尾崎 直哉

Asteroid exploration has been attracting more attention in recent years. Nevertheless, we have just visited tens of asteroids, whereas we have discovered more than 1 million bodies. As our current observation and knowledge should be biased, it is essential to explore multiple asteroids directly to better understand the remains of planetary building materials. One of the mission design solutions is utilizing asteroid flyby cycler trajectories with multiple Earth gravity assists. An asteroid flyby cycler trajectory design problem is a subclass of global trajectory optimization problems with multiple flybys, involving a trajectory optimization problem for a given flyby sequence and a combinatorial optimization problem to decide the sequence of the flybys. As the number of flyby bodies grows, the computation time of this optimization problem expands maliciously. This paper presents a new method to design asteroid flyby cycler trajectories utilizing a surrogate model constructed by deep neural networks approximating trajectory optimization results. Because one of the bottlenecks of machine learning approaches is the heavy computation time to generate massive trajectory databases, we propose an efficient database generation strategy by introducing pseudo-asteroids satisfying the Karush–Kuhn–Tucker conditions. The numerical result applied to Japan Aerospace Exploration Agency’s DESTINY+ mission shows that the proposed method is practically applicable to space mission design and can significantly reduce the computational time for searching asteroid flyby sequences.

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.

Recent low-thrust space missions have highlighted the importance of designing trajectories that are robust against uncertainties. In its complete form, this process is formulated as a nonlinear constrained stochastic optimal control problem. This problem is among the most complex in control theory, and no practically applicable method to low-thrust trajectory optimization problems has been proposed to date. This paper presents a new algorithm to solve stochastic optimal control problems with nonlinear systems and constraints. The proposed algorithm uses the unscented transform to convert a stochastic optimal control problem into a deterministic problem, which is then solved by trajectory optimization methods such as differential dynamic programming. Two numerical examples, one of which applies the proposed method to low-thrust trajectory design, illustrate that it automatically introduces margins that improve robustness. Finally, Monte Carlo simulations are used to evaluate the robustness and optimality of the solution.

Micro/nano-satellite technology development first started as either an educational or research tool primarily at universities and has spread rapidly across the world and found many practical applications. Recent technology advancement has enabled micro/nano satellites to become one of the platforms for deep space science and exploration missions. In order to provide an opportunity for students and researchers worldwide to propose a mission of deep space science and exploration using a micro/nano satellite, the 7th Mission Idea Contest (MIC) will focus on these missions. This is because the technological field of LEO satellites is already well-established, and so we consider that creation of a deep space mission will give the young generation more motivation towards the “Frontier.” MIC has provided aerospace engineers, college students, consultants, scientists, and anybody interested in space with opportunities to present their creative ideas and gain attention internationally. The 7th Mission Idea Contest (MIC7) is to open a door to new opportunities for proposing a deep space mission using a micro/nano satellite. There are several examples and emerging opportunities for micro/nano satellites in these types of missions. In this paper, we briefly introduce the past Mission Idea Contests and present examples of deep space missions using micro/nano satellites for MIC7, and follow this with a discussion on why it is important for UNISEC-Global to organize a “Mission Idea Contest for Deep Space Science and Exploration .

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.

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. 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.

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.

Autonomous orbit determination method using active maneuvers and intersatellite ranging between multiple spacecraft is applicable to general dynamics situations, but large uncertainty of information about maneuvers results in inaccurate orbit estimation. This paper proposes an accurate and robust estimation method using sequential filter, RTS smoother, and EM algorithm. Proposed method estimates not only states but also maneuver results. Results from simulations of Mars-Phobos system show that the proposed method improve orbits determination accuracy.

Temporarily-captured orbiters (TCOs) are a new population of asteroids that are temporarily gravitationally bound around the Earth-Moon system. Because of its small geocentric distance and energy, short-term exploration with small Ay is expected possible. This study aims to construct low-energy transfers to 2006 RH120, one of the TCOs, from low-Earth orbit using an analogy with the Earth-Moon low-energy transfers. The initial guess was sought by back-propagating perturbed 2006 RH120's trajectory, then it was optimized through the direct multiple shooting method. The result shows that various transfers are possible with rendezvous Ay below 100 m/s, and they form diverse family structures.

