The design of mission specific aerospace power controllers is constrained by cost and development time, often leading to poorly adapted and less then optimal heritage solutions. We want to provide modularity on circuit level to be able to deliver truly bespoke units. This is only possible if manual labor in the development process is reduced by design automation and tailored processes. Our approach is applied to the whole development effort from architectural exploration to detailed design. In this paper we focus on the prototyping stage that aims at delivering a correct-by-design functional model of the product under time and budgetary constraints. We demonstrate the application of our Avionics Design AutoMAtioN Toolkit (ADAMANT) for aerospace power controller design. A functional model of the controller is then implemented using our Modular Breadboard prototyping system. The prototype is demonstrated on the DLR Core Avionics Testbed (CAT), a platform to perform subsystem component tests in a mission like electrical environment. The state of the overall process and toolset development, lessons learned from this first application and future activities are presented.

Considerable research is being directed toward active debris removal (ADR) in space, especially by small, cost-effective satellites. In the descent phase, thrust vector control (TVC) is often utilized to maintain the satellite’s trajectory and thrust vector by gimbaling the thrust vector. A common challenge in performing TVC is that its effectiveness depends on the mass, moment of inertia and location of the thrusting point with respect to the center of mass. Since these properties change with fuel consumed during a mission, a designer needs to consider such parameters as they strongly affect the resulting dynamics. This paper introduces spacecraft dynamics with a gimbaled thruster and proposes an adaptive controller for TVC with model uncertainties. Through numerical examples, the effectiveness of the proposed controller is demonstrated for improved control performance.

The issues of degradation of data of on-board sensors, deterioration of thermal control function, and so on are known to occur by depositing molecular contaminants on the surface of spacecraft. To solve these issues, the application of adsorbent is proposed in the present research. Pure porous material (MCM-41) and 4 types of adsorbents consisting of MCM-41 as the base material with different binders (methyl cellulose 4000 or water glass) were prepared. Some samples were heat-treated at 150°C for 3 hours. To clarify their adsorption characteristics and apply them to space use, they were evaluated by scanning electron microscopy (SEM), gas adsorption method, outgassing test, adsorption test toward hexamethyl cyclotrisiloxane (D3) method, and thermogravimetry analysis combined with in situ evolved gas analysis by mass spectrometry (TG-GC-MS). These analyzes revealed that all of the adsorbents have a high adsorption capability to D3. In particular, the two types of the adsorbents with MC4000 binder showed an improved adsorption capacity of D3. The D3 desorption process started at a temperature of around 200 °C for the 4 types of manufactured adsorbents, while the pure MCM-41 powder desorbed D3 completely below 200 °C. The forming process of the adsorbents, such as mixing binder components and heat treatment conditions, would influence the desorption process of D3.

The measurement of exhaust plasma is useful for predicting thrust, and 2D distribution can estimate the thrust vector or plasma physics near the thruster exit region. However, measuring the spatial distribution in a narrow, millimeter-scale area of a miniature plasma thruster for miniature satellites is difficult. This study aims to obtain the qualitative spatial distribution of the electron and neutral-particle number densities of the miniature RF plasma thruster with a diameter of 3.4 mm using a simple optical system around the discharge-tube exit. An optical system capable of simultaneous magnified imaging at two wavelengths is developed, and the intensity of the 2D distribution of ArI of 420 nm and ArII of 488 nm emission lines is measured near the exit of the miniature RF plasma thruster. The local profile in the region of 10 mm × 10 mm near the thruster exit is obtained through Abel inversion, from the 2D distribution of the integrated values in the line-of-sight direction. The optically measured results qualitatively agree with the probe measurement at the downstream end of the measurement area.

With the increasing usage of microsatellites in recent years, the demand for propulsion systems has also increased. Therefore, in this study, we developed a Microsatellite-Friendly Multi-Purpose PROPulsion system (MFMP-PROP) using 60 wt.% H₂O₂ as a basic propellant. MFMP-PROP can operate in the monopropellant mode for small impulse bit and the bipropellant mode for a relatively large thrust with a comparably high specific impulse. In this paper, the bipropellant mode is introduced and described, including reports on the progress concerning their research and development. As an approach to stable combustion, the effects of mixing ratio, preheating temperature, and mass flow rate on the operating characteristics were evaluated. As a result, we demonstrated a reliable and stable combustion range. The lifetime and continuous operating time were evaluated to investigate durability. As a result, 132 ignitions and 7,900 s of total operating time were achieved, during which the propellant consumption reached 1,375 g.

