川勝 康弘

In the modern space development, small-scale deep space mission should be realized to promote frequent and challenging deep space mission. Therefore, the efficient and quick design method to construct Earth escape trajectory with high flexibility in the boundary condition such as escape velocity, direction and timing is strongly demanded. In this paper, the families of Moon-to-Moon transfers with sequential lunar swing-by on a hyperbolic orbit are computed and stored in a database. These families are useful to enhance the Earth escape energy and to change escape direction which could lead a spacecraft to further destinations.

Japan has been preparing for a Phobos sample return mission. In this mission, it is worth taking the opportunity to also observe Deimos. This paper discusses trajectory options in the case of utilizing a flyby as a Deimos-observation method. There are two problems when designing mission orbits that meet requirements. As a result of mission analysis about various options, it was revealed which mission can be conducted and how much the probe satisfies the requirements. Finally, the effective trajectory to the case that the probe has rendezvous with a satellite of Mars, and flyby with the other satellite is shown.

© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All Rights Reserved. This study is devoted to explore space trajectory for DESTINY (Demonstration and Experiment of Space Technology for INterplanetary voYage), which was proposed to ISAS (Institute of Space and Astronautical Science) Epsilon-class small program in 2013 based on the \Space Science & Exploration Roadmap” which is proposed by ISAS and later approved by the government committee of space policy. In the DESTINY mission, spacecraft is first injected into a low elliptical orbit by the Epsilon rocket, and it raises the altitude to reach the Moon using an ion engine system (IES). After that it is injected into a transfer orbit of L2 Halo orbit of the Sun-Earth system through gravity assist of the Moon. While the spacecraft revolves around the Earth for several hundred times, it increases its altitude little by little, and thus, launch time and the thrusting profile must be chosen properly. It is very important to note that there are many conicting requirements such as reduction of fuel consumption, total flight time, and the maximum eclipse time and so forth. To satisfy these requirements, many-objective evolutionary computation is applied to search for a better orbital design.

© Copyright 2017 by the International Astronautical Federation (IAF). All rights reserved. Martian Moons eXploration (MMX) is a mission under study in ISAS/JAXA to be launched in 2020s. This paper introduces the concept of MMX mission. "How was water delivered to rocky planets and enabled the habitability of the solar system?" This is the key question to which MMX is going to answer. Solar system formation theories suggest that rocky planets must have been born dry. Delivery of water, volatiles, organic compounds etc. from outside the snow line entitles the rocky planet region to be habitable. Small bodies as comets and asteroids play the role of delivery capsules. Then, dynamics of small bodies around the snow line in the early solar system is the issue that needs to be understood. Mars was at the gateway position to witness the process, which naturally leads us to explore two Martian moons, Phobos and Deimos, to answer to the key question. The goal of MMX is to reveal the origin of the Martian moons, and then to make a progress in our understanding of planetary system formation and of primordial material transport around the border between the inner- and the outer-part of the early solar system. On the origin of Martian moons, there are two leading hypotheses, "Captured primordial asteroid" and "Giant Impact". We decide to collect samples from a Martian moon to conclude this discussion, and on the conclusion, to investigate further to improve our understanding of material distributions and transports at the edge of the inner part of the early solar system as well as of planetary formations. Moreover, circum-Martian environment will be measured and Martian atmosphere will be observed to improve our views of evolutions of Martian moons as well as Mars surface environmental transition. In the conceptual design phase, the goals and objectives of the mission are defined, and the feasibility of the mission is evaluated. Fundamental engineering options are listed up, and trade-off studies are conducted to define baseline plan. Key technology issues are identified and their technology readiness is evaluated. The results will be shown in the paper.

© 2017 The Japan Society for Aeronautical and Space Sciences. This paper describes development strategies and on-orbit results of the attitude determination and control system (ADCS) for the world's first interplanetary micro-spacecraft, PROCYON, whose advanced mission objectives are optical navigation or an asteroid close flyby. Although earth-orbiting micro-satellites already have ADCSs for practical missions, these ADCSs cannot be used for interplanetary micro-spacecraft due to differences in the space environments of their orbits. To develop a new practical ADCS, four issues for practical interplanetary micro-spacecraft are discussed: initial Sun acquisition without magnetic components, angular momentum management using a new propulsion system, the robustness realized using a fault detection, isolation, and recovery (FDIR) system, and precise attitude control. These issues have not been demonstrated on orbit by interplanetary micro-spacecraft. In order to overcome these issues, the authors developed a reliable and precise ADCS, a FDIR system without magnetic components, and ground-based evaluation systems. The four issues were evaluated before launch using the developed ground-based evaluation systems. Furthermore, they were successfully demonstrated on orbit. The architectures and simulation and on-orbit results for the developed attitude control system are proposed in this paper.

