川勝 康弘

Martian moons exploration, MMX, is the new sample return mission planned by the Japan Aerospace Exploration Agency (JAXA) targeting the two Martian moons with the scheduled launch in 2024 and return to the Earth in 2029. The major scientific objectives of this mission are to determine the origin of Phobos and Deimos, to elucidate the early Solar System evolution in terms of volatile delivery across the snow line to the terrestrial planets having habitable surface environments, and to explore the evolutionary processes of both moons and Mars surface environment. To achieve these objectives, during a stay in circum-Martian space over about 3 years MMX will collect samples from Phobos along with close-up observations of this inner moon and carry out multiple flybys of Deimos to make comparative observations of this outer moon. Simultaneously, successive observations of the Martian atmosphere will also be made by utilizing the advantage of quasi-equatorial spacecraft orbits along the moons' orbits.

The science operations of the spacecraft and remote sensing instruments for the Martian Moon eXploration (MMX) mission are discussed by the mission operation working team. In this paper, we describe the Phobos observations during the first 1.5 years of the spacecraft's stay around Mars, and the Deimos observations before leaving the Martian system. In the Phobos observation, the spacecraft will be placed in low-altitude quasi-satellite orbits on the equatorial plane of Phobos and will make high-resolution topographic and spectroscopic observations of the Phobos surface from five different altitudes orbits. The spacecraft will also attempt to observe polar regions of Phobos from a three-dimensional quasi-satellite orbit moving out of the equatorial plane of Phobos. From these observations, we will constrain the origin of Phobos and Deimos and select places for landing site candidates for sample collection. For the Deimos observations, the spacecraft will be injected into two resonant orbits and will perform many flybys to observe the surface of Deimos over as large an area as possible.

© 2020 COSPAR Ballistic landers enable orbiting asteroid missions to perform surface science at limited additional cost and risk. Due to asteroids’ weak gravity and irregular terrain, lander deployment trajectories will consist of several chaotic bounces. Although impacts on regolith-covered asteroids are numerically expensive to model, impacts on rocky asteroids can be modeled with simpler, impulsive contact models. One such model is that by Stronge, which was successfully used in large-scale Monte Carlo studies of asteroid lander deployment. This model parameterizes impacts with (fixed) material restitution and friction coefficients, but has not been validated for the low-velocity regime of an assembled, nonspherical body. This paper uses an air-bearing setup to perform 2D experiments of a rectangular floating assembly impacting a concrete block with V⩽25 cm/s. The impact velocity, assembly attitude, and block attitude are varied across 2,400 experimental runs of both normal and tangential impacts. Optical tracking is used to extract the pre- and post-impact velocities of the assembly. In a majority of cases, Stronge's model can be fit to the experiments to extract the corresponding restitution and friction coefficients. We find that the coefficients are not fixed with respect to the impact velocity and attitude, but that their variation is seemingly random. In some tangential impact cases, the model even fails to reproduce the observed behavior althogether. This suggests that there may not be a simple way to reconcile Stronge's fixed-material-coefficient model with reality, although it may retain practical use if the coefficients are randomly varied in each impact of a simulation.

Martian Moons eXploration (MMX) is a mission to Martian moons under development in JAXA with international partners to be launched in 2024. This paper introduces the system definition and the latest status of MMX program. “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 in the context of our minor body exploration strategy preceded by Hayabusa and Hayabusa2. Solar system formation theories suggest that small bodies as comets and asteroids were delivery capsules of water, volatiles, organic compounds etc. from outside of the snow line to entitle the rocky planet region to be habitable. 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. The mission is to survey two Martian moons, and return samples from one of them, Phobos. In view of the launch in 2024, the phase-A study was completed in February, 2020. The mission definition, mission scenario, system definition, critical technologies and programmatic framework are introduced int this paper.

© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Solar electric propulsion is a key enabling technology that has improved the efficiency of space transport. With specific impulses that are typically ten times higher than the chemical counterpart, electric motors allow a considerable saving in propellant mass at the expense of longer times of flight. However, the length of the transfer process and the specific operational needs require to develop a different operational concept for the navigation and orbit control that can be sustained during the different phases of the mission. In this paper, a trade-off is performed among several operational concepts and solutions for multi-revolutions SEP transfers with application to the DESTINY+ mission. The GTO-to-Moon low-thrust transfer is first computed in a high-fidelity model with a tangential thrust strategy and later optimized with a five-order Legendre-Gauss-Lobatto collocation method. The impact of eclipses, radiation, thrust outages and misfires, and orbit tracking is analyzed in detailed and included in the transcript optimal problem as algebraic constraints where possible. Numerical results show that the driving factors for the optimal trajectory are related to the operations of the spacecraft rather than the final mass or time of flight.

© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Planetary orbit insertion maneuverer is one of the most critical events in planetary exploration missions. Once spacecraft fails to decelerate enough to stay in the planetary orbit, the spacecraft escape from the vicinity of the planet. In some cases, such as JAXA’s Akatsuki mission, the feasible recovery trajectory was found after engine malfunction; however, in many cases, the mission may fail because the spacecraft will end its mission life before re-encountering the planet. To mitigate the risk, researchers have studied robust orbit insertion strategies, which reduces the total probabilistic ∆V by sacrificing nominal ∆V and shortens the flight time to re-encounter the planet. The previous studies consider the possibility that POI fails completely and use the free-return trajectories as the recovery trajectories. This paper extends the robust orbit insertion method by implementing V-infinity leveraging transfer as the recovery trajectories. This extension enables to treat general probability models that describe the uncertainties of orbit insertion maneuvers. The numerical example of JAXA’s Martian Moons eXploration mission shows that proposed method enhance performance, such as the expected ∆V and ∆V99, as well as robustness.

© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Stable Quasi-Satellite Orbits (QSOs) have gained a lot of attention as suitable candidate orbits to explore remote planetary moons. Despite many studies on QSO and its orbital stability are found in literature, the problem of efficient transfer between two QSO is still unsolved. Previous works on transfers between two QSOs includes transfers between planar QSOs using impulsive maneuvers and utilizing the bifurcated multiple-revolutional periodic QSOs (MP-QSOs). Purpose of this paper is to explore the three-dimensional case and design transfer trajectories from lower altitude QSOs to three-dimensional QSOs being considered for upcoming sample return missions like Martian Moons eXploration (MMX) and PHOOT-PRINT. Specifically, we utilize the unstable 3D QSO to generate manifolds and connect them to the planar QSO in the framework of Mars-Phobos Circular Hill Problem with Ellipsoidal secondary (HPE). Numerical simulations suggest that transfer between a planar and spatial QSO around Phobos is possible with ∆V as low as 8.277 m/s with TOF of 4.19 days and similarly transfer between a spatial and planar QSO with ∆V of 8.286 m/s with TOF of 6.75 days. As a result, transfer design space between planar and spatial QSOs is established using the invariant manifolds.

© 2020, The Author(s). Quasi-satellite orbits are of great interest for the exploration of planetary moons because of their dynamical features and close proximity with respect to the surface of scientifically relevant objects like Phobos and Deimos. This paper explores the equations of the elliptical Hill problem, offering a new analytical insight into the long-term evolution of mid-altitude quasi-satellite orbits. Our developments are based on the Yamanaka–Ankersen solution of the Tschauner–Hempel equations and capture the effects of the secondary’s gravity and orbital eccentricity on the shape and orientation of near-equatorial retrograde relative trajectories. The analytical solution of the in-plane and out-of-plane components of the secular motion is achieved by averaging over the relative longitude of a spacecraft as seen from the co-rotating frame of the two primaries. Developments are validated against the numerical integration of quasi-periodic trajectories that densely cover the surface of three-dimensional invariant tori. This analysis confirms the stable nature of quasi-satellite orbits and provides new tools for future spacecraft missions such as the Martian Moons eXploration envisaged by JAXA.