Analogy between the hypersonic turbulent transition and directed percolation (DP), a simple stochastic model, was verified to provide a new insight for predicting the transition which increases the aerodynamic heating. A direct numerical simulation was performed on a Mach number 6 over a flat plate to obtain the values required for the v erification. A blowing and suction disturbance was introduced to the laminar solution to trigger the transition to turbulence. By analyzing the flow behavior in the turbulent transition region, one of the critical exponents of DP was estimated and was compared with the theoretical value. The turbulent fraction at each Reynolds number based on the distance from the leading edge of the flat plate is required to obtain the critical e xponent. The critical exponent depends on the threshold value and the distance from the wall, and some of the critical exponents did not agree with the theoretical value or did not reproduce the turbulent fraction distribution. Conversely, some of the critical exponents obtained from the present numerical simulation results were close to the theoretical value. This suggests that hypersonic turbulent transitions could also belong to DP universality class.

Towards industrial use of dielectric-barrier-discharge (DBD) plasma actuators, a practical DBD plasma actuator system including a driving circuit assumed to equip unmanned aerial vehicles (UAVs) is proposed in this study. It is preferable to use a high-power battery already equipped with a UAV as a power supply of the DBD plasma actuator. We develop the driving circuit, which consists of a DC-power supply (around 500 V), simple booster circuits, and a Cockcroft-Walton voltage multiplier. The DBD plasma actuator is configured with assuming to divide an ionization role and a charged-particle acceleration role. At the ionization part, approximately 2 kV pulses are repetitively generated with a frequency of 100 kHz, successfully forming stable surface DBD. At the acceleration part, a high-DC voltage of approximately 10 kV is applied. The electrical characteristics of the developed DBD plasma actuator shows a similar trend to that of a nanosecond-pulse-driven DBD plasma actuator. Also, the peak velocity of the induced ionic wind exceeds 2 m/s, indicating that the performance of the developed DBD plasma actuator is comparable to the conventional DBD plasma actuators. The developed driving circuit considerably saves the additional weight and cost compared to the conventional high-power and high-voltage power supply, advancing a step toward the industrial use of the DBD plasma actuators for drastic improvement of the aerodynamic characteristics of unmanned aerial vehicles.

Laser velocimetry techniques such as Particle Image Velocimetry (PIV) and Laser Doppler Velocimeter (LDV) capture the scattered light from the laser beam on tracer particles added to the fluid. However, these laser velocimetry techniques can be difficult to apply to high-speed fluids, where adding tracer particles and tracking particles are difficult. This study attempts to analyze images of high-speed impinging jet flows by applying the optical-flow image analysis to typical shadowgraph time-series images to calculate the velocity field without tracer particles. In the image analysis of such high-speed fluid phenomena, the spatiotemporal resolution of a high-speed video camera is a major issue. This study investigates the range of flow velocities for which the optical flow image analysis can be applied to high-speed impinging jets in the range of NPR = 1.5 to 2.7, using the latest Phantom high-speed video camera with high spatiotemporal resolution. The optical flow image analysis allows all NPRs to clearly capture the shear layer oscillations of the jet and the advection of vortices caused by the shear layer instability and the secondary vortices near the wall, respectively. Furthermore, the peak frequency of the optical flow associated with the vortex shedding coincides well with that of the wall pressure. These results show the possibility of extracting velocity fields in high-speed flows by the optical flow image analysis without tracer particles based on shadowgraph images.

The combustion characteristics of hybrid rocket motors have been widely studied and many equations for fuel regression rates have been proposed. Among these, the equation devised by Marxman et al. has been used most often. However, while the results obtained from this equation are in good agreement with the experimental data for some fuels, it is inaccurate for many other fuels. This method was derived using a constant wall temperature model that tends to give results inferior to a variable wall temperature model. Thus, the present work employed a variable wall temperature to derive a theoretical mass flux exponent. The resulting exponent was compared with that predicted by the Marxman equation and with an experimental value. As a result, it is smaller than the value reported by Marxman et al. and slightly closer to the experimental value. Therefore, it was found that the fuel regression rate has a larger error than the error that the surface temperature is constant.