© 2016 The Society of Instrument and Control Engineers - SICE. This paper presents the planning, flight results and lessons learned of flyby guidance experiments of interplanetary micro-spacecraft PROCYON. PROCYON is the world's first interplanetary micro-spacecraft and was launched on 3rd December, 2014. Orbital control of interplanetary micro-spacecraft is challenging because of severe restriction and lower reliability on spacecraft system. For guidance strategy of PROCYON, we have introduced an innovative guidance strategy by two-stage stochastic programming for thrust-direction-constrained problem. Although the flight experiment has many difficulties especially on navigation, the flight result shows that we successfully demonstrate that PROCYON has been guided to the target point with objective guidance accuracy, which is within 100[km] on B-plane at 3,000,000[km] distance from the Earth. These results contributes the future flyby navigation and guidance for interplanetary micro-spacecraft, which has severe constraints and lower reliability on spacecraft system.

© 2016 IEEE. PROCYON is a first full-scale, 50-kg-class probe featuring most of the key technologies for deep-space exploration. It was developed by the University of Tokyo and ISAS/JAXA and launched with Hayabusa 2 on 3 Dec 2014. PROCYON has a newly developed X-band telecommunication system fully compatible with the frequency range, up- and down-link turn-around ratio, modulation scheme, and DDOR tones following CCSDS-recommended standards, and it can establish X-band coherent two-way communication and ranging links with deep-space stations as larger deep-space probes have done. The total mass of the onboard telecommunication system is 7.3 kg excluding its RF coaxial harness, and total power consumption during two-way communication, 15 W of RF output power at SSPA, is 54.3 W. After launch, PROCYON's telecommunication system has been successfully working according to the system design. These achievements will provide core technologies for next-generation deep-space exploration by ultra-small probes.

<p>DESTINY: the Demonstration and Experiment of Space Technology for Interplanetary Voyage, which is a candidate mission of Epsilon Launch Vehicle, plans to execute scientific observations using instruments with the mass of up to about 10 kg on the transfer and Halo orbit of the sun to earth Lagrangian point L1/L2 or on the fly-by orbit of near earth objects (NEO). Potential scientific objects include in-situ observation and remote sensing from these space are solar system explorations, such as, the observations of plasma and energetic particles around the terrestrial magnetosphere, inter-planetary and inter-stellar dust, and NEO. It is also considered to be useful for the pilot observations for future infrared, gamma-ray, and cosmic-ray space astronomical telescope in the deep space. Applied missions of DESTINY will be able to go to deep space with higher mass of payloads. Using the Epsilon Launch Vehicle, it will convey instruments of up to 50 kg to the space between Venus and Mars. DESTINY launched by the improved launch vehicle with the power of M-V rocket will carry payloads of up to 200 kg into the orbit of Venus and Mars. In these phases, Explorations for Venus, Mars, and multiple NEO, and astronomical observations from the deep space observatory will be realized by low cost deep space missions.</p>

Copyright © 2016 by the authors. All rights reserved. The Exploration Mission-1 (EM1) is the first test flight of NASA's new Space Launch System. Scheduled for launch in 2018, EM1 will carry the Orion Multi-Purpose Crew Vehicle (MPCV) into a cislunar orbit, together with a secondary payload composed by 13 cubesat. Two of these cubesat are currently proposed by JAXA: EQUULEUS, a 6U Earth-Moon Lagrangian-Point orbiter (in collaboration with the University of Tokyo); and SLSLIM, a 6U Moon lander. This paper presents the mission analysis work for EQUULEUS, while a second paper presents the mission analysis work for SLSLIM. EQUULEUS mission objectives are demonstrating cubesat orbit control techniques within the Sun-Earth-Moon regions; understanding the Earth's radiation environment; characterizing the flux of impacting meteors at the far side of the Moon; and demonstrating future exploration scenarios with a deep-space port at the Lagrange points. Following MPCV disposal, EQUULEUS is separated by the upper stage towards a lunar flyby, which, if not corrected, would result in an Earth escape trajectory. For this reason, after one-day orbit determination a trajectory correction maneuver is performed by the onboard thrusters to pump up the flyby perilune and put the spacecraft into an Moon-return orbit. Exploiting Sun perturbations, multiple lunar flybys and small trajectory correction maneuvers, EQUULEUS will be finally placed into a libration orbit around the Earth-Moon L2 point. We present the trajectory design process and a few sample trajectories, with the current baseline and the launch window analysis. Several astrodynamics techniques are described, including the search for Lunar-return orbits in the Earth-Sun Circular Restricted Three-Body Problem (first introduced by Lantoine in [1], and further developed by Garcia [2] for EQUULEUS and other applications); and the design of Libration orbits and low-energy transfers in real ephemeris.

After the successful launch on the world first spacecraft, Sputnik 1 by the former Soviet Union in 1957, 58 years has passed. In 1960, Pioneer 5 of the United States escaped the Earth's gravity at the first time, and since then many interplanetary explorers had set to sail interplanetary. However, even in the present day, interplanetary voyages are not still easy. First, interplanetary missions require large amounts of delta-V, and second, the opportunity to get to the destination opens only every synodic period with the destination celestial body. For example, the synodic period with Mars is about 2 years, which means the opportunity to get to Mars opens every 2 years. For such circumstances, this paper proposes a new type of low-thrust orbit design method, "Interplanetary Parking Method" that realizes "anytime" launch of deep-space explorers. The proposed interplanetary parking method enables to make an Earth return orbit with an arbitrary time-of-flight connecting to the minimum energy transfer orbit to a destination. While the time-of-flight of the transfer orbit is fixed, the Earth return orbit with the arbitrary time-of-flight virtually eliminates the severe launch window constraint in interplanetary missions. As application of the proposed method, the paper demonstrates dual launch trajectory design of explorers to different destinations i.e., Mars and Venus. The proposed method will widen the scope of opportunity for interplanetary missions.