© 2020, American Astronautical Society. This paper presents the trajectory design for EQUilibriUm Lunar-Earth point 6U Spacecraft (EQUULEUS), which aims to demonstrate orbit control capability of CubeSats in the cislunar space. The mission plans to observe the far side of the Moon from an Earth-Moon L2 (EML2) libration point orbit. The EQUULEUS trajectory design needs to react to uncertainties of mission design parameters such as the launch conditions, errors, and thrust levels. The main challenge is to quickly design science orbits at EML2 and low-energy transfers from the post-deployment trajectory to the science orbits within the CubeSat’s limited propulsion capabilities. To overcome this challenge, we develop a systematic trajectory design approach that 1) designs over 13,000 EML2 quasi-halo orbits in a full-ephemeris model with a statistical stationkeeping cost evaluation, and 2) identifies families of low-energy transfers to the science orbits using lunar flybys and solar perturbations. The approach is successfully applied for the trajectory design of EQUULEUS.

Copyright © 2019 by the International Astronautical Federation (IAF). All rights reserved. Martian Moons eXploration (MMX) is a mission to Martian moons under study in JAXA with international partners to be launched in 2024. This paper introduces the mission definition and the latest status of MMX program. “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 in the context of our minor body exploration strategy preceded by Hayabusa and Hayabusa2. Solar system formation theories suggest that small bodies as comets and asteroids were delivery capsules of water, volatiles, organic compounds etc. from outside the snow line to entitle the rocky planet region to be habitable. 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. The mission is to survey two Martian moons, and return samples from one of them. In view of the launch in 2024, the phase-A study is to be completed in this year. The mission definition, mission scenario, system description, and programmatic framework are introduced int this paper.

<p>Failure of the orbit insertion maneuver has a significant impact on the entire mission, for the trajectory of a spacecraft is largely deflected by swing-by. The risk can be reduced by targeting a point on the B-plane where the spacecraft reaches the free-return (FR) trajectory with the target body in the case of insertion failure. The backup orbit must also satisfy conditions suitable for the mission. We investigated the type of orbit insertion that is both robust to failure and reasonable for the mission requirements. We call this method FR ensured orbit insertion. Among various failure modes of the orbit insertion maneuver, we focus on the complete maneuver failure. The impact parameters on the B-plane to achieve the orbit insertion are formulated based on the geometry of velocity vectors at swing-by. The necessary deflection angle α_FR at swing-by must be smaller than the possible maximum deflection angle α_max for the target body. When we introduce proper scaling factors, the relation of α_max and α_FR is characterized by a single parameter λ. Using polar orbit insertion as an example, maps which show the reachability of FR trajectory after the insertion failure for each approaching condition are presented. The derived maps can be used as a tool to assess the applicability of the method in the mission design. Finally, as an application to practical mission design, we demonstrate the use of FR ensured orbit insertion in JAXA's MMX mission.</p>

Copyright © 2019 by the International Astronautical Federation (IAF). All rights reserved. DESTINY+ (Demonstration and Experiment of Space Technology for INterplanetary voYage, Phaethon fLyby and dUSt analysis) is a small-sized high-performance deep space vehicle proposed at ISAS/JAXA. The trajectory design of DESTINY+ is divided into several phases. First phase is an orbit injection into an extended elliptical orbit launched by the Epsilon rocket with the additional solid kick motor. Second phase is many revolutions transfer to raise apogee altitude by low thrust propulsion system to the moon orbit nearby. And at third phase, the distant flyby and the swing-by around the moon is designed to give DESTINY+ momentum to escape Earth gravitational field. At an interplanetary phase, DESTINY+ goes to an Asteroid Phaethon for flyby observation. After the Phaethon flyby, DESTINY+ is planned to go back toward Earth for gravity assist and go to another asteroid 2005UD which thought to have split from Phaethon. This paper discusses DESTINY+'s low-thrust trajectory design. As for the many revolution transfer 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 of time of radiation belt, the work time of low thrust propulsion system and the maximum eclipse period. After the spacecraft reaches to the moon's orbit, it utilizes the moon swing-by several times to connect to the transfer trajectory for Asteroid Phaethon. From these studies, we can show the feasibility of the mission design of DESTINY+,.

