the basic theory to design which can take into account such constraints.

DESTINY which stands for "Demonstration and Experiment of Space Technology for INterplanetary voYage" is a mission candidate for the third mission of ISAS small science satellite series. DESTINY is launched by an Epsilon launch vehicle, JAXA's next-generation solid fuel rocket, and is firstly placed into a low elliptical orbit. It raises its altitude by the use of ion engine and reaches the Moon. Then, it is injected into transfer orbit for L 2 Halo orbit of the Sun-Earth system by using lunar gravity assist. On the way to L2 Halo orbit, DESTINY conducts demonstration and experiment on key advanced technology for future deep space missions. This paper presents the overview of DESTINY mission as well as its importance and significance in Japanese space science program.

The trajectory design of the interplanetary mission "Demonstration and Experiment of Space Technology for INterplanetary voYage, DESTINY" is discussed. The trajectory optimization of low-thrust spacecraft pose a difficult design challenge. Moreover, DESTINY mission requires many constraints for the orbital design. In advance of optimizing the whole transfer mission, we investigated the basic theory to design which can take into account such constraints.

The Japanese Venus explorer "Akatsuki (PLANET-C)", which now rotates about the Sun, will approach to Venus again in 2015. For the Venus orbit reinsertion, several trajectory strategies were devised. In this paper, we introduce the difficulties we faced in redesigning the trajectory of Akatsuki after the failure of the first Venus Orbit Insertion (VOI) in 2010 and report some newly devised trajectory control strategies including Gravity Brake Method, which will make the most of the solar perturbations to conduct the Venus orbit insertion for the second time.

DESTINY which stands for "Demonstration and Experiment of Space Technology for INterplanetary voYage" is a mission candidate for the third mission of ISAS small science satellite series. DESTINY is launched by an Epsilon launch vehicle, JAXA's next-generation solid fuel rocket, and is firstly placed into a low elliptical orbit. It raises its altitude by the use of ion engine and reaches the Moon. Then, it is injected into transfer orbit for L-2 Halo orbit of the Sun-Earth system by using lunar gravity assist. On the way to L-2 Halo orbit, DESTINY conducts demonstration and experiment on key advanced technology for future deep space missions. This paper presents the overview of DESTINY mission as well as its importance and significance in Japanese space science program.

In this study, we show the multiple revolution orbits design in Elliptic Restricted 3-Body Problem (ER3BP) of Sun, Earth and a particle 3-body system. For the deep-space observation, the location around L2 point is suitable because of the wide field of view to the outer space. In near future, some missions are planned to be putted into the Sun-Earth L2 point orbits. Depending on the mission requirements, the periodic orbits may not be the necessary condition. Therefore, the multiple revolution orbit closed in configuration space under arbitrary conditions are considered. Additionally, a preliminary calculation of the orbit maintenance is provided.

This work will present a Multi-Objective approach to the design of the initial, Low-Thrust orbit raising phase for JAXA's proposed technology demonstrator mission DESTINY. The proposed approach includes a simplified model for Low Thrust, many-revolution transfers, based on an analytical orbital averaging technique, and a simplified control parameterisation. Eclipses and J(2) perturbation are also accounted for. This is combined with a stochastic optimisation algorithm to solve optimisation problems in which conflicting performance figures of DESTINY's trajectory design are concurrently optimised. It will be shown that the proposed approach provides for a good preliminary investigation of the launch window and helps identifying critical issues to be addressed in future design phases.

AKATSUKI, the Japanese Venus explorer, once failed to inject itself into an orbit around Venus in 2010. AKATSUKI is now on its way to re-encounter Venus in 2015. However, due to a malfunction in the propulsion system, AKATSUKI can be only injected into the orbit much higher than that originally planned. It causes a couple of issues to be considered in its orbit design around Venus. One of which is long eclipse, which is the main topic of this paper. This paper introduces an orbit design strategy to avoid the long eclipse. An example of design result is shown as well.