DESTINY+ is a small-seized and high performance deep space vehicle proposed for public offering small-sized plan space science mission of ISAS/JAXA. DESTINY+ is injected into an extended elliptical orbit launched by Epsilon rocket. The orbit is spiraled upward by the low-thrust of IES. And the swing-by is designed to give DESTINY momentum to Asteroid Phaethon flyby. After Phaethon flyby, DESTINY+ plan to go back to Earth for gravity assist and go to another asteroid. DESTINY+ has several mission objectives, including: demonstration and experiment of space technology of interplanetary voyage; Phaethon flyby with reusable probe; compact avionics as for Engineering mission, and the investigation of the process to the end of evolution of primitive body; the limitation of initial state and the evolution process of the meteor shower dust as for Science mission. This paper discusses DESTINY+'s low-thrust trajectory design and the related system analysis. As for the spiral upward trajectory phase, the low-thrust trajectory is optimized by the multi-objective optimization using genetic algorithm. In this phase, we minimize the time of flight, the passage time of radiation belt, the work time of IES and the shadow time. After the spacecraft reaches to the moon's path, it utilizes the moon swing-by several times to connect to the transfer trajectory for Asteroid Phaethon. Parallel to the trajectory design, the radiation effect analysis, thermal environmental analysis, attitude analysis and ground station visibility analysis for operation are achieved. From these study, we can show the feasibility of the mission design of DESTINY+.

Copyright © 2016 by the authors. All rights reserved. The Mars system has been the target of many space exploration missions, yet the origin of the moons Phobos and Deimos is still not understood. The three prevailing hypothesis are: 1) the moons are type C or D asteroids, 2) the moons are remnants of Mars formation, or 3) the moons were formed by accretion of ejecta from a planetesimal impact on Mars. Unraveling the mystery of the origins of Deimos and Phobos would impact our understanding of the origin and formation of the solar system, and of the nature of the matter that was incorporated. For this reason, in the last decade several space agencies have studied mission options dedicated to the exploration of the moons of Mars. In 2011 the Russian Space Agency launched the sample return Phobos-Grun], but the spacecraft has failed to escape the Earth orbit. More recently, several mission scenarios were proposed by ESA and NASA (Phootprint, PANDORA, PADME, and MERLIN), including flyby missions, orbiters, landers, and sample return missions. Because of the scientific importance of such a mission, and building on the experience of asteroid sample return missions Hayabusa and Hayabusa2, the Japanese space agency is now planning a new sample return mission to be launch in the early 2020s. The mission is currently the main candidate for the next Japanese large-class spacecraft, following Hayabusa 2 (currently flying) and the space observatory ASTRO-H (launch in 2016). This paper presents the mission analysis work carried for the mission, with focus on the transfers from the launch to the moons, and from the moons back to the Earth. Several architectures are considered, depending on the propulsion system used and on the moon or moons chosen for the landing. An analytical formulation is also presented for quick estimate of the three-maneuver orbit insertion strategy.

This work explores the target selection and trajectory design of the mission candidate for ISAS/JAXA's small science satellite series, DESTINY. This mission combines unique aspects of the latest satellite technology and exploration of transition bodies to fill a technical and scientific gap in the Japanese space science program. The spacecraft is targeted to study the comet-asteroid transition body (3200) Phaethon through a combination of low-thrust propulsion and Earth Gravity Assist. The trajectory design concept is presented in details together with the launch window and flyby date analysis. Alternative targets for a possible the mission extension scenario are also explored.

Two types of boundary values problems of low-thrust trajectories are considered: the position matching problem and the reachability problem. We perform experiments on three approaches to solve and to analyze such boundary value problems, particularly the feasible regions of the solutions. A linear approximation method is applied which can compute the attainable sets of solutions efficiently and accurately for short transfer arcs as compared with the nonlinear constraint satisfaction method. An optimization approach that can map out the feasible set of solutions for long transfer arcs is also examined. Our method can provide an initial estimate for broad searches of multi-leg low-thrust trajectories.

© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. This work discusses on designing fast and efficient Mars escape trajectory for Martian Moons eXplorer in the three-body system using chemical and electric propulsion. Chemical propulsion is used for fast low-energy escape from Mars and electric propulsion is used to increase v-infinity and to re-encounter with Mars for a gravity assist. We propose a method called “all-three-body method" and we compare the new method with the patched three-and-two-body method, and a parametric study is carried out. Using electric propulsion soon after Mars escape injection done by chemical propulsion, the all-three-body method would consume less fuel than patched three-and-two-body method. Limiting the use of chemical propulsion could also increase the final spacecraft mass when electric propulsion is used within 10 days after impulsive maneuver.