Recent low-thrust space missions have highlighted the importance of designing trajectories that are robust against uncertainties. In its complete form, this process is formulated as a nonlinear constrained stochastic optimal control problem. This problem is among the most complex in control theory, and no practically applicable method to low-thrust trajectory optimization problems has been proposed to date. This paper presents a new algorithm to solve stochastic optimal control problems with nonlinear systems and constraints. The proposed algorithm uses the unscented transform to convert a stochastic optimal control problem into a deterministic problem, which is then solved by trajectory optimization methods such as differential dynamic programming. Two numerical examples, one of which applies the proposed method to low-thrust trajectory design, illustrate that it automatically introduces margins that improve robustness. Finally, Monte Carlo simulations are used to evaluate the robustness and optimality of the solution.

Low-thrust propulsion is a key technology for space exploration, and much work in astrodynamics has focused on the mathematical modeling and the optimization of low-thrust trajectories. Typically, a nominal trajectory is designed in a deterministic system. To account for model and execution errors, mission designers heuristically add margins, for example, by reducing the thrust and specific impulse or by computing penalties for specific failures. These conventional methods are time-consuming, done by hand by experts, and lead to conservative margins. This paper introduces a new method to compute nominal trajectories, taking into account disturbances. The method is based on stochastic differential dynamic programming, which has been used in the field of reinforcement learning but not yet in astrodynamics. A modified version of stochastic differential dynamic programming is proposed, where the stochastic dynamical system is modeled as the deterministic dynamical system with random state perturbations, the perturbed trajectories are corrected by linear feedback control policies, and the expected value is computed with the unscented transform method, which enables solving trajectory design problems. Finally, numerical examples are presented, where the solutions of the proposed method are more robust to errors and require fewer penalties than those computed with traditional approaches, when uncertainties are introduced.

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.

Low-thrust propulsion is a key technology for space exploration, and much work in astrodynamics presents low-thrust trajectory design methods. Typically, a nominal trajectory is designed in a deterministic system. To account for model and execution errors, mission designers heuristically add margins - for example, by reducing the thrust and specific impulse or by computing penalties for specific failures. These conventional methods are time-consuming, done by hand by experts, and lead to conservative margins. This paper introduces a new method to compute nominal trajectories, taking into account disturbances. The proposed method, tube stochastic differential dynamic programming, is a modified algorithm of stochastic differential dynamic programming to handle the control constraints. The proposed algorithm, which is inspired by the tube model predictive control in the field of robotics, employs the sigma points to create a tube and computes the expected value by the unscented transform. Finally, numerical examples present that the proposed solutions automatically introduce the margin against uncertainty and therefore gain good robustness against uncertainties.

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.

Earth observation satellites can improve the flexibility of observation sites by having “maneuverability,” and low-thrust obtained by ion thruster will be a promising method for orbital change for micro-satellites. Designing low-thrust trajectories for these satellites is a multi-revolution and multi-objective (time/fuel-optimal) optimization problem, which usually requires high computational cost to solve numerically. This paper derives an analytical and approximate optimal orbit change strategy between two circular orbits with the same semi-major axis and different local time of ascending node, and proposes a graph-based method to optimize the multi-objective criteria. The optimal control problem results in a problem to search a switching point on the proposed graph, and mission designers can design an approximate switching point on this graph, by using two heuristic and reasonable assumptions that 1) the optimal thrust direction should be tangential to orbit and 2) the optimal thrust magnitude should be bang-bang control with an intermediate coast. Finally, numerical simulation with feedback control algorithm taking thrust margin demonstrates that the proposed method can be applicable in the presence of deterministic and stochastic fluctuation of aerodynamic disturbances.

© 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.

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.

This paper presents a method to design robust trajectories against possible faults of lover-thrust engines. Recently, increasing number of space probes with love-thrust engines have. been developed. Due to the insufficient reliability of low-trust propulsion system, almost all probes with ion engines have experienced the failure of engines. Conventional methods to design the low-thrust trajectory have pursued fuel minimum solution to only one target celestial body, which does not necessarily mean the maximization of mission achievement. In other words, it Can enhances accomplishment to intelligently change the trajectory when engine failures occur. In this research, the objective function is defined as expected scientific gain and the optimal solution is searched by a proposed method. Resulting method. Bellman Rapidly-exploring Random Trees (Bellman RRT) are efficiently extended by applying the Bellman equation. Finally, the numerical simulation demonstrated that the Bellman RRT improved the mission achievement using real deep space exploration scenario. (C) 2016, IFAC (International Federation Control) Hosting By Elsevier Ltd. All rights reserverd.

