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Interplanetary micro-spacecraft have been recently demonstrated for various missions. Orbital control of micro-spacecraft is challenging because the spacecraft systems have severe constraints. This paper presents a constrained optimal flyby guidance algorithm. We iteratively introduce two-stage stochastic programming and achieve rapid and global optimization by using delta-v mappings and cross-correlation technique, which is well-known in the field of image and signal processing. Finally, numerical examples and Monte-Carlo simulation show that the proposed algorithm is efficient in terms of the expected total delta-v.

This paper proposes a novel control technique for a new panel quasi-static deployment and retraction method using electromagnetic force. This panel system is modeled as a multibody dynamic system which will be used for numerical simulations. Although the electromagnetic deployment has various beneficial points, it still has a major technical issue that vibration on panels continues for a certain time, and hence its motion does not converge quickly. To eliminate or damp these vibrations, we propose both feed-forward and feedback control methods using magnetism generated by electronic current. These methods can successfully reduce convergence time of deployment motion by at most 60 percents. Additionally, to realize feedback control, it is required to estimate the state variables of panels, that is, the angle between the panels. For that purpose, we introduce a new approach to determine these angles by sensing electromagnetic field generated by the deployment system with a magnetometer. This new method accomplishes the determination of the angles between the panels to an accuracy of at worst four degrees and one degree on average.

This paper highlights the power system design of the equilibrium lunar–Earth point 6-unit spacecraft (EQUULEUS), which will fly to the second Earth–Moon Lagrange point and perform a variety of scientific and technological missions. The limited power generation and high uncertainties of power generation and consumption are major difficulties encountered by the electrical power system of EQUULEUS for ensuring the feasibility of its mission. To address these issues, some methodologies are described in this paper in detail, which include the unique component arrangement, maximum power-point tracking function, and probabilistic analysis of power management. For the probabilistic analysis, Monte Carlo simulations are conducted based on a numerical model, and the feasibility of the power balance between generation and consumption is verified. [ABSTRACT FROM AUTHOR]

DELPHINUS is a camera system mounted on EQUULEUS, which is planned to be launched using NASA's Space Launch System EM-1 in 2021. DELPHINUS aims to investigate size distribution, influx ratio, and daily variation of meteoroids in the cislunar space through observations of lunar impact flashes (LIFs) from the far side of the moon. DELPHINUS will observe the moon's surface with the 60-fps camera modules to capture the flashes that are short duration phenomena. All image data cannot be downlinked due to constraints in memory size and communication capability. Therefore, an on-board image processing algorithm was developed to reduce downlinked data size by extracting only necessary pixel data including LIFs. Three experiments using three simulators were demonstrated to verify the real-time processing performance and detection capability. This paper reports the details of the proposed algorithm and the verification results. [ABSTRACT FROM AUTHOR]

he amount of debris in space is increasing every year, and its e↵ect on satellite missions is worsening. In this research, a method for removing debris using laser radiation pressure from a ground station is proposed. This method does not require a satellite to be launched to remove debris, and the system applied is less likely to be considered as a threat because it is relatively safe even if the laser erroneously irradiates another satellite, because the laser intensity is no more than 108 W/m2. However, the power applied to debris is small, such that it requires a great deal of time and energy to deorbit the debris. Therefore, this study aims to evaluate the feasibility of debris removal by laser radiation pressure from the viewpoint of energy consumption and time. In addition, in order to increase the feasibility of this method, a sub-optimized laser irradiation method for reducing the time and energy consumption required for removal is proposed. This study addresses how much energy and time is needed for a particular piece of debris to be removed by laser under three parameters: power of the laser, time that the laser is powered, and the location of the laser ground station. By sub-optimizing these parameters, the energy needed to remove a particular piece of space debris within a particular time is minimized. Furthermore, a laser irradiation method for deorbiting multiple pieces of debris using multiple laser ground stations is discussed. This study contributes to show that the new debris removal method using laser radiation pressure from laser ground station can e↵ectively and practically remove multiple debris in terms of time and energy cost. [ABSTRACT FROM AUTHOR]

The University of Tokyo has proposed a water resistojet thruster with a high certainty of liquid-vapor separation and low power consumption. In this propulsion system, liquid water is periodically vaporized in a pulsating manner to generate thrust. A vaporization chamber with a labyrinth-shaped flow path catches droplets using their surface tension to separate the liquid and vapor, and the droplets vaporize under normal temperature to reduce the input power by reusing the heat from the surrounding components. In this study, we designed and fabricated a flight model of the proposed propulsion system for 6U CubeSat and evaluated the performance of this propulsion system, including the control method. The results confirm the concept of the proposed liquid-vapor separation method and its low power consumption. Moreover, we revealed the relationships between the vaporizing duty cycle, input power, and thrust. [ABSTRACT FROM AUTHOR]