In this paper, we investigate the low-energy escape trajectory design for a mission called Martian Moons eXplorer to achieve the world's first sample return from Martian moon. The hybrid usage of chemical and electric propulsion with combination of the three-body and two-body problems has come into consideration in order to seek a fast low-energy escape from Mars. We first study the needs of pre-departure sequence. Then, we determine the transition point from a low-energy three-body phase to a low-thrust two-body phase, in which the tube dynamics is employed for the low-energy three-body phase. We finally develop charts to reveal the relation between the velocity in Mars Escape Injection maneuver and the required time of flight.

On-board Orbit Determination (OD) using the sun sensor and the Optical Navigation Camera (ONC) for the Autonomous Navigation (AutoNav) is focused on in this paper. In deep space missions, the OD has been performed by Range and Range-Rate (RARR), which is the traditional ground tracking approach by radio wave. The RARR enables the accuracy of the OD to be higher than the other methods. However, such the radio navigation has the inevitable problems, e.g. the delay of radio wave, the reduction of radio wave strength and the transmitter limitation. The influence of these problems becomes significant especially for deep space missions. Furthermore, people must stay and operate the spacecraft on the ground station, which makes the operating cost considerable. Therefore, there has been a growing interest in the AutoNav of the spacecraft in recent years because the AutoNav can solve the above-mentioned problems. The realization of the AutoNav in deep space can eliminate the complexity of operation on the ground station, and especially it has the significant impact on the operation cost reduction. This paper focuses on the case of the Earth resonant trajectory as an actual mission. The usefulness of the sun observation by sun sensors is discussed for the proposed method. Then, the observation of asteroids is also argued.

© 2016 IAA Inspired by successful extended missions such as the ISEE-3, an investigation for the extended mission that involves a lunar encounter following a Sun-Earth halo orbit mission is considered valuable. Most previous studies present the orbit-to-orbit transfers where the lunar phase is not considered. Intended for extended missions, the present work aims to solve for the minimum phasing ∆V for various initial lunar phases. Due to the solution multiplicity of the two-point boundary value problem, the general constrained optimization algorithm that does not identify multiple feasible solutions is shown to miss minima. A two-step differential corrector with a two-body Lambert solver is developed for identifying multiple solutions. The minimum ∆V associated with the short-way and long-way approaches can be recovered. It is acquired that the required ∆V to cover all initial lunar phases is around 45 m/s for the halo orbit with out-of-plane amplitude Az greater than 3.5×105 km, and 14 m/s for a small halo orbit with Az=1×105 km. In addition, the paper discusses the phasing planning based on the ∆V result and the shift of lunar phase with halo orbit revolution.

© 2016 The Japan Society for Aeronautical and Space Sciences. This paper investigates the Earth escape for spacecraft in a Sun-Earth halo orbit. The escape trajectory consists of first ejecting to the unstable manifold associated with the halo orbit, then coasting along the manifold until encountering the Moon, and finally performing lunar-gravity-assisted escape. The first intersection of the manifold tube and Moon's orbit results in four intersection points. These four manifold-guided encounters have different relative velocities (v∞) to the Moon; therefore, the corresponding lunar swingbys can result in different levels of characteristic energy (C3) with respect to the Earth. To further exploit these manifold-guided lunar encounters, subsequent swingbys utilizing solar perturbation are considered. A graphical method is introduced to reveal the theoretical upper limits of the C3 achieved by double and multiple swingbys. The numerically solved Sun-perturbed Moon-to-Moon transfers indicate that a second lunar swingby can efficiently increase C3. Compared to the direct low-energy escape along the manifold, applying a portion of the lunar swingbys before escape is shown to be more advantageous for deep-space mission design.

After the successful launch on the world first spacecraft, Sputnik 1 by the former Soviet Union in 1957, 58 years has passed. In 1960, Pioneer 5 of the United States escaped the Earth's gravity at the first time, and since then many interplanetary explorers had set to sail interplanetary. However, even in the present day, interplanetary voyages are not still easy. First, interplanetary missions require large amounts of delta-V, and second, the opportunity to get to the destination opens only every synodic period with the destination celestial body. For example, the synodic period with Mars is about 2 years, which means the opportunity to get to Mars opens every 2 years. For such circumstances, this paper proposes a new type of low-thrust orbit design method, "Interplanetary Parking Method" that realizes "anytime" launch of deep-space explorers. The proposed interplanetary parking method enables to make an Earth return orbit with an arbitrary time-of-flight connecting to the minimum energy transfer orbit to a destination. While the time-of-flight of the transfer orbit is fixed, the Earth return orbit with the arbitrary time-of-flight virtually eliminates the severe launch window constraint in interplanetary missions. As application of the proposed method, the paper demonstrates dual launch trajectory design of explorers to different destinations i.e., Mars and Venus. The proposed method will widen the scope of opportunity for interplanetary missions.

