Hall thrusters typically exhibit a discharge-current oscillation with a few tens of kilohertz, which negatively affects the power supply and induces electromagnetic interference (EMI). A pulsating boost chopper power supply (chopper PS), which applies a varying voltage to the Hall thrusters, was developed to shift the discharge-current frequency to an acceptable frequency because discharge-current oscillation is unpredictable and can emit EMI. In this study, a 1-kW anode-layer-type Hall thruster was driven with a chopper PS to evaluate the dependence of performance on the chopping frequency because past investigations tested the chopper PS with only magnetic-layer-type Hall thrusters. The experiments showed that the thruster using a conventional DC power supply induced discharge-current oscillation at 32 kHz (i.e., thruster-originated natural frequency). Meanwhile, the chopper-PS drive shifted the discharge-current frequency to the chopping frequency unless the chopping frequency was more than 52 kHz. Conversely, for a chopping frequency over 52 kHz, the dominant frequency of the discharge current returned to the thruster-originated natural frequency. Moreover, the thruster performance above the 40-kHz chopping frequency was comparable to that using a conventional DC power-supply drive. Furthermore, the chopper-PS drive successfully suppressed amplitudes of discharge-current oscillation and ion wall loss to the guard ring by 50%.

OMOTENASHI is a CubeSat that will be launched by a NASA SLS rocket. Its mission is to demonstrate that a CubeSat can make a semi-hard landing on the Moon. The 6U-size spacecraft consists of an orbiting module (OM), a rocket motor (RM) for decelerating, and a surface probe (SP) on the landing module. Its mission will prove successful when the accelerometer mounted on the SP sends its information back to Earth. Under the tight resource constraints, a unique mechanism was designed for separating the RM and SP from the OM, based on small volume and low mass. Its function has been verified through several tests. Moreover, the attitude disturbance at separation in orbit has been estimated by numerical simulation. The nutation angle has been evaluated through spin-drop testing and simulations, and using a correlation model, the nutation angle in orbit and the effect of several error factors have been estimated. Results reveal that the nutation angle is not determined only by the spin rate. The nutation before separation predominates at lower spin rates in determining the nutation angle of the RM after separation. This paper presents the design overview, test results, and attitude estimation.

Hopping mobility is effective for exploration robots present on small celestial bodies including asteroids and comets and their surfaces are covered in certain places with granular materials such as fine grain sand. However, existing studies do not address hopping mobility on the aforementioned types of granular materials in detail. Therefore, it is necessary to analyze hopping motion on sandy surface for future missions. This study presents a parametric analysis of the hopping motion on granular materials based on a discrete element method (DEM) simulation. In particular, the DEM allows for the numerical simulation of the dynamic behaviors of numerous fine particles. The DEM defines a sand particle as a simple sphere and computes mechanical interaction between each particle which follows some simple laws of dynamics. To analyze the hopping motion, we focus on two characteristics of the hopping motion: the initial hop velocity and hop angle of the robot. In DEM simulations, we change the friction and rolling friction coefficient of the sand particles, the acceleration of gravity and the angular velocity of the robot. We perform a simplified analysis and assume that the terrain surface in DEM simulation is smooth and flat and that the robot exhibits a cubic shape. Based on the research results, we qualitatively evaluate the exploration robot's motion on a sandy surface by DEM analysis when compared with its motion on a rigid surface.

Evaporation of a single n-decane droplet at high pressures and high temperature has been studied experimentally under normal and microgravity conditions. Ambient pressure was set at 0.10 and 0.50 MPa. Ambient temperature was fixed at 773 K. At normal gravity, initial droplet diameter was varied from 0.4 to 0.80 mm. To examine unsteadiness of droplet evaporation, instantaneous droplet evaporation rate coefficient was measured. The effect of natural convection on instantaneous droplet evaporation rate coefficient was evaluated as a function of the Grashof number. The unsteadiness of droplet evaporation appears in entire period of evaporation lifetime. The effect of natural convection on instantaneous droplet evaporation rate coefficient changes during evaporation and its dependence on Grashof number in the range where the effect of natural convection was regarded as quasi-steady was expressed as the empirical equations.