© 2019, Univelt Inc. All rights reserved. The candidate science orbits for two JAXA missions heading to three-body system, EQUULEUS and MMX, are presented in this paper. Both of these missions need to conduct science observations while coping with tight engineering constraints such as thermal and power budgets. Due to these requirements, eclipses may become a significant issue for both missions. To minimize or avoid eclipses, we introduce key design parameters for synodic resonant periodic orbits: the synodic ratio and the elongation angle between the Sun and the two primaries. By combining these parameters, we can obtain science orbits that avoid or minimize eclipses in a full ephemeris model. This approach is demonstrated for the science orbit design of both EQUULEUS and MMX.

© 2019, Univelt Inc. All rights reserved. This paper explores the application of multi-revolutional periodic orbits to transfer between single-revolutional quasi-satellite orbits around the Martian moon Phobos. Multi-revolutional periodic orbits are retrograde trajectories that repeat after multiple revolutions around Phobos. At first, we generate them by conventional predictor-corrector scheme and bifurcation analysis and find many candidate options for trajectory design analyses. Next, we explore transfer solutions between different single-revolutional periodic quasi-satellite orbits. We find that, if we choose appropriate multi-revolutional periodic orbits, we can reduce the transfer time and ΔV, as well as increase the robustness of the transfers from a contingency analysis standpoint.

© 2019, Univelt Inc. All rights reserved. Analysis of tube dynamics under two dimensions has been analyzed by Koon et al.2 However, analysis in three dimensions has not been done sufficiently. In this work, we discuss on designing Mars escape trajectory in the three-body system via invariant manifolds. In contrast with the conventional methods, we propose to use invariant manifolds from quasi-periodic invariant tori and investigate quasi-halo orbits and their associated tubes. Finally, we show how the invariant manifolds of quasi-periodic orbits can provide new escape trajectories from the Martian system in the circular restricted three-body problem.

Copyright © 2019 by the International Astronautical Federation (IAF). All rights reserved. The Martian Moon eXploration mission, currently under development by JAXA, will be launched in 2024 with the goal of retrieving pristine samples from the surface of Phobos. Soon after arrival, the spacecraft will inject into retrograde relative trajectories known as quasi-satellite orbits and study the geophysical environment of the Martian satellite for more than three years. This paper presents the orbit design and maintenance strategy of MMX in the framework of the elliptical Hill problem with ellipsoidal secondary. Specifically, we introduce a numerical continuation procedure on the eccentricity of Phobos to generate families of two-dimensional quasi-periodic invariant tori that replace purely periodic solutions that exist in the circular case. Two-dimensional torus maps can be then constructed and used to represent physical quantities of interest (e.g., spacecraft longitude and altitude), as well as to generate reference trajectories for tracking purposes. Sensitivity and stability analyses are carried out to investigate the dynamical properties of retrograde relative trajectories in the time-periodic case. Finally, a Linear Quadratic Regulator is implemented in order to assess the robustness of the computed trajectories under injection, navigation, and execution errors. Monte Carlo simulations demonstrate that all of the computed quasi-satellite orbits can be maintained with as low as 18.065 m/s per month.

Copyright © 2019 by the International Astronautical Federation (IAF). All rights reserved. In this paper, we propose a new method of designing efficient escape trajectories from a gravity field of a planet. In particular, we study to design escape trajectories from the Martian moon Phobos with these processes in the context of the three-dimensional Sun-Mars-Spacecraft Circular Restricted Three-Body Problem (CR3BP). Our method consists of two design steps for realizing low-energy transit trajectories. The first step is to design reference trajectories escaping from a vicinity of Phobos. In this step, we use a halo orbit as a hub, and numerically propagate trajectories which is along both stable and unstable manifolds. Each stable and unstable invariant manifold asymptotically approaches to a halo orbit forward and backward in time respectively. Therefore, it is possible to systematically design transit trajectories passing through the vicinity of the halo orbit with some lower energy using the properties of the invariant manifolds. The second step is to modify such designed trajectories to reduce its thrust (?V) of departing for practical missions. Therefore, we apply a method called differential correction to renew trajectories by iterating analytical approximations. We target to states of a spacecraft whose ?V is to be lower, and the number of thrusting a spacecraft is reduced in order to improve its mission operability. Finally, we illustrate that we can obtain the efficient escaping trajectories from the Mars vicinity.