On-board Orbit Determination (OD) using the sun sensor and the Optical Navigation Camera (ONC) for the Autonomous Navigation (AutoNav) is focused on in this paper. In deep space missions, the OD has been performed by Range and Range-Rate (RARR), which is the traditional ground tracking approach by radio wave. The RARR enables the accuracy of the OD to be higher than the other methods. However, such the radio navigation has the inevitable problems, e.g. the delay of radio wave, the reduction of radio wave strength and the transmitter limitation. The influence of these problems becomes significant especially for deep space missions. Furthermore, people must stay and operate the spacecraft on the ground station, which makes the operating cost considerable. Therefore, there has been a growing interest in the AutoNav of the spacecraft in recent years because the AutoNav can solve the above-mentioned problems. The realization of the AutoNav in deep space can eliminate the complexity. of operation on the ground station, and especially it has the significant impact on the operation cost reduction. This paper focuses on the case of the Earth resonant trajectory as an actual mission. The usefulness of the sun observation by sun sensors is discussed for the proposed method. Then, the observation of asteroids is also argued.

This work explores the target selection and trajectory design of the mission candidate for ISAS/JAXA's small science satellite series, DESTINY. This mission combines unique aspects of the latest satellite technology and exploration of transition bodies to fill a technical and scientific gap in the Japanese space science program. The spacecraft is targeted to study the comet-asteroid transition body (3200) Phaethon through a combination of low-thrust propulsion and Earth Gravity Assist. The trajectory design concept is presented in details together with the launch window and flyby date analysis. Alternative targets for a possible the mission extension scenario are also explored.

Two types of boundary values problems of low-thrust trajectories are considered: the position matching problem and the reachability problem. We perform experiments on three approaches to solve and to analyze such boundary value problems, particularly the feasible regions of the solutions. A linear approximation method is applied which can compute the attainable sets of solutions efficiently and accurately for short transfer arcs as compared with the nonlinear constraint satisfaction method. An optimisation approach that can map out the feasible set of solutions for long transfer arcs is also examined. Our method can provide an initial estimate for broad searches of multi-leg low-thrust trajectories.

In this paper, we investigate the low-energy escape trajectory design for a mission called Martian Moons eXplorer to achieve the world's first sample return from Martian moon. The hybrid usage of chemical and electric propulsion with combination of the three-body and two-body problems has come into consideration in order to seek a fast low-energy escape from Mars. We first study the needs of pre-departure sequence. Then, we determine the transition point from a low-energy three-body phase to a low-thrust two-body phase, in which the tube dynamics is employed for the low-energy three-body phase. We filially develop charts to reveal the relation between the velocity in Mars Escape Injection maneuver and the required time of flight.

© 2016 Nakamura et al. AKATSUKI is the Japanese Venus Climate Orbiter that was designed to investigate the climate system of Venus. The orbiter was launched on May 21, 2010, and it reached Venus on December 7, 2010. Thrust was applied by the orbital maneuver engine in an attempt to put AKATSUKI into a westward equatorial orbit around Venus with a 30-h orbital period. However, this operation failed because of a malfunction in the propulsion system. After this failure, the spacecraft orbited the Sun for 5 years. On December 7, 2015, AKATSUKI once again approached Venus and the Venus orbit insertion was successful, whereby a westward equatorial orbit with apoapsis of ∼440,000 km and orbital period of 14 days was initiated. Now that AKATSUKI's long journey to Venus has ended, it will provide scientific data on the Venusian climate system for two or more years. For the purpose of both decreasing the apoapsis altitude and avoiding a long eclipse during the orbit, a trim maneuver was performed at the first periapsis. The apoapsis altitude is now ~360,000 km with a periapsis altitude of 1000-8000 km, and the period is 10 days and 12 h. In this paper, we describe the details of the Venus orbit insertion-revenge 1 (VOI-R1) and the new orbit, the expected scientific information to be obtained at this orbit, and the Venus images captured by the onboard 1-μm infrared camera, ultraviolet imager, and long-wave infrared camera 2 h after the successful initiation of the VOI-R1.

Copyright © 2015 by the American Institute Federation of Aeronautics and Astronautics. Inc. All rights reserved. PROCYON is the first deep-space micro-spacecraft; it was developed at low cost and short time (about one year) by the University of Tokyo and JAXA, and was launched on December 3rd, 2014 as a secondary payload of the H II A launch of Hayabusa2. The mission primary objective is the technology demonstration of a microspacecraft bus for deepspace exploration; the second objectives are several engineering and science experiments, including an asteroid flyby. This paper presents PROCYON high-fidelity, very-low-Thrust trajectory design and implementation, subject to mission and operation constraints. Contingency plans during the first months of operations are also discussed. All trajectories are optimized in high-fidelity model with jTOP, a mission design tool first presented in this paper. Following the ion engine failure of March 2015, it was found the nominal asteroid could not be targeted if the failure was not resolved by mid-April. A new approach to compute attainable sets for low-Thrust trajectories is also presented.

This study objectives are to devise orbit design methods and propose orbits fulfilling orbit constraints. The orbits were designed as follows: The orbiter's longitude of ascending node and argument of periapsis, in a Mars-Sun fixed coordinate system, are taken as design variables and the orbit constraint, proposed by a science group, is used as an evaluation function. A curved line that expresses a change history of orbiter's elements is draw in the plane. As a result, it is possible to find out rough initial values of longitude of ascending node and argument of periapsis suitable for the mission visually by moving the curved line, and also rough values of periapsis altitude, apoapsis altitude and inclination by choosing a form of orbit profile matching to the evaluation function.