The Venus explorer, Akatsuki, was launch on 20 May 2010. After 200-day journey through the interplanetary transfer orbit, it reached the Venus at the altitude of 550 km on 7 Dec 2010. However, it experienced a trouble of the explorer's propulsion system and was not able to be the Venus orbiter. It now orbits the Sun with the period of 203 days. In this paper, we discuss the trajectory design strategies for Akatsuki mission by introducing the constraints which come from the observation orbit and the spacecraft system. The details of planning and the results of orbital maneuvers are also shown in this paper.

A 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 solar polar region from the orbit largely inclined with the ecliptic plane. In order to achieve this severe mission target, possible trajectory sequences are investigated considering the application of various trajectory manipulation techniques. Three major trajectory options are listed up to achieve the mission objective. The first option, "SEP option", is characterized by the usage of solar electric propulsion (SEP) combined with the earth gravity assists. The other options are characterized by the usage of planetary gravity assists, and SEP is not used in these options. They are named "Jupiter option" and "Venus option" respectively. The comparison among the trajectory options is discussed in the paper, not only from the aspect of the orbital mechanics, but also from the aspects of the spacecraft design and operation.

Japanese Venus Climate Orbiter/AKATSUKI was proposed in 2001 with strong support by international Venus science community and approved as an ISAS (Institute of Space and Astronautical Science) mission soon after the proposal. The mission life we expected was more than two Earth years in Venus orbit. AKATSUKI was successfully launched at 06:58:22JST on May 21, 2010, by H-IIA F17. After the separation from H-IIA, the telemetry from AKATSUKI was normally detected by DSN Goldstone station (10:00JST) and the solar cell paddles' expansion was confirmed. The malfunction happened on the propulsion system during the Venus orbit insertion (VOI) on Dec 7, 2010. We failed to make the spacecraft become a Venus orbiter, and the spacecraft entered an orbit around the Sun with a period of 203 days. Most of the fuel still had remained, but the orbital maneuvering engine was found to be broken. We decided to use only the reaction control system (RCS) for orbital maneuver and three minor maneuvers in Nov 2011 were successfully done so that AKATSUKI will meet Venus in 2015. We are considering several scenarios for VOI using only RCS. Copyright © (2012) by the International Astronautical Federation.

The Space Infrared Telescope for Cosmology and Astrophysics (SPICA) is a 3.2m cooled (below 6K) telescope mission which covers mid- and far-IR waveband with unprecedented sensitivity. An overview of recent design updates of the Scientific Instrument Assembly (SIA), composed of the telescope assembly and the instrument optical bench equipped with Focal Plane Instruments (FPIs) are presented. The FPI international science and engineering review is on-going to determine the FPI suite onboard SPICA: at present the mandatory instruments and functions to perform the unique science objectives of the SPICA mission are now consolidated. The final decision on the composition of the FPI suite is expected in early 2013. Through the activities in the current preproject phase, several key technical issues which impact directly on the instruments' performances and the science requirements and the observing efficiency have been identified, and extensive works are underway both at instrument and spacecraft level to resolve these issues and to enable the confirmation of the SPICA FPI suite. © 2012 SPIE.

This study investigates the effective method to correct the transfer trajectory into the Halo-orbit by using the stable manifolds. The Halo-orbits around collinear Lagrange points are recently in the spotlight because of its periodicity, large field of view to the deep-space, and stable thermal environment. Japan Aerospace Exploration Agency (JAXA) is currently planning the first Japanese astronomical mission putted into the Halo-orbit named Space Infrared telescope for Cosmology and Astrophysics (SPICA). SPICA is designed to utilize the stable manifolds to transfer into the Halo-orbit. The stable manifolds, which are constructed by dynamical system theory (DST), are the strong feature to insert the spacecraft naturally into Halo-orbit. In this paper, we want to show the stable manifolds prediction using scaling method at first and efficient trajectory correction method by means of predicted stable manifolds in the Circular Restricted Three-Body Problem (CR3BP).

On December 7, 2010, AKATSUKI, the Japanese Venus explorer reached its destination and tried to inject into a closed orbit around Venus. However, due to a malfunction of the propulsion system, the maneuver was interrupted and AKATSUKI escaped the Venus into an interplanetary orbit. Telemetry data from AKATSUKI suggests the possibility to perform orbit maneuvers to reencounter Venus and retry Venus orbit insertion. Reported in this paper is an orbit plan investigated under this situation. The latest results reflecting the maneuvers conducted in the autumn 2011 is introduced as well.