We consider the interplanetary trajectory design problem posed by the 8th edition of the Global Trajectory Optimization Competition and present the end-to-end strategy developed by the team ACT-ISAS (a collaboration between the European Space Agency's Advanced Concepts Team and JAXA's Institute of Space and Astronautical Science). The resulting interplanetary trajectory won 1st place in the competition, achieving a final mission value of J = 146.33 [Mkm]. Several new algorithms were developed in this context but have an interest that go beyond the particular problem considered, thus, they are discussed in some detail. These include the Moon-targeting technique, allowing one to target a Moon encounter from a low Earth orbit; the 1-k and 2-k fly-by targeting techniques, enabling one to design resonant fly-bys while ensuring a targeted future formation plane; the distributed low-thrust targeting technique, admitting one to control the spacecraft formation plane at 1,000,000 [km]; and the low-thrust optimization technique, permitting one to enforce the formation plane's orientations as path constraints.

This paper proposes a robust-optimal trajectory design method for uncertain system to minimize the expected value of cost-to-go function in Dynamic Programming. The fundamental idea is introducing Stochastic Differential Dynamic Programming (SDDP), which solves stochastic-optimal control problem by the second-order expansion of Bellman's equation around reference trajectory. Most recent studies have focused on trajectory optimization assuming that the spacecraft can control the trajectory perfectly as planed; however, the assumption is violated in realistic operations where uncertain events, such as navigation error or uncertainty on dynamical system, perturb the predetermined trajectory. Conventionally, experienced specialists empirically determine "margin" on optimal low-thrust trajectory by duty cycle or forced coast period. A proposed SDDP autonomously provides "margin" in optimization for future feedback as well. Numerical results by V-infinity leveraging problem show that SDDP has "margin" without duty cycle or coast period. Monte-Carlo simulation shows the SDDP solution has better performance than DDP considering uncertainty.

We propose thrust vector management by correctly positioning the thruster on a spacecraft by thrust vector measurement to decrease unwanted torque of thrust vector misalignment. A ground test was performed to measure 2-dimensional ion current distribution of 10W-class miniature ion thruster by electrostatic probe. The thrust vector measurement test showed that the thrust vector inclining angle was 1.4^º from the geometrically symmetric axis of the thruster. The thruster was positioned on the first interplanetary micro-spacecraft: PROCYON after redesigning thruster bracket. Thrust vector estimation in the initial on-orbit operation of 6.5 hours showed that thrust vector passes through within 5mm of the PROCYON's center of gravity.

This paper proposes a novel radiative cooling system for a high temperature superconducting (HTS) coil for an attitude orbit control system in nano- and micro-spacecraft missions. These days, nano-spacecraft (1-10 kg) and micro-spacecraft (10-100 kg) provide space access to a broader range of spacecraft developers and attract interest as space development applications. In planetary and high earth orbits, most previous standard-size spacecraft used thrusters for their attitude and orbit control, which are not available for nano- and micro-spacecraft missions because of the strict power consumption, space, and weight constraints. This paper considers orbit and attitude control methods that use a superconducting coil, which interacts with on-orbit space plasmas and creates a propulsion force. Because these spacecraft cannot use an active cooling system for the superconducting coil because of their mass and power consumption constraints, this paper proposes the utilization of a passive radiative cooling system, in which the superconducting coil is thermally connected to the 3 K cosmic background radiation of deep space, insulated from the heat generation using magnetic holders, and shielded from the sun. With this proposed cooling system, the HTS coil is cooled to 60 K in interplanetary orbits. Because the system does not use refrigerators for its cooling system, the spacecraft can achieve an HTS coil with low power consumption, small mass, and low cost. (C) 2014 COSPAR. Published by Elsevier Ltd. All rights reserved.

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.

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

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.