© 2015, American Institute of Aeronautics and Astronautics Inc. All rights received. This study explores many-objective trajectory designs for the future low-thrust space- craft proposed in ISAS/JAXA. The various trajectory profiles are identified by multiobjective evolutionary algorithm (MOEA) for many-objective optimization. One difficulty of the low-thrust transfer problem is that there are many possible trajectory profiles due to many revolutions during the orbit raising phase. The objective functions of the low-thrust transfer problem are designed to minimize (1) the operation time of the Ion Engine System, (2) the flight time to reach the lunar orbit, (3) the maximum eclipse time and to maximize (4) initial mass of the spacecraft. For improving the performance of the evolutionary computation with large populations, CHEbyshev-Epsilon opTimizer AlgoritHm (CHEETAH) is adopted. The CHEETAH combines ∊-indicator and Chebyshev achievement function into a ranking method, and is designed for the parallel evaluation. The parallel CHEETAH with large populations is conducted at "K" supercomputer. The analysis of nondominated solutions reveals various trade-off relations and correlations among the objective functions. Furthermore, the analysis results provide useful knowledge to design trajectory profiles in the low-thrust spacecraft with the ion engine.

This study investigates the effective orbit maintenance maneuver for the periodic orbits around collinear equilibrium point (Lagrangian point) in Elliptic Restricted Three-Body Problem (ER3BP) using the Dynamical Systems Theory (DST). The assumed state displacement is stabilized by nullifying the unstable relative motion. The other displacement, which will not be deviated from the reference are permitted in this maneuver plan. The periodic orbits in the ER3BP are considered as multiple revolution orbit and, in total, by choosing the invariant unstable direction to calculate the stabilizing maneuver reduces control budget from choosing the unstable direction of each revolution obtained.

We investigate new ideas of direct transcription method for modeling low thrust trajectories as an improvement to the Sims-Flanagan model. First, the sequential impulsive AV transcription is replaced by integrated continuous thrust arcs to improve the fidelity of the dynamical model. Next, to enable the simulation of trajectories with multiple revolutions, we adopt a transformation in the independent variable from time to true anomaly. The obtained new algorithm is able to produce an operational trajectory accounting for the real spacecraft dynamics and adapting the segment duration on-line improving the final trajectory optimality.

Halo orbit missions are of many applications and become popular. An investigation on the extended mission following halo orbit missions would be worthwhile. In a previous study, the strategy of using the unstable manifolds associated with the Sun-Earth L-1/L-2 halo orbit and lunar gravity assists for Earth escape was analyzed to be advantageous for extending the mission. However, in an extension mission where the halo orbit mission is not pre-phased for a lunar swingby, the fuel cost for phasing the halo-to-Moon transfers should be investigated. The current paper aims to give the insight of the minimum phasing AV to encounter the Moon for various lunar phases with respect to the halo orbit. Efforts are made to tackle the problem of multiple optimization directions. The phasing planning is briefly discussed as well.

© 2015 IAA.Published by Elsevier Ltd. All rights reserved. In this study, we propose a flexible orbit design method that enables anytime launch of a deep-space explorer. Based on the Electric Delta-V Earth Gravity Assist (EDV-EGA) scheme, (Kawaguchi, 2001, 2002) [1,2] the proposed interplanetary parking method enables the explorer to make an Earth return orbit at an arbitrary time-of-flight by connecting to the minimum energy transfer orbit to destination. While the time-of-flight of the transfer orbit is fixed, the Earth return orbit with the arbitrary time-of-flight significantly alleviates the severe launch window constraint in interplanetary missions. We offer two case examples of applications of this method. The first is the dual launch of a Mars explorer with a geostationary transfer orbit (GTO) mission payload. The second is a dual launch of Mars and Venus explorers by a single launch vehicle. In the first case, we assume that a small Mars explorer is dual launched into a GTO for a secondary payload. With this assumption, the secondary payload cannot choose a desirable launch epoch for itself because the launch window to Mars is very narrow and opens only every 2 years. Moreover, the GTO, whose orbital period is approximately 10 h, repeatedly passes through the Van Allen belt wherein the radiation level is very high. Hence, the explorer has to escape from the GTO as soon as possible. However, our proposed interplanetary parking method enables the explorer to reach the destination within the limits of a practical mass resource, regardless of the Earth departure epoch. In the second case, the explorers traveling to different destinations, i.e., Mars and Venus, are dual launched by a single launch vehicle, and they fly to each destination via an interplanetary parking orbit. Our proposed method will widen the scope of opportunity for interplanetary missions.