Multi-objective design exploration (MODE) is applied to the trajectory design optimization problem for DESTINY mission, whose spacecraft will go to the moon by the ion engine from the large ellipse orbit. This approach revealed some important design knowledge such as tradeoff information among minimization of the operation time of ion engine, maximization of the final orbit energy, and minimization of the transit time under 20000 km altitude. This study also shows possibility of simultaneous design of trajectory and spacecraft. Copyright© (2012) by the International Astronautical Federation.

Multiobjective design exploration (MODE) framework is applied to the trajectory design optimization problem for SOLAR-C Plan A solar electric propulsion option. Present approach presents many designs that are better than the nominal design in all design objectives. Present approach also revealed some important design knowledge such as tradeoff information among maximization of final mass, maximization of final velocity, and maximization of minimum distance from the Sun. Present study also shows possibility of simultaneous design of trajectory and spacecraft.

This study investigates the trajectory design of the small scientific spacecraft, DESTINY (Demonstration and Space Technology for INterplanetary voYage), which aims to be launched by the third Japanese next-generation solid propellant rocket (Epsilon rocket) around 2017. In the DESTINY mission, the spacecraft will go to the moon by the ion engine from the large ellipse orbit. Afterward, by using the lunar swing-by, the spacecraft will put into the periodic orbit in the vicinity of the libration point (Halo orbit) of the Sun-Earth L2. This study focuses on the transfer trajectories from the moon to the Halo orbit.

The moon is attracting attention again as a target of space exploration. To focus on the lunar transfer, there is a sequence named "phasing orbit" which is different from the widely used direct transfer sequence. In this method, the spacecraft is not directly injected into the translunar orbit, but stays on a long elliptical phasing orbit for revolutions. A major merit of this method is that, by using the phasing orbit as a buffer, the translunar orbit can be fixed for acceptable width of launch window. Discussed in the paper is the design method of the phasing orbit using a chart developed by the author named "TM diagram.".

Multiobjective design exploration (MODE) framework is used to obtain useful knowledge from the Pareto-optimal solutions of a Solar observatory trajectory design problem. The trajectory design optimization problem is defined as a multiobjective design optimization problem and solved with a multiobjective evolutionary computation. Many designs that are better than the nominal design in all design objectives are found. The obtained Pareto-optimal solutions visualize tradeoff between the design objectives. Some data mining approaches are applied to the database of the obtained Pareto-optimal designs and show useful design information such as sensitivity of the design parameters to the design objectives. The present results shows MODE approach is useful for trajectory design optimization problems. © 2012 by Akira Oyama, Yasuhiro Kawakatsu, and Kazuko Hagiwara.

A study was conducted to investigate the three-dimensional (3-D)-resonant hopping strategy and present an automated trajectory design method for Japan Aerospace Exploration Agency's (JAXA) proposed Jupiter Magnetospheric Orbiter (JMO). The first part of the study showed that the Tisserand constant was the main problem parameter and it introduced the 3-D-Tisserand graph, which provided information about the problem. Some new analytical formulas were used to compute fixed-altitude gravity assists connecting resonant orbits. The second part of the study implemented the formulas in a branch-and-bound algorithm. The algorithm was applied to the JMO mission design, providing around 10,000 solutions in a few minutes. The algorithm and the solution space were the second main results of this study.