Thermophysical properties of the surface of asteroid 162173 Ryugu have been investigated through global, local and close-up thermal imaging from the Hayabusa2 spacecraft using the thermal infrared imager TIR, a two-dimensional thermographic camera. The set of thermal images during one rotation of asteroid allowed to map thermal inertia of the asteroid, indicating the characteristics of large boulders and rocks covering the surface. The close-up thermal images have resolved the temperature of each geologic unit and boulders one by one, indicating the variation of rocks. Advantages of using TIR are to investigate the surface physical state even from the home position, 20 km from the surface, and to track the artificial targets against the asteroid surface while they are not well seen with a visible imager at the same time.

Parachute drawing by its cover contributes to simplicity in mechanism of a sample return capsule. Attachment of a band part to suspension lines of the parachute cover is presented to improve attitude stability of the flat plate shaped cover. Aerodynamic characteristics of the cover with the band were evaluated through vertical and horizontal flow wind tunnel tests. The results show that the attachment of the band with an appropriate band perimeter and gap between the band and the cover surface can improve the stability remarkably. Escape route where air flowing inside the band is able to run away is necessary for the stabilization, which is similar as that stability of a parachute relates to air permeation through porosity of the parachute.

To realize innovative transportation systems for intercontinental flight and orbital spaceflight, it is essential to establish the necessary aerodynamic, structural, propulsion, and control technologies for a vehicle to fly at high altitudes and speeds within the atmosphere and to verify those technologies by using small unmanned supersonic experimental aircraft. This paper describes the switching conditions at waypoints from one flight mode to another for fully autonomous flight and the verification results of the flight path from takeoff to landing via a circuit by a 3 kg low-speed model airplane. It is confirmed that the switching conditions worked well and thereby the airplane flew on the designated flight path. Based on the results, a flight path is designed for a small unmanned supersonic airplane.

Design and manufacture of a hypersonic pre-cooled turbojet engine have been developed so far. Wind tunnel tests of the engine have been conducted using the Ramjet Engine Test Facility (RJTF) at JAXA Kakuda Space Center, which simulates Mach 4 flight conditions. The shock oscillation was found and mass capture ratio (MCR) was around 40.1%, which is lower than the design target. This paper shows CFD analysis to clarify details of the experiment results and to find the way to improve the intake performance. In this paper, a virtual nozzle is used to simulate experiment results. From CFD analysis, the intake buzz occurs. Shock patterns from CFD analysis agree with the experiment results quantitatively. When back pressure decreases, the intake buzz doesn’t occur and MCR increases to 79.1% and max TPR is 43.6%.

In future sample return mission, a large-scale sample return capsule will be required to take more amount of the materials. We performed a conceptual study of a parachute system for such mission. In this mission, a two-stage parachute system with drogue and main parachutes has been proposed to avoid the large opening shock with single stage parachute system. For evaluating the drag area and the loss of the drogue chute, low speed wind tunnel tests are performed. In addition, the drag area and the opening load factor of the main chute are evaluated with the helicopter drop tests.

This study modeled the aerodynamical effect of airflow on liquid droplets. The purpose is to simulate the deformation of droplets using the particle method without calculating the air flow field. This paper proposes the potential flow model, in which the airflow at the upstream side of a droplet is assumed to be potential flow around a circular cylinder and the corresponding pressure is applied to surface particles. It is compared with the Newton flow model, that was previously developed by the authors, by simulating droplet breakup in two-dimensions. Breakup patterns such as sheet stripping and catastrophic breakup were observed. In all cases (We = 15, 45, 150, 1000), the droplet shape was improved from when the Newton flow model was used. The model will be studied further to be applied to problems concerning the breakups and trajectories of droplets, such as water splash around aircrafts and fuel atomization.

On-board data processing is one of the prior on-orbit activities that improves the performance capability of in-orbit space systems such as deep-space exploration, earth and atmospheric observation satellites, and CubeSat constellations. However, on-board data processing encounters higher energy consumption compared to traditional on-board space systems. This is because the traditional space systems employ simple processing units such as single-core microprocessors as the systems do not require heavy data processing. Moreover, solving the radiation hardness problem is crucial in space, and adopting a new processing unit is challenging. In this paper, we consider a Graphics Processing Unit (GPU) accelerated in-orbit space system for on-board data processing. According to prior works, there exist radiation-tolerant GPU, and the computing capability of systems is improved by using heterogeneous computing method. We conduct experimental observations of energy consumption and computing potential using this heterogeneous computing method in our GPU accelerated in-orbit space systems. The results show that the proper use of GPU increases computing potential with 10-140 times and consumes between 8-130 times less energy. Furthermore, the entire task system consumes 10-65% of less energy compared to the traditional use of processing units.

In this work we demonstrate how SystemC-AMS can accompany widely used modelling tools and languages in spacecraft system design. The language bridges a wide range of abstraction levels from a high behavioural description down to the detailed design. To showcase its abilities models of a scalable spacecraft power system are presented. These incorporate distributed analog control loops, digital communication, power electronics, a complex environment and embedded software using the different modelling regimes provided by SystemC-AMS. We show how these models can be used throughout the whole spacecraft system design cycle. The use of abstract models for initial sizing of the power system and high fidelity models for the detailed design of subsystem components is shown. We illustrate how our PCDU Modular Breadboard approach bridges the gap between simulation model and hardware, by making use of the Platform based Design Methodology. In the end we provide an overview of how SystemC-AMS can be embedded into a setup using distributed model execution, configuration control and continuous integration techniques that are focused on software engineering.

This study deals with the behavior of a jumping/hopping exploration robot under microgravity. Through a dynamics simulator developed based on Open Dynamics Engine, we discuss the reachability of a hopping exploration robot to a target point on rough terrain of a small asteroid. In this paper, the leg arrangement of exploration robots is supposed on isotropic regular polyhedrons. First, by adding a constant input torque, the relations between the leg arrangement and rover's hopping speed/angle are investigated. Then, to make the hopping angle lower, two times of input torques are examined. After that, the reachability to a destination point is stochastically investigated for the different leg arrangements, considering unknown local surface slopes which are generated randomly.

To simulate the crater formation into the granular material covering the majority of the surface of a planet, a satellite, a comet nucleus or an asteroid by high-speed impact of a natural or artificial object, the Euler equations with the Compressible and Non-Expanding (CNE) fluid model were numerically solved by the finite volume method in the framework of the macroscopic continuum fluid dynamics. The irreversible nature of the granular material in the compression process was considered assuming a higher speed of sound at unloading than compression, which enables the fluid to have higher density after unloading than the initial uncompressed density. The pressure that represents the resistance against the compression was defined as a function of the density, and the hardening effect was considered assuming the pressure function whose slope increases with the increase in the density. In the two-dimensional crater formation problem, the result with the CNE fluid model seems more realistic than that with the conventional compressible fluid model, because the high-density zone compressed by the impact was allowed to remain at the bottom of the crater in the former case. The effects of the hardening and the sand box size on the crater formation were numerically investigated by the CNE fluid model. Finally, the usefulness of the present method for the development of impact probes was discussed.

Experimental and numerical studies were conducted to investigate the flow over a backward-facing step with finite aspect ratio. Measurements using pressure-sensitive paint, infrared thermal imaging, oil-flow visualization and schlieren imaging were conducted in the Hypersonic and High Enthalpy Wind Tunnel at the University of Tokyo. The results show the complex three-dimensional structure of the flow over a backward-facing step. The numerical simulation was also conducted, and the simulation results reveal that the fluid flow from the open side of the model has significant three-dimensional effect on the flow features. Comparison between CFD and experimental results is also made in this paper. A recirculation bubble found after the step, generated by the interaction of side expansion and recirculating flow, is observed in CFD results. A symmetric local peak pressure and heating region named ‘twin peak’ is also founded by both numerical simulation and experiments.

ALE-1 is a 68 kg microsatellite with the mission to generate artificial meteors. This satellite was developed by Tohoku University in collaboration with ALE Co., Ltd. It was launched into orbit by the 4th Epsilon rocket on January 18, 2019. In the artificial meteors mission, a pellet is released from the satellite as a meteor source. The source emits plasma when it reenters the atmosphere, so that it can be observed as an artificial meteor from the ground. Highly accurate orbit prediction at the time of the pellet release is required to achieve this mission. For orbit prediction, initial orbit state and orbit propagation are needed. Currently, highly precise orbit propagator exists. Therefore, the focus of this research is on a method to determine accurate initial state. In particular, online processing is noted because it is suitable for real-time mission. To obtain initial state in online processing, Kalman-Filter-based orbit determination (OD) methods are often used. However, they have the problems of vulnerability to outliers and insufficient accuracy. In this study, to solve these problems, two methodologies were proposed. The first methodology is to realize filter robustness against outliers by combining data from multiple Global Positioning System Receivers (GPSR). The second is to selectively adopt plausible orbit state by sequentially estimating the increase rate of propagation. As a result of applying the proposed methodology, highly accurate and reliable OD was realized and the orbit prediction accuracy satisfied the mission requirements.

A small satellite program is under study to demonstrate the technology for the Solar Power Satellite (SPS). This small satellite is supposed to conduct a demonstration test of the onboard transmitting antenna in a low earth orbit, and it is considered to measure the antenna pattern on the ground to evaluate its effectiveness. In the ground measurement of the satellite antenna pattern, without using the movement and attitude control of the satellite itself, it is effective to set up the receiving stations at multiple points on the ground and compare the reception levels with each other. However, the number and arrangement of these receiving stations and the method of processing data including errors have not been studied in detail. The relationship between the measurement accuracy of each receiving stations and the boresight position, and the dependency of the measurement error by the receiving station placement method are shown. As a result, the boresight position can be estimated at 1/20 or less of the beam width when the measurement of the receiving station is performed within 3 dB. Throughout these results, we obtained a guideline for accurate pattern measurement.

A thrust prediction system using a neural network for controlling Hall thrusters automatically is under development. This network has described the time variation of discharge current within 1% error, but calculation of the current was very cumbersome (2100 s) and overfitting occurred. In order to reduce the calculation cost and to prevent the neural network from overfitting, we have adopted a stacked auto encoder and optimized the network model using a genetic algorithm. The calculation time was reduced from 2100 s to 100 s without overfitting. The present system cannot yet describe hysteresis of discharge current; this will be addressed in future work.

In this study, the transferring problem to a halo orbit by a low propulsion is considered. This problem is reduced to a nonlinear optimal control problem to minimize a cost function which represents energy consumption. As a method of solving a nonlinear optimal control problem, we use continuous-time differential dynamic programming (DDP) with input saturation constraints, which we have proposed. The proposed DDP is an iterative solution based on the local quadratic approximation, and the correction term of the control input at each iteration is obtained by solving the Riccati differential equation, whose solution can be easily numerically computed. Then, solving an optimal control problem by the proposed DDP, we design an optimal transfer trajectory from an Earth parking orbit to a halo orbit in the Earth-Moon system with a stable manifold and a low propulsion.

ALE-1 is a technology demonstration microsatellite with the mission objective of generating artificial shooting stars. It was launched in January 2019 to an orbital altitude of 500 km. There are manned missions at a lower altitude in the International Space Station (ISS), therefore the altitude of ALE-1 needs to be lowered below the ISS before the mission starts. The de-orbit technology being used is a deployable and separable drag sail. In this paper, a novel non-intrusive method to collect and process dynamic behavior data of deployable space structures using a Time-of-Flight monitoring camera system is presented. This camera system collects distance information on the sail, and the collected data is downlinked to the ground for analysis. With this module, better insight into the dynamic behavior of the flexible components of the de-orbiting system, namely the boom and the drag sail, can be obtained. This paper also presents the data processing method for the shape estimation and surface reconstruction of the drag sail. The acquired analysis results can be utilized to better understand the dynamic behavior of deployable space structures and drag sails, in order to achieve more accurate de-orbiting simulations and predictions.

An international scientific micro-satellite named RISESAT (Rapid International Scientific Experiment Satellite) was launched on January 18th, 2019 on the Epsilon rocket. The main objective of RISESAT is Earth observation and technology demonstration. RISESAT carries six international scientific payloads from Taiwan, Czech Republic, and Japan. For controlling all payload instruments, an on-board payload handling computer named Science Handling Unit (SHU) was developed based on Space Plug-and-Play Avionics (SPA). In order to connect with the SHU, each payload is equipped with a Micro Remote Terminal Unit (μRTU), which converts the payload to a SPA-compliant component. Therefore, the hardware interface from the payloads to the SHU is standardized and payloads can be connected to an arbitrary connector of the SHU. This way of integration can also be applied to the rest of the on-board avionics in the future and has the potential to make the system integration of satellite easier. The soundness of the functional performance of the developed plug-and-play system, as well as the concept of system integration have been verified through ground electrical tests and orbit operation of RISESAT.

The exploration of minor bodies, such as asteroids and comets using robotics is a necessary step for the study of the Solar System's evolutionary process. For such purpose, a multi-legged ground-gripping robot was proposed to perform precise locomotion towards specific places of interest. Moreover, stable locomotion is expected, since this robot has grippers capable of grasping the rocky and uneven terrain of the minor bodies, maintaining the attachment to the ground and preventing the flotation of the robot on the microgravity environment. Decreasing the forces induced on the grippers is essential to keep the stable locomotion in such environments and, in this paper, a gait control method is proposed to reduce accelerations and motion reactions on the robot, avoiding high reaction forces on the contact points with the ground. This method focuses on generating trajectories to be traveled by parts of the robot. A numerical simulation was developed to investigate the effectiveness of the proposed method in decreasing reactions on the robot during the motion of the robot on uneven terrains by comparing different trajectories' parameters.

This paper describes the design and tests of 0.4-N class gas-generator-type hybrid thruster using laser ignition for onboard propulsion systems. Conventional hybrid rocket engines have advantages such as variable thrust and combustion interruption. However, the oxidizer-to-fuel ratio (O/F) is varied from the optimal value during thruster firing because the fuel-grain cavity is enlarged owing to combustion, and this reduces specific impulse. Hence, we propose a gas-generator-type hybrid thruster using laser ignition for onboard propulsion systems. In the thruster, the gas generator produced fuel-rich combustion gas using the oxidizer and solid fuel at a constant mass flow rate to maintain O/F at the optimum value. The semiconductor laser is used to ignite the solid fuel for enabling the multi-start capability. The firing tests showed that the use of hydroxyl terminated polybutadiene/ammonium perchlorate/carbon black=50/50/0.5 wt% allowed the gas generator to produce the fuel-rich combustion gas at the target mass flow rate, and that combustion was interruptible by discontinuing the oxidizer supply.

This paper describes the development status of a laser ignition system for a solid rocket motor. This system is being developed as a simple, lightweight, and small design with a high resistance to electrical disturbances and a high level of safety. The most notable advantage of this system is that its high level of safety can decrease the cost of launching rockets into space. A laser initiator and a laser safe-and-arm device (laser S/A), which are essential components of the proposed system, were developed. In particular, prototypes of the laser initiator and laser S/A for the ignition of an upper stage rocket motor were manufactured, and some environmental tests, which are required for space rocket devices, were conducted. In addition, the lowest laser energy that is needed to ignite the laser initiator was determined by changing the laser power and operating time of the laser S/A. Furthermore, a small rocket motor vacuum fire test was successfully conducted.