© 2019, The Author(s), under exclusive licence to Springer Nature Limited. The surface of the Martian moon Phobos exhibits two distinct geologic units, red and blue, characterized by their spectral slopes. The provenance of these units is uncertain yet crucial to understanding the origin of the Martian moon and its interaction with the space environment. Here we present a combination of dynamical analyses and numerical simulations of particle dynamics to show that periodic variations in dynamic slopes, driven by orbital eccentricity, can cause surface grain motion. For regions with steep slopes that vary substantially over one Phobos orbit, the surface is excavated at a faster rate than the space weathering timescale. Our model predicts that this new mechanism is most effective in regions that coincide with blue units. Therefore, space weathering is the likely driver of the dichotomy on the moon’s surface, reddening blue units that represent pristine endogenic material.

© 2019 Elsevier Ltd Landing on Phobos and bringing samples from its surface would settle the debate on the origin of the Martian moons and support future manned exploration to Mars. To fulfill these scientific objectives, JAXA is planning to send a sample return probe to Phobos by the first half of the next decade, named the Martian Moons eXploration (MMX) mission. In order to explore scientifically interesting regions of Phobos, as well as to support the sampling operations of MMX, a number of Deployable CAMera 5 payloads are proposed to be deployed from quasi-satellite orbits (QSOs) around the Martian moon. This paper explores the feasibility of ballistic deployments from QSOs under realistic dynamical environment and surface constraints in order to guarantee surface settlement within the lifespan of DCAM5. First, we analyze the dynamical environment and escape speeds from Phobos by means of the Circular Hill Problem. Then, the surface coefficient of restitution is estimated by generic impacts onto Phobos regolith via discrete element method simulations. By combining these two analyses, maximum allowable impact velocities for surface settling are calculated and applied to downselect the number of feasible ballistic landings from QSOs. It is found that access to Phobos surface is possible especially along the leading and trailing sides of the Martian moon and in agreement with the engineering requirements of DCAM5.

The candidate science orbits for two JAXA missions heading to three-body system, EQUULEUS and MMX, are presented in this paper. Both of these missions need to conduct science observations while coping with tight engineering constraints such as thermal and power budgets. Due to these requirements, eclipses may become a significant issue for both missions. To minimize or avoid eclipses, we introduce key design parameters for synodic resonant periodic orbits: the synodic ratio and the elongation angle between the Sun and the two primaries. By combining these parameters, we can obtain science orbits that avoid or minimize eclipses in a full ephemeris model. This approach is demonstrated for the science orbit design of both EQUULEUS and MMX.

This paper explores the application of multi-revolutional periodic orbits to transfer between single-revolutional quasi-satellite orbits around the Martian moon Phobos. Multi-revolutional periodic orbits are retrograde trajectories that repeat after multiple revolutions around Phobos. At first, we generate them by conventional predictor-corrector scheme and bifurcation analysis and find many candidate options for trajectory design analyses. Next, we explore transfer solutions between different single-revolutional periodic quasi-satellite orbits. We find that, if we choose appropriate multi-revolutional periodic orbits, we can reduce the transfer time and Delta V, as well as increase the robustness of the transfers from a contingency analysis standpoint.

Martian Moons eXploration (MMX) mission focus on the Martian moons. The spacecraft will make close-up remote sensing and in-situ observations of both moons, and collect a sample from one of the moons to bring back to Earth.

Analysis of tube dynamics under two dimensions has been analyzed by Koon et al.(2) However, analysis in three dimensions has not been done sufficiently. In this work, we discuss on designing Mars escape trajectory in the three-body system via invariant manifolds. In contrast with the conventional methods, we propose to use invariant manifolds from quasi-periodic invariant tori and investigate quasi-halo orbits and their associated tubes. Finally, we show how the invariant manifolds of quasi-periodic orbits can provide new escape trajectories from the Martian system in the circular restricted three-body problem.