This paper presents the trajectory design for the Jupiter Magnetospheric Orbiter (JMO) in the Jovian system. The JMO is the Japanese contribution to the Europa Jupiter System Mission, and its science objectives include in-situ exploration of different regions of the magnetosphere and the remote sensing of the plasma torus from high latitudes. These science objectives require placing the spacecraft into a low-inclination, high-Apojove orbit, with subsequent gravity assists of the Galilean satellites to raise inclination and lower apojove. The trajectory design is the result of a complex multi-objective optimization problem: to minimize the mission cost and complexity, the spacecraft avoids high-radiation regions; at the same time, the trajectory minimizes the transfer time and the propellant mass and maximizes the final inclination to Jupiter's equator. The optimal strategy for the JMO consists of one Ganymede gravity assist followed by a sequence of gravity assists at Callisto. This design is also robust to changes in the trajectory requirements, which is an important feature in any preliminary design. The main result of this work is the analysis of the multidimensional solution space and the discussion of the trade-offs between solutions in the Pareto front. Copyright © 2011 by Serhat Hosder.

On December 7, 2010, AKATSUKI, the Japanese Venus explorer reached its destination and tried to inject into a closed orbit around Venus. However, due to a malfunction of the propulsion system, the maneuver was interrupted and AKATSUKI escaped the Venus into an interplanetary orbit. Telemetry data from AKATSUKI suggests the possibility to perform orbit maneuvers to reencounter Venus and retry Venus orbit insertion. Reported in this paper is an orbit plan investigated under this situation. The latest results reflecting the maneuvers conducted in the autumn 2011 is introduced as well.

The moon is attracting attention again as a target of space exploration. To focus on the lunar transfer, there is a sequence named "phasing orbit" which is different from the widely used direct transfer sequence. In this method, the spacecraft is not directly injected into the translunar orbit, but stays on a long elliptical phasing orbit for revolutions. A major merit of this method is that, by using the phasing orbit as a buffer, the translunar orbit can be fixed for acceptable width of launch window. Discussed in the paper is the design method of the phasing orbit using a chart developed by the author named "TM diagram."

We present the overview and the current status of SPICA (Space Infrared Telescope for Cosmology and Astrophysics), which is a mission optimized for mid- and far-infrared astronomy with a cryogenically cooled 3.2 m telescope. SPICA has high spatial resolution and unprecedented sensitivity in the mid- and far-infrared, which will enable us to address a number of key problems in present-day astronomy, ranging from the star-formation history of the universe to the formation of planets. To reduce the mass of the whole mission, SPICA will be launched at ambient temperature and cooled down on orbit by mechanical coolers on board with an efficient radiative cooling system, a combination of which allows us to have a 3-m class cooled (6 K) telescope in space with moderate total weight (3.7t). SPICA is proposed as a Japanese-led mission together with extensive international collaboration. ESA's contribution to SPICA has been studied under the framework of the ESA Cosmic Vision. The consortium led by SRON is in charge of a key focal plane instrument SAFARI (SPICA Far-Infrared Instrument). Korea and Taiwan are also important partners for SPICA. US participation to SPICA is under discussion. The SPICA project is now in the "risk mitigation phase". The target launch year of SPICA is 2022. © 2012 SPIE.

This paper investigates the preliminary orbit maintenance for the next-generation infrared astronomical mission, Space Infrared Telescope for Cosmology and Astrophysics, which is to be launched into the sun-Earth L2 halo orbit. Particularly, the impact of the reaction wheel unloading ΔV on the spacecraft trajectory from the viewpoint of orbit maintenance is analyzed. In addition, the nondivergent and suppressive methods using the characteristics of the dynamical theory for unloading ΔV are proposed and validated. Copyright © 2012 by the American Institute of Aeronautics and Astronautics, Inc.

This study investigates the trajectory design of the small scientific spacecraft, DESTINY (Demonstration and Space Technology for INterplanetary voYage), which aims to be launched by the third Japanese next-generation solid propellant rocket (Epsilon rocket) around 2017. In the DESTINY mission, the spacecraft will go to the moon by the ion engine from the large ellipse orbit. Afterward, by using the lunar swing-by, the spacecraft will put into the periodic orbit in the vicinity of the libration point (Halo orbit) of the Sun-Earth L2. This study focuses on the transfer trajectories from the moon to the Halo orbit.

Multiobjective design exploration (MODE) framework is applied to the trajectory design optimization problem for SOLAR-C Plan A solar electric propulsion option. Present approach presents many designs that are better than the nominal design in all design objectives. Present approach also revealed some important design knowledge such as tradeoff information among maximization of final mass, maximization of final velocity, and maximization of minimum distance from the Sun. Present study also shows possibility of simultaneous design of trajectory and spacecraft.