When a meteoroid impacts the Moon at several tens of km/s, a brilliant flash, referred to as a lunar impact flash, can be observed at the point of impact by ground-based telescope. Lunar impact flashes observed from the ground are biased due to atmospheric extinction, background illumination of earthshine, and short observation windows, typically a few hours per day during a period of approximately 10 days. NASA's meteoroid impact program took 10 years to detect 400 lunar impact flashes. EQUULEUS will demonstrate low-energy trajectory control techniques, such as multiple lunar flybys, proposed by the University of Tokyo and JAXA, within the Earth-Moon region. The spacecraft will be launched by NASA-SLS in 2019. The DELPHINUS camera system will be placed onboard EQUULEUS to observe lunar impact flashes while the spacecraft stays in halo orbit around an Earth-Moon L2 point. Thus, the lunar distance from the spacecraft is approximately one tenth that from ground-based observation. We estimate that DELPHINUS will detect 1,607, 2,699, and 4,534 lunar impact flashes during its six-month mission phase by assuming the limiting magnitude of its camera to be the 4.5th, 5.0th, and 5.5th magnitudes, respectively. The present study describes the DELPHINUS camera system and the first method for space-based observation of lunar impact flashes [ABSTRACT FROM AUTHOR]

This paper proposes an accurate aerodynamic model for membrane structures in free-molecular flow that allows the calculation of torque and force exerted on them, and which can be used for precise analysis of the attitudes and orbital dynamics of the structures. Recently, various types of deployable thin membrane structures attached to satellites for deorbit purposes have been developed and demonstrated on low-Earth orbits (LEOs), and their attitude dynamics have also been investigated. However, in early studies, deorbit sails were modeled as simple flat plates because there were no precise models for membranes that could express wrinkles, billowing, and asymmetric shapes. In this paper, a calculation method using tensor expressions of aerodynamic disturbance is proposed for the design of deorbit devices using high-fidelity membrane models. The proposed method is tested in the attitude stability analysis of membranes on LEOs, and the results show that the asymmetric shapes of the membranes affect their attitude stability. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

Initial flight operations of the miniature propulsion system I-COUPS (Ion Thruster and COld-gas Thruster Unified Propulsion System) are presented with problems found in space and its countermeasures for them. The I-COUPS was developed by the University of Tokyo and installed on a 70 kg space probe, PROCYON as main propulsion system to verify propulsive capability of the first micropropulsion in deep space. The PROCYON was successfully launched on December 3rd, 2014 and inserted into an orbit around the Sun. The PROYON project team started flight operation on the interplanetary orbit. Up to today, the cold-gas thrusters have successfully conducted unloading maneuvers since the launch. The ion thruster overcame several problems and achieved 223 hours operation with the averaged thrust of 346 μN. The I-COUPS will become the first electric propulsion and reaction control system operated on a small space probe (<100 kg) on an interplanetary orbit. [ABSTRACT FROM AUTHOR]

Solar Power Sail is a novel spacecraft with hybrid propulsion of large-area solar sail and electric propulsion driven by thin-film solar panel. The Solar Power Sail spacecraft is currently developed to probe into the outer-solar system bodies out to the Jupiter orbit, which was not well precedently explored. This mission aims to explore Trojan asteroids and interplanetary dusts by remote sensing and in-situ sampling using orbital rendezvous or fly-by technique. Scientific objective of this mission is not only planetary science but also novel astronomy; measurement of the infrared extragalactic background light without foreground contamination of the zodiacal light thanks to low-density environment at deep space, polarization measurement of the gamma-ray burst and accurate determination of its direction based on the interplanetary network technique. The Solar Power Sail mission will thus develop a new direction of space astronomy and planetary science providing us an interplanetary telescope site and will play an important roll to form a new interdisciplinary science field. [ABSTRACT FROM AUTHOR]

Enceladus is the only icy satellite known to exhibits on-going geological activity of water-rich plumes derived from the interior ocean. Here, we propose a sample return and in-situ measurement mission for Enceladus' plume materials. Depending on the cost, mission duration, and propulsion system, we propose three types of missions to Enceladus; type 1: free-return trajectory, type 2: trajectory orbiting Saturn, and type 3: trajectory orbiting Enceladus. Type 2 and 3 are preferable to type 1 in order to achieve lower encountering velocity to the plumes (> 4 km/s and 0.2 km/s for type 2 and 3, respectively) and, thus, to collect multiple and intact samples. High resolution mass spectroscopy of the gas components will provide essential information to understand the physical and chemical conditions of both the interior ocean and the solar nebula. Furthermore, detailed onboard and onshore analyses of returned samples could provide geochemical, preboplogical, and, potentially, biological context in the interior ocean of Enceladus. [ABSTRACT FROM AUTHOR]

This paper presents the extended missions of the IKAROS solar sail interplanetary spacecraft. IKAROS entered an extended operation phase at the beginning of 2011. During this phase, its spin rate was reduced to observe deformation of the sail membrane under low centrifugal forces, and on Oct. 18, 2011 the spin was reversed to enhance knowledge of the membrane’s stiffness against solar radiation pressure. We also investigated changes in IKAROS's attitude motions under reverse spin conditions. At the end of 2011, IKAROS switched to hibernation mode due to an increased Sun angle reducing on-board electricity generation. We later searched for IKAROS considering its attitude and orbital motion during hibernation, and on Sep. 6, 2012, we succeeded in locating the spacecraft, which came out of hibernation. A solar power sail can be part of a hybrid propulsion system, with electrical power generated by thin-film solar cells on a solar sail membrane being used to operate ultra-high specific impulse ion engines. This paper also introduces a follow-on mission to IKAROS, a round trip to the Trojan asteroid using hybrid electric photon propulsion. [ABSTRACT FROM AUTHOR]