第28回アストロダイナミクスシンポジウム (2018年7月30-31日. 宇宙航空研究開発機構宇宙科学研究所), 相模原市, 神奈川県資料番号: SA6000135060レポート番号: C-7

Copyright © 2018 by the International Astronautical Federation. EQUULEUS is a Lunar L2 orbiter and a 6-Unit CubeSat by JAXA and the University of Tokyo. OMOTENASHI is a 6-Unit CubeSat by JAXA, the world's smallest Lunar lander. EQUULEUS and OMOTENASHI are among the 13 secondary payloads selected by NASA to be launched with Exploration Mission-1 in 2019. Despite their limited size and cost, EQUULEUS and OMOTENASHI are challenging missions, especially in terms of trajectory design and control. EQUULEUS exploits the Earth-Sun-Moon chaotic dynamics and enter a libration point orbit around the L2 of the Earth-Moon system, using a new water propulsion system with low thrust and little propellant. This “Orbit Control Experiment” is one of the main objectives of the mission. OMOTENASHI executes a semi-hard landing that requires breaking the spacecraft to a stop just a few-hundred meters above the Moon's surface. Both missions present new and unique challenges, where the design of the nominal trajectory is mainly driven by the constrains on orbital control capabilities, and operational and robustness considerations. This paper presents the current baselines, and give an overview of the new techniques developed for their design.

© 2018 Univelt Inc. All rights reserved. DESTINY+ flies by asteroid named Phaethon at a speed of 33 km/s. At flyby phase, it is planned to image Phaethon by 2 kinds of camera. To take good quality images in terms of science, it is necessary that the view of cameras continue to catch Phaethon. In this paper, to achieve this mission, how to track Phaethon and how to control the view of camera during flyby are showed. And it is indicated by numerical analysis that DESTINY+ can take the image which is satisfied scientific request.

Copyright © 2018 by the International Astronautical Federation. All rights reserved. Small deployable landers are promising candidates to enhance the science return of minor body missions due to their relatively low cost, low complexity, and low risk of operation. These landers are generally passive and deployed on a ballistic trajectory, i.e. no trajectory control is available. Therefore, many challenges remain in the design of small body landers due to the complex dynamical environments near minor bodies as well as partially inelastic collisions with their surfaces. In this research, we introduce a novel mission design toolbox that combines orbital and attitude simulations in realistic gravitational environments with a state-of-the-art contact dynamics analyzer. The mission designer can derive and analyze scientific and engineering requirements for the mission based on the results of the toolbox analysis, including covariance ellipses on the final landing site and local regolith and gravity information. The capabilities of the toolbox are demonstrated with a feasibility study of a cylindrical and cuboid landers to be deployed on the surface of Phobos. We find that landing can be accomplished in both cases despite the high touchdown velocities imposed by the dynamics of the Martian system.

Copyright © 2018 by the International Astronautical Federation (IAF). All rights reserved. The Martian Moons eXploration mission is currently under development at JAXA and will be the first spacecraft mission to retrieve pristine samples from the surface of Phobos. In preparation for the sampling operations, MMX will collect observations of Phobos from stable retrograde relative trajectories also known as quasi-satellite orbits or QSOs. This paper offers a semi-analytical analysis of mid- and high-altitude QSOs in terms of relative orbit elements. Our analysis is not limited to planar orbits and takes into account the eccentricity of the moon's orbit. Furthermore, we introduce a numerical map between mean and osculating orbit elements to study the long-term evolution of MMX and derive a Lyapunov control law for orbit maintenance purposes. The nonlinear controller is based on mean relative orbit element differences and tested with respect to injection errors.

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. In the case of initial guess in the trajectory design to use the swing-by, Zero Patched Conics Method is often used. In this method, the sphere of influence of the swing-by target body is treated as a point. It can be that the trajectory is calculated by the two body problem with the central body and connected analytically. However, the trajectory obtained by this method often cannot be connected due to multi body problem. Therefore, in this study, we propose a method called “Improved Zero Patched Conics Method” to make a connectable pair of the trajectories before and after swing-by that conforms to multi-body problem. This method can improve the connection the trajectories before and after the swing-by in multi body problem.

© 2018 Univelt Inc. All rights reserved. Martian Moons eXploration (MMX) mission focus on the Martian moons. The spacecraft will make close-up remote sensing and in-situ observations of both moons, and collect a sample from one of the moons to bring back to Earth.