Two spacecraft or more are assumed to be in a state of loose formation flying around a collinear Lagrangian point in the Sun-Earth circular restricted three-body problem (CR3BP) system. The orbit reference of choice for the leader is a halo orbit, and the followers are assumed to follow nearby and be constrained either geometrically or in size. This type of formation could be useful in the future for constructing space ports, space telescopes, astronomical spacecraft requiring sun shields and, with greater numbers, spacecraft swarm missions. The formation design method is constructed by firstly seeking the local coordinate system from the monodromy matrix through extraction of the independent bases that span the space of the halo orbit. To nullify diverging and converging motion, we confine the relative motion to within the periodic subspaces. We observe two modes of relative motion within these subspaces, long-term and short-term motion. In this study, we approximate the long-term motion by deriving a discrete formulation of independent directions based on the eigenvectors of the monodromy matrix, while for the short-term motion we approximate the fundamental set solutions using Fourier series and additional linear functions. Since the size of the formation discussed is significantly smaller than that of the halo orbit, the formation design method can fundamentally be stated as a process of linearly combining these approximations to achieve the desired formation. Consequently, use of this approach transforms formation design from a differential equation problem into an algebraic one, and furthermore enables the long-term and short-term motion design problems to be handled either jointly or separately. A set of design examples is presented to demonstrate the validity of the design method. © 2012 The Japan Society for Aeronautical and Space Sciences.

This study investigates the effective method to correct the transfer trajectory into the Halo-orbit by using the stable manifolds. The Halo-orbits around collinear Lagrange points are recently in the spotlight because of its periodicity, large field of view to the deep-space, and stable thermal environment. Japan Aerospace Exploration Agency (JAXA) is currently planning the first Japanese astronomical mission putted into the Halo-orbit named Space Infrared telescope for Cosmology and Astrophysics (SPICA). SPICA is designed to utilize the stable manifolds to transfer into the Halo-orbit. The stable manifolds, which are constructed by dynamical system theory (DST), are the strong feature to insert the spacecraft naturally into Halo-orbit. In this paper, we want to show the stable manifolds prediction using scaling method at first and efficient trajectory correction method by means of predicted stable manifolds in the Circular Restricted Three-Body Problem (CR3BP).

Two spacecrafts are assumed to fly in a loose formation flying around a collinear lagrangian point of the Sun-Earth Circular Restricted Three-Body Problem (CRTBP) system. Orbit reference of choice for the leader is a halo orbit and the follower is conceived to follow nearby and constrained geometrically or in size. This type of formation could be useful in future for constructing space based ports, space telescopes, astronomical spacecrafts requiring sun shield, and with more constituents, spacecraft swarm missions. The formation design method is constructed by firstly sought the local coordinate system from the Monodromy matrix by extracting the independent directions which spanned the periodic subspace of the Halo orbit. We then approximate the relative motion of each direction by using Fourier series and linear function. Since the size of formation discussed here is significantly small compared with the size of the Halo orbit, we can devise formation design method simply as to how linearly combine the approximations. This consequently means that we transform the formation design problems into algebraic problems. Suppressing oscillation in the Sun-Earth direction is shown here among other things as an example to demonstrate the formation design method.

This study analyzes the accessibility to asteroids for the Small Solar Power Sail Demonstrator, IKAROS mission. IKAROS was launched in May 21, 2010 with the Venus orbiter, Akatsuki, and IKAROS flied by the Venus in December, 2010. This study focuses on searching the close approach to asteroids after Venus flyby.

This study investigates the availability of the loose station keeping using the dynamical system theory considering the SPICA shading constraint. As a result, the amount of the correction maneuver is approximately 0.1 m/s/y, which achieve the mission requirement enough.