The solar power sail is a deep space probe that will be powered by a hybrid propulsion system with solar photon acceleration and ion engines to explore the outer planetary regions of the Solar System without having to rely on nuclear power. The Japan Aerospace Exploration Agency (JAXA) launched the world's first solar sail demonstration spacecraft, “IKAROS” (Interplanetary Kite-craft Accelerated by Radiation Of the Sun), in 2010. This spacecraft successfully demonstrated several key technologies related to the use of a solar power sail in a deep space flight environment. JAXA is currently planning an outer Solar System exploration mission using the demonstrated solar power sail technology, where the spacecraft will fly to Jupiter and perform a swing-by for a Jovian Trojan asteroid. This study undertook a trajectory design and system feasibility analysis for this mission. Candidate target asteroids were selected based on ballistic trajectory analysis, and then, electric-propulsion, continuous-thrust trajectory design was conducted to verify the conditions assumed for the ballistic analysis (e.g., estimate of steering loss due to the low-thrust trajectory). It was found that out of over 4000 Trojan asteroids, only 7 are feasible candidates considering the preliminary system design results. To broaden the choice of target asteroids, it will be necessary to reduce the weight of the solar power sail itself, its deployment mechanism, and the bus electronics. [ABSTRACT FROM AUTHOR]

Solar sail is a spacecraft that has a large-scale membrane to utilize the solar radiation pressure for its thrust. Hence, maintaining the membrane structure during space flight is a critical issue to keep thrust performance of the spacecraft. In this paper, we focused on the electrostatic force due to spacecraft charging on the membrane as one of the possible factor to cause the deformation of the membrane structure. We had estimated the electrostatic force via charging simulation for the IKAROS spacecraft in solar wind plasma at 1.0 AU. We had also made a structural analysis for the deployed membrane of IKAROS with the electrostatic force. The structural analysis showed that the electrostatic force could hardly affect the membrane structure in this case. [ABSTRACT FROM AUTHOR]

Japan Aerospace Exploration Agency (JAXA) are planning an outer solar system exploration and sample return mission from a Jovian Trojan asteroid using a 3000 m2 solar power sail. The difficulty of this mission is a severe restriction on the weight; only 300 kg is allocated for sampling and returning to Earth. In this weight, fuel for trajectory and attitude control, sampling mechanism, re-entry capsule, and other systems required to return to the Earth are included. In this paper, a preliminary analysis of this sample return mission is conducted. Three scenarios for sampling are proposed; sampling with a 3000 m2 solar power sail, with a detachable small solar power sail using electric propulsion systems and with a small probe using chemical propulsion systems. The mission analysis shows that the most feasible configuration is to conduct the sampling with the 3000 m2 solar power sail using an extension mast. [ABSTRACT FROM AUTHOR]

An attitude model for a general spinning solar sail spacecraft under the influence of solar radiation pressure is presented. This model, called "Generalized Spinning Sail Model", can be applied to realistic sails with nonflat surfaces that have nonuniform optical properties. The unique behaviors predicted by the generalized spinning sail model are verified by actual operation of the Japanese spinning solar sail spacecraft IKAROS. It is shown how imperfections in the sail surface affect the attitude motion of spinning sails, and a compact mathematical model that can precisely reproduce the spin-averaged motion of the spinning sails is derived. The stability conditions and a reduced model that preserves the key characteristics of the generalized spinning sail model are also derived to reveal the unique properties of the attitude behavior of spinning sails. [ABSTRACT FROM AUTHOR]

This study proposes a micropropulsion system unifying ion thrusters and resistojet thrusters and assessing that propulsive capability. The remarkable features of the system are the usage of water propellant and unification of the two types of thrusters by the single propellant. Water has been regarded as an attractive propellant in the view points of safety, availability, handling ability, low molecular mass, and future procurement in space. A multimode propulsion system is an attractive solution for the increasing demand for nano-/microsatellite missions. The proposal is to use microwave discharge water ion thrusters, tolerant for oxidization by water, and low-temperature water resistojet thrusters, enabling reuse of the waste heat. As a result of the assessment, it was expected that the propulsion system would have 3U size (10×10×30 cm³) and 3.70 kg mass, which realize in total a 6U and 10 kg satellite with 3U and 6 kg satellite bus system. The ion thruster would provide the maximum ΔV of 630  m/s by 47 W system power and the resistojet thruster would have 3.80 mN thrust and 72 s specific impulse by 19.4 W. Additionally, reuse of the waste heat from ion-thruster power supplies would enable the simultaneous operations of the two thrusters even at 50 W, which is almost the same power as the single ion thruster operation. [ABSTRACT FROM AUTHOR]