A small-scale supersonic flight experiment vehicle (OHWASHI) is being developed at Muroran Institute of Technology as a flying testbed for verification of innovative technologies for high-speed atmospheric flights which are essential to nextgeneration aerospace transportation systems. Dynamic derivatives due to angular rates of attitude motion are essential for the analysis of the six-degree-of-freedom flight path and the design of autonomous guidance and control systems for the experiment vehicle. They have been evaluated by subsonic wind-tunnel tests and CFD analysis. In this study, mechanism of dynamic derivatives due to roll and yaw rates is investigated by comparing smoke-wire visualization and CFD analysis. As a result, followings are confirmed: For 𝐶_lr in the shorter-nose configuration, i.e. Nose-A, the lateral difference in the vortex flow on the main wing generates a rolling moment in the direction opposite to the yaw motion. For C_nr in the Nose-A configuration, the vertical tail has a narrower streamline spacing on its left side, which will degrade the yaw damping. In the longer-nose configuration, i.e. Nose-C, the streamline spacing is narrower on the right side of the vertical tail, which would cause yaw damping enhancement. For 𝐶_lp, roll damping is considered to take place due to pressure difference between the right and left main wings.

The concept of a small precipitation radar constellation has been proposed by the Japan Aerospace Exploration Agencyas one of the future precipitation observation missions. If realized, this can improve the quality of the Global Satellite Mapping of Precipitation, or GSMaP. This study evaluated the impact on the accuracy of GSMaP resulting from increased spaceborne radar　observations over Japan. The Japan Meteorological Agency provides high-accurate and high-resolution ground precipitation radar data calibrated by rain gauges. These ground observation data were used to generate the pseudo future precipitation radar data and inserted into the operational GSMaP processing in some temporal intervals. Both the sampling errors and retrieval errors were taken into consideration by utilizing Dual-frequency Precipitation Radar aboard the Global Precipitation Measurement Core Observatory. The relationship between temporal interval of insertion to the GSMaP processing, and improvement of GSMaP accuracy were verified. Results indicated that accuracy improved as pseudo Precipitation Radar (PR) data were inserted into the GSMaP processing more frequently. Furthermore, the GSMaP accuracy improved not only with the time when the pseudo PR data was inserted but also several hours after the insertion, which can be due to the characteristics of the GSMaP algorithm.

Small rockets using hybrid rocket engines have been developed as a part of space development projects by student groups at various universities. The actively developed technologies include supersonic rockets and self-built engines. However, development based on aerodynamic technology has few precedents. In this paper, a case is described in which the Nagoya University Aerospace Flight Technology NAFT demonstrated attitude control technology, which is indispensable for rockets, by using aerodynamic technology. The main mission of the Masterpiece-05 Leidenschaft launch vehicle used for this launch demonstration was to achieve autonomous attitude control by rolling with movable fins. The goal was to roll the rocket by 45 deg after the engine burnout was stopped. The control law for the roll attitude control was proportionalderivative control, and the value of the gain was determined by estimating the torque in the roll direction generated by the movable fins using computational fluid dynamics. The launch took place during the Izu-Oshima Joint Launch Experiment (Izu-Oshima Space Event) held on November 9–11, 2018. The altitude and attitude of the airframe were logged. After stopping the roll of the airframe, attitude control was performed from 3.90 to 8.45 s after ignition. As a result, the phase difference between the roll attitude angle at the start of attitude control and the converged roll attitude angle at the end of attitude control was approximately 51.8 deg. The mission was confirmed to have been accomplished based on the numerical values of the roll attitude angle obtained from the data logger. These efforts are significant from the viewpoint of educational value in space engineering because there are few examples of extracurricular activities wherein students take such initiatives.

We developed an integrated attitude and orbit control system (AOCS) module for advanced CubeSat missions. The module is 1U and 1kg and is installed with all sensors and actuators required for accurate three-axis control. The target attitude control accuracy is 10 arcsec, which is the same as for other existing integrated AOCS modules, such as XACT and iADCS. In addition, we strongly focused on the transparency and expandability of the software from the user's viewpoint, and we decided to use open-source software for flight software and ground testing tools. This study shows an overview of the hardware and software design of the module. The development of the engineering and flight models was completed, and the hardware tests were successfully executed. The flight algorithms were developed based on the open-source software C2A and tested with the open-source simulator S2E and ground station software WINGS. These design and development results show the effectiveness of the open-source strategy in improving the transparency and expandability of the AOCS module.

In this study, we investigated numerically the supersonic aerodynamic characteristics of re-entry capsules designed with hyperbolic contours and compared them with Hayabusa re-entry capsule. The numerical results showed that the drag coefficients increased using the hyperbolic contours compared to Hayabusa re-entry capsule without drastically changing the capsule’s shape. This result suggests that shape design based on the hyperbolic contours can improve the aerodynamic characteristics of re-entry capsules.