Two spacecrafts are assumed to fly in a loose formation flying around a collinear lagrangian point of the Sun-Earth Circular Restricted Three-Body Problem (CRTBP) system. Orbit reference of choice for the leader is a halo orbit and the follower is conceived to follow nearby and constrained geometrically or in size. This type of formation could be useful in future for constructing space based ports, space telescopes, astronomical spacecrafts requiring sun shield, and with more constituents, spacecraft swarm missions. The formation design method is constructed by firstly sought the local coordinate system from the Monodromy matrix by extracting the independent directions which spanned the periodic subspace of the Halo orbit. We then approximate the relative motion of each direction by using Fourier series and linear function. Since the size of formation discussed here is significantly small compared with the size of the Halo orbit, we can devise formation design method simply as to how linearly combine the approximations. This consequently means that we transform the formation design problems into algebraic problems. Suppressing oscillation in the Sun-Earth direction is shown here among other things as an example to demonstrate the formation design method.

This paper presents, the operations scenario, specifically the initial phase operation, of the Attitude and Orbit Control System (AOCS) of PLANET-C: Venus Climate Orbiter Mission "AKATSUKI". AKATSUKI is the first exploration program from Japan to Earth's inferior planetary neighbor, Venus, and will be launched by an H-A vehicle on May 18th 2010 (Japan Standard Time: JST) from Kagoshima Space Center. AKATSUKI aims at understanding the atmospheric characteristics of Venus. To accomplish this mission, AKATSUKI has five camera instruments onboard to image ultraviolet and thermal infrared wavelengths, to detect lightning with a high-speed imager, and to observe the vertical structure of the atmosphere using radio science techniques. To achieve these objectives, AKATSUKI has four operation phases; such as launch and initial check, cruise, Venus orbit injection (VOI) and nominal Venus observation. The VOI phase is a particularly critical phase to enter equatorial elongated Venus orbit. To achieve this trajectory, AKATSUKI requires an AOCS with the tolerance for the stringent high temperature environment produced in the inferior planet. © 2010 SICE.

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 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.

Trojan asteroids exist in nearby of triangular Lagrangian points of the Sun-Jupiter system and are believed to contain primitive information on the early formation of our Solar System. Additionally their origin is also unclear. Considering these significances, we report in this paper mission analysis on the Trojans rendezvous mission. In the mission design two major steps were conducted, target selection and sequence improvements. In the target selection we explored three types of trajectories, direct transfer and the use of gravity assist of Mars and Jupiter and we finally introduced ΔVEGA and low-thrust propulsion utilization to the detailed design to refine the selected sequences.

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.

Transfers from the Earth to the Halo orbits by exploiting the stable manifolds, which are the characteristic of the dynamical system in the three-body problem are investigated. First, the conditions for the transfer to the Halo orbits such as the altitude of perigee and the orbital inclination are shown. Next, the launch window for the lift-off to the Halo orbit using stable manifold at the Japanese launch site is discussed.

Preliminary mission design for the next-generation infrared astronomical mission, SPICA, which is to be launched into the Sun-Earth L2 Halo orbit, was studied. Our specific focus is on investigating the impact of translational force (unloading ΔV), caused by the thruster unloading for accumulated angular momentum on the spacecraft reaction wheels, with respect to the orbit maintenance. In addition, suppression methods for orbital disturbances caused by the unloading ΔV are proposed and validated.

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.

Preliminary mission design for the next-generation infrared astronomical mission, SPICA, which is to be launched into the Sun-Earth L2 Halo orbit, was studied. Our specific focus is on investigating the impact of translational force (unloading Delta V), caused by the thruster unloading for accumulated angular momentum on the spacecraft reaction wheels, with respect to the orbit maintenance. In addition, suppression methods for orbital disturbances caused by the unloading Delta V are proposed and validated.

Trojan asteroids exist in nearby of triangular Lagrangian points of the Sun-Jupiter system and are believed to contain primitive information on the early formation of our Solar System. Additionally their origin is also unclear. Considering these significances, we report in this paper mission analysis on the Trojans rendezvous mission. In the mission design two major steps were conducted, target selection and sequence improvements. In the target selection we explored three types of trajectories, direct transfer and the use of gravity assist of Mars and Jupiter and we finally introduced Delta VEGA and low-thrust propulsion utilization to the detailed design to refine the selected sequences.