Copyright © 2014 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. A B-type near-Earth asteroid, (3200) Phaethon, is the parent body of the Geminid meteor shower. Unlike most parent bodiesof meteor showers, Phaethonisdynamicallyan asteroid with few cometary features. Asteroids (155140) 2005 UD and (225416) 1999 YC are likely fragments originating from Phaethon, collectively known as the Phaethon-Geminid complex. A mission to this group could provide key information on their origins and solve fundamental issues in thermal and dynamic evolution of comet-asteroid transition bodies. This study assesses the feasibility of a multiple flyby mission for Phaethon, 2005 UD, and 1999 YC by a small-class mission. The objective is to design a simple multiple flyby mission based on ballistic transfers combined with gravity-assisted maneuvers that fly by some or all members of the Phaethon-Geminid complex. The results show periodic launch opportunities to all three asteroidswith the best case for Phaethon requiring less than1 km/s of Earth excess velocity.Nodirect transfer can be made to 1999 YC with less than 4 km/s. However, with a gravity-assist maneuver at Mars, an Earth-Mars-1999 YC transfer requires less than 3 km/s. It isalso found that, with a maximum of 3 km/s, there isnot a single transfer that connects all asteroids. However, launch windows in the years 2026 and 2027 allow a flyby of Phaethon and later 2005 UD by conducting an Earth gravity-assist maneuver.

© 2015 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. This paper presents a method of trajectory optimization based on the well-known primer vector theory, which is modified to accommodate weights in the cost function. This change arises from the need of a fast and accurate analysis obtained with an indirect method that takes into account the velocity increment used for departure from the planet and, particularly for flyby missions, the disregard of the last rendezvous impulse.Adetailed derivation of the weighted cost function and its gradient is presented, followed by a discussion on the values of the weights specifically for flyby and rendezvous missions. To test the optimization method, realistic test cases are selected and their results compared against a trajectory using the solution of the Lambert problem and optimization by a nonlinear programming solver. The proposed method showed a fast design of a trajectory with a midcourse impulse, which costs less than the trajectories calculated by the other two methods.

© 2015, American Astronautical Society. Akatsuki (“dawn” in Japanese) is the JAXA Venus orbiter that was scheduled to enter orbit around Venus on Dec. 7th, 2010. Following the failure of the main engine during the orbit insertion maneuver, the spacecraft escaped Venus on a 200-day orbit around the Sun, only to return in early 2017. This paper presents the design and implementation of the recovery trajectory, which involves perihelion maneuvers to re-encounter Venus in late 2015. Relying only on the onboard propellant, the trajectory rescued the mission by (1) anticipating the beginning of the science phase within the nominal lifetime of the spacecraft, and (2) halving the Δv requirements for the orbit insertion maneuver. Several trajectories are designed with an innovative use of a technique called non-tangent V-Infinity Leveraging Transfers (VILTs). Candidate solutions are then recomputed in higher fidelity models, and one solution is finally selected for its low Δv requirements and for programmatic reasons. The results of the perihelion maneuver campaign are also presented.

PROCYON (PRoximate Object Close flyby with Optical Navigation) is a 50kg-class micro-spacecraft developed by the University of Tokyo and the Japan Aerospace Exploration Agency (JAXA), to be launched in an Earth resonant trajectory at the end of 2014 as a secondary payload with Hayabusa 2 mission. The mission objective is to demonstrate low cost and applicability of a micro-spacecraft bus technology for deep space exploration and proximity flyby to asteroids performing optical navigation. This paper introduces the spacecraft and mission design for PROCYON, as well as, the operation strategy mainly for the deep-space cruising period

PROCYONは,超小型衛星による小惑星フライバイ探査ミッションを目的として,はやぶさ2のピギーバック搭載機会である2014年12月の打ち上げを目指している.超小型衛星による深宇宙探査ミッションは,大型衛星とは異なる信頼性基準,コストのバランスによって成立させる必要がある.特に搭載重量や発生電力の制約条件は大きく,従来の宇宙用通信コンポーネントの設計概念を大きく変えて積極的な民生部品の活用と小型軽量化に最適な技術の導入が不可欠である.本稿では,PROCYON通信系構成,並びに,各コンポーネントの詳細を紹介する.

Copyright ©2014 by the International Astronautical Federation. All rights reserved. PROCYON (PRoximate Object Close flY by with Optical Navigation) is world's first mission aimed to demonstrate the technology of a micro spacecraft deep space exploration and proximity flyby to asteroids. The mission is developed by the University of Tokyo in collaboration with ISAS, JAXA. The spacecraft is scheduled to be launched as a secondary payload in late 2014 with Hayabusa 2 spacecraft. PROCYON will first target back to the Earth using its miniature ion engine; then it will transfer to the target asteroid using Earth gravity assist; finally it will use optical navigation to perform proximity flyby of the asteroid. Due to the very low thrust and limited propellant of the mission, it is therefore important to ensure that the mission objective and requirements can still be satisfied under different conditions and parameters. In this paper, we present the results of a broad sensitivity study of PROCYONs trajectory due to various launch dates and mission parameters.

Copyright © 2014 by the International Astronautical Federation. All rights reserved. Recently with new trajectory design techniques and use of low-thrust propulsion systems, missions have become more efficient and cheaper with respect to propellant. As a way to increase the mission's value and scientific return, secondary targets close to the main trajectory are often added with a small change in the transfer trajectory. Due to their large number, importance and facility to perform a flyby, asteroids are commonly used as such targets. However, particularly for low-thrust mission, the trajectory design can be long and complicated requiring not only to optimize the fuel consumption but also the optical navigation time. Usually, design of trajectories to secondary targets are made through non-linear programming gradient based direct methods with a single objective function. In this work, the indirect method known as Primer Vector theory is used to define the direction and magnitude of the thrust for a minimum mass problem. The control law defined by the Primer Vector is implemented into the system's dynamics and genetic algorithm using optical navigation time as the objective function is applied to optimize the problem. This procedure allows for a global search, solving the problem's constraints while maximizing the optical navigation time and spacecraft's final mass, along with the asteroid flyby. Results are presented for a Galileo like mission using an ionic propulsion system.

The paper investigates the escape strategy using the unstable manifolds of the Sun-Earth L2 halo orbit and lunar gravity assist. There are four cases of intersection of the manifold tubes associated with halo orbits and the orbit of the Moon. The four intersections have different V<inf>∞</inf> with respect to the Moon. The corresponding lunar gravity assists can result in a range of escape trajectories, granting choices for the extended mission of halo orbits. In order to satisfy the lunar encounter requirements, the strategy and ΔV costs of phasing maneuvers are presented as well.

© 2014 SPIE. SPICA (Space Infrared Telescope for Cosmology and Astrophysics) is an astronomical mission optimized for mid- and far-infrared astronomy with a 3-m class telescope which is cryogenically cooled to be less than 6 K. The SPICA mechanical cooling system is indispensable for the mission but, generates micro-vibrations which could affect to the pointing stability performances. Activities to be undertaken during a risk mitigation phase (RMP) include consolidation of micro-vibration control design for the satellite, as well as a number of breadboarding activities centered on technologies that are critical to the success of the mission. This paper presents the RMP activity results on the microvibration control design.

© 2014 SPIE. We present 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 is expected to achieve 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. We have carried out the "Risk Mitigation Phase" activity, in which key technologies essential to the realization of the mission have been extensively developed. Consequently, technical risks for the success of the mission have been significantly mitigated. Along with these technical activities, the international collaboration framework of SPICA had been revisited, which resulted in maintenance of SPICA as a JAXA-led mission as in the original plan but with larger contribution of ESA than that in the original plan. To enable the ESA participation, a SPICA proposal to ESA is under consideration as a medium-class mission under the framework of the ESA Cosmic Vision. The target launch year of SPICA under the new framework is FY2025.

Japanese Venus Climate Orbiter/AKATSUKI was proposed in 2001 with strong support by international Venus science community and approved as an ISAS (The 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' deployment was confirmed. After a successful cruise, the malfunction happened on the propulsion system during the Venus orbit insertion (VOI) on Dec. 7, 2010. The engine shut down before the planned reduction in speed to achieve. The spacecraft did not enter the Venus orbit but 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 and unusable. However, we have found an alternate way of achieving orbit by using only the reaction control system (RSC). We had adopted the alternate way for orbital maneuver and three minor maneuvers in Nov. 2011 were successfully done so that AKATSUKI would meet Venus in 2015. We are considering several scenarios for VOI using only RCS. © 2013 IAA.

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. © 2013 2013 California Institute of Technology.

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. © 2013 2013 California Institute of Technology.

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 J2 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. © 2013 2013 California Institute of Technology.

This paper presents a method of trajectory optimization based on a well known theory, primer vector, which is modified to accommodate weights in the cost function; arising from the need of a more accurate analysis that takes into account the velocity increment used for the planet's departure and, particularly for flyby missions, the disregard of the last rendezvous impulse. A detailed derivation of the weighted cost function and its gradient is presented, followed by a discussion on the adjustment of the weights specifically for flyby missions. In order to test the optimization method, a realistic test case is selected and its results are compared against a trajectory using the solution of the Lambert problem and classical primer vector optimization. It is possible to clearly evaluate from the results the advantages of the proposed method, enabling to design a trajectory with a midcourse impulse which costs less than the trajectories calculated by the other methods and provides a more realistic analysis of the planet's departure.

Akatsuki is the JAXA Venus orbiter that was scheduled to enter a closed orbit around Venus on Dec. 7th, 2010. However, because of a failure of the main engine during the orbit insertion maneuver, the spacecraft escaped Venus on a 200-day orbit around the Sun, and if no deep space maneuvers were performed, would return close to Venus only in early 2017. This paper presents faster and safer recovery options, including the trajectory that was eventually selected by the mission team and that Akatsuki is currently flying. The selected trajectory involves a large perihelion maneuver to re-encounter Venus in late 2015. The maneuver also reduces the cost of the new orbit insertion opportunity, so that no Δv penalty is incurred. The recovery trajectories presented in this paper are computed first in the two-body problem model, assuming a circular orbit for Venus and a coplanar orbit for the spacecraft; several options are quickly computed using a technique called non-tangent V-Infinity Leveraging Transfers (VILTs). Candidate solutions are then recomputed in higher fidelity models, and one solution is finally selected for its low Av requirements and for programmatic reasons. The results of the perihelion maneuver campaign are also presented.