SHIKISAI, the Global Change Observation Mission-Climate (GCOM-C) satellite, was launched on 23 Dec. 2017 and began global observation with a Second-generation Global Imager (SGLI) on 1 Jan. 2018. SGLI has seventeen channels from near-ultraviolet (NUV, 0.38 m) to thermal infrared (12 m) wavelengths and two polarization channels at the red and nearinfrared wavelengths with a spatial resolution of 250 m or 1 km and a swath width of 1150 km or 1400 km. The goals of GCOM-C are to continuously monitor changes in the global environment and acquire the knowledge needed to improve our prediction of the earth-system environment. For example, the unique SGLI channels and spatial resolution enabled us to observe the following atmosphere and ocean surface phenomena. Thick aerosols from the wildfires in eastern Siberia were characterized by SGLI polarimetry and NUV channels, which are advantageous in detecting small-particle aerosols and optically absorbing aerosols. The 250-m resolution multi-channel observation enabled us to evaluate the spectral characteristics of water-leaving reflectance at patches of high chlorophyll-a concentration in coastal areas, including various kinds of red tide. Pumice rafts produced by a submarine volcano, Fukuoka-no-ba, were detected by the SGLI 250-m resolution images in 2021. We derived the fine resolution ocean surface current using pairing sensors of similar spatial resolution with tens of minutes time lag. GCOM-C/SGLI can be expected to play a role in the earth system's observation networks by continuing the observations over more periods than the planned 5-year lifetime.

This paper describes the space related human resources development project conducted by KOSENs in 2020. Satellite development is effective for practical education in KOSENs. As such, KOSEN Space Collaboration Group have held bootcamp style education programs for KOSEN students since 2015 and online lectures on space science and technology since 2018. Herein, we have proposed a new contest called the All KOSEN Space Contest to cultivate satellite mission planning skills. To evaluate the efficiency of this proposed contest, we conducted a pre-contest in 2020 and researched the educational benefits of this pre-contest. From the results, we were able to extract useful insights to guide the planning of the main contests with potential to bridge the gap between our previous educational programs and actual satellite development. Consequently, the result of this study indicates that the satellite development skills of KOSEN students can be further improved.

The plan to construct a Lunar Orbital Platform-Gateway (LOP-G) is underway for further development in the cis-lunar space. The highly elliptical orbit around the Moon named Near Rectilinear Halo Orbit (NRHO) is a candidate for the LOPG mission. This paper is on the transfer trajectories from Low-Earth Orbit to NRHO in the minimum time for a given fuel. Two cases are assumed for the transfer trajectories: Two-Impulse Transfer and Powered Lunar-Flyby Transfer. The trajectories are obtained from the optimization calculation with the Multiple Shooting Method in the Circular Restricted Three-Body Problem (CRTBP). This paper shows the trend of the flight time and Δ𝑉 for the transfer trajectories in two cases as the results. Furthermore, the stability of the mission duration is investigated with the position of LOP-G.

Magnetoplasma sail (MPS) is a space propulsion system using the interactions between the solar wind and the magnetopause generated by the onboard coil. MPS performs magnetospheric inflation by plasma injection. There are two mechanisms of magnetospheric inflation: the effect of extending the magnetic field line due to dynamic pressure and the effect of expanding the magnetic moment due to thermal pressure. In this study, the influence of the pressure types on the angular dependence of thrust of MPS considering both the thermal pressure and the dynamic pressure of the injection plasma at different circumferential injection angles was investigated using MHD calculation. The results show that the maximum thrust gain is obtained at 0° in the case of the injection plasma with no initial velocity. However, in the case of the injection plasma with the dynamic pressure, the thrust gain was taken as the thrust peak at 10°. For a higher injection angle, the acceleration in the magnetopause after injections is sufficient to cancel the magnetopause current, and the thrust gain becomes small.

For making better use of the Sun-Earth libration points, small-amplitude periodic orbits controlled by Solar Radiation Pressure (SRP) were previously proposed, whose amplitude is much smaller than the classical halo orbits, and the geometry with respect to the Earth is nearly fixed. However, the mechanics were investigated only in the Circular Restricted Three-Body Problem (CR3BP), and only typical computation examples were shown. To further improve the study, we extend the study in two aspects in this paper. First, we establish an orbit design scheme for arbitrary-amplitude periodic orbits in the nonlinear CR3BP using an optimization problem and find a family of small-amplitude periodic orbits controlled by SRP. Second, we establish an orbit design method in the high-fidelity ephemeris model, where the influence of the lunar gravity is significant. A reference periodic orbit designed in the CR3BP is translated into a quasi-periodic orbit. We also propose a strategy to avoid eclipse in the design of quasi-periodic orbits in the ephemeris model.

This research investigates a low impact stable self-deployment of a shape memory polymer convex shell. The convex shell generates a self-deployment force. After deployment, it has a higher bending stiffness with a reduced thickness. The self-deployment force depends on the initial geometry, such as the degree of flatness or curvature of the convex shell when it is stowed. To examine the effect of the initial condition, we investigate the three different forces required to prepare the flat shape before stowing. Theoretical analysis is performed to derive the work required by the weight to flatten the shell. Then, the deployment behaviors for each condition are compared by determining the deployment ratio. As a result, the deployment ratio and rate depended on the force used to make the stowed condition. In addition, repeatability of the deployment is discussed with multiple deployment tests. Furthermore, theoretical analysis is presented to discuss the required force and work done by weights used to flatten the shell. Finally, some nonlinear finite element analyses of a single convex shell are also presented to show the predicted deformation of the convex shell.

A corner reflector (CR) is a passive device that reflects electromagnetic energy to return to the source of energy. It has been implemented as a known location and displacement for the calibration and validation of SAR data. Moreover, it can be employed for accurate surface displacement monitoring. In this study, we experimented with 2 different sizes of CRs at the playground of the university including 50 cm small CR and 141.4 cm large CR (side length dimension). The CR reflectors have been set up with appropriate orientation and tilt angles regarding targeted SAR sensors. Afterward, radar cross-section (RCS) and signal-to-clutter ratio (SCR) of CR were performed and compared with the theoretical expectations. SCR value can be used to estimate displacement error. This study found that large CR offered significant visibility and stability than small CR. Large CR provided higher SCR more than 28 dB (equivalent to 0.12 mm of phase error in a single SAR data), and small CR provided SCR at 14.58 dB (equivalent to 0.58 mm of phase error). The evaluation of this study is beneficial for gaining insights into the potential and limitations of CR performance with specific SAR data. This can aid deciding an optimal CR and calibration site for accurate displacement monitoring in the future.

Currently, a demonstration experiment using a small satellite is planned to realize future solar power satellites (SPS). There are two purposes for the experiment. One is to demonstrate satellite-mounted equipment in space, such as the highprecision beam-forming antennas used in SPS. The other is to conduct small-scale space-ground power transfer demonstrations, if possible. As the transmitting (Tx) antenna for the SPS, an active phased array antenna is used because the satellite is capable of forming a Tx beam in an arbitrary direction. For the onboard test, it is necessary to accurately evaluate the beam shape transmitted from the satellite on the ground. For this purpose, as an experiment on ground reception, we propose a method for installing receiving stations in multiple straight lines and arranging them in a direction orthogonal to the moving direction of the satellite. We calculated the theoretical pattern from the satellite's phased array antenna, which was 1.9 mφ, and tried to reproduce the Tx antenna pattern from the measured value of the reception level of each receiving station on the ground. The pattern was approximated by a quadratic function and sinc function, and reproduction up to the main beam and the first side lobe was confirmed, respectively. We calculated the boresight position error with respect to the number and spacing of the receiving stations and showed that it can be estimated with an error of 1 m or less by approximating the quadratic function. In addition, two different orbit inclination angles were calculated as candidate satellite orbits, two receiving station sites were selected, and their superiority and inferiority were compared. As a result, appropriate guidelines were obtained regarding the arrangement and spacing of the receiving stations in the ground evaluation of the SPS demonstration test satellite. Furthermore, the requirements for candidate sites for the receiving station and the satellite orbit are summarized. This makes it possible to reproduce the shape of the transmission antenna beam and the antenna pointing direction.

A shock stationary frame method is proposed to simulate efficiently an attenuating shock wave due to the interaction with a boundary layer in an expansion tube. Numerical simulations of the expansion tube were conducted and compared with two methods: the proposed shock stationary frame method and an adaptive mesh refinement technique in laboratory frame, which was used in a previous study. As a result of comparing the boundary layer development and the shock speeds in the numerical simulations by the two methods, they showed reasonable agreement and reproduced the analytical solution. The shock stationary frame method can reduce the computational cost, which is expected to be applied to calculations for the turbulent boundary layer development in long-distance propagation of the shock wave.

In recent times deep neural networks have consistently given state-of-the-art results in the remote sensing domain. However, deep neural networks need a larger dataset for training, this is one of the bigger challenges in case of the Synthetic Aperture Radar (SAR) satellite data application. Building a larger training dataset for different SAR satellites have the following constraints- 1) The interpretation of SAR images needs expert knowledge, so annotation is a time-consuming and resource-intensive procedure. 2) Most of the SAR satellites are commercial, so gathering enough scenes just for training data preparation for each satellite is not practical. 3) Many new satellites are getting launched