© 2018 Mars Moon eXplorer (MMX) is JAXA's next candidate flagship mission to be launched in the early 2020s. MMX will explore the Martian moons and return a sample from Phobos. This paper presents the mission analysis work, focusing on the transfer legs and comparing several architectures, such as hybrid options with chemical and electric propulsion modules. The selected baseline is a chemical-propulsion Phobos sample return, which is discussed in detail with the launch- and return-window analysis. The trajectories are optimized with the jTOP software, using planetary ephemerides for Mars and the Earth; Earth re-entry constraints are modeled with simple analytical equations. Finally, we introduce an analytical approximation of the three-burn capture strategy used in the Mars system. The approximation can be used together with a Lambert solver to quickly determine the transfer Δv costs.

© 2017, American Astronautical Society. This work explores the target selection and trajectory design of the mission candidate for ISAS/JAXA’s small science satellite series, DESTINY PLUS or 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 mission extension scenario are also explored.

Copyright © 2018 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 2024. This paper introduces the mission design 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. The mission is to survey two Martian moons, and return samples from one of them. Following the mission concepts study results presented in the previous conference, the following items will be reported in this paper. First, based on the mission goals and objectives defined, the requirements to the systems and operations are derived and their feasibility is evaluated. Second, as to the key technologies issues identified, partial models are built and their performance is evaluated. And third, collaborations with overseas space agency are discussed and the programmatic framework is defined.

© 2018 IAA This study proposes a solar sailing method for angular momentum control of the interplanetary micro-spacecraft PROCYON (PRoximate Object Close flYby with Optical Navigation). The method presents a simple and facile practical application of control during deep space missions. The developed method is designed to prevent angular momentum saturation in that it controls the direction of the angular momentum by using solar radiation pressure (SRP). The SRP distribution of the spacecraft is modeled as a flat and optically homogeneous plate at a shallow sun angle. The method is obtained by only selecting a single inertially fixed attitude with a bias-momentum state. The results of the numerical analysis indicate that PROCYON's angular momentum is effectively controlled in the desired directions, enabling the spacecraft to survive for at least one month without momentum-desaturation operations by the reaction control system and for two years with very limited fuel usage of less than 10 g. The flight data of PROCYON also indicate that the modeling error of PROCYON's SRP distribution is sufficiently small at a small sun angle (<10°) of the order of 10−9 Nm in terms of its standard deviation and enables the direction of the angular momentum around the target to be maintained.

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.

© Copyright 2017 by the International Astronautical Federation (IAF). All rights reserved. DESTINY+ (Demonstration and Experiment of Space Technology for Interplanetary Voyage, Phaethon Flyby and Dust Science) is a candidate of ISAS Epsilon class small program. The mission of DESTINY+ is to validate key technologies for our future deep space exploration. DESTINY+ will demonstrate the high performance electric propelled vehicle technology and execute the flyby exploration of asteroid 3200 Phaethon. DESTINY+ starts its voyage from a low elliptic orbit, spirals up the orbits, fly-by the Moon, escapes from the Earth, and depart for the asteroid 3200 Phaethon. It will detect and analyze interplanetary and interstellar dust particles during deep space cruise. This paper will introduce an overview of the DESTINY+ mission.

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, Springer Science+Business Media Dordrecht. 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. As a result of their large number, importance and facility to perform a flyby, asteroids are commonly used as such targets. This work uses the Primer Vector theory to define the direction and magnitude of the thrust for a minimum fuel consumption problem. The design of a low-thrust trajectory with a midcourse asteroid flyby is not only challenging for the low-thrust problem solution, but also with respect to the selection of a target and its flyby point. Currently more than 700,000 minor bodies have been identified, which generates a very large number of possible flyby points. This work uses a combination of reachability, reference orbit, and linear theory to select appropriate candidates, drastically reducing the simulation time, to be later included in the main trajectory and optimized. Two test cases are presented using the aforementioned selection process and optimization to add and design a secondary flyby to a mission with the primary objective of 3200 Phaethon flyby and 25143 Itokawa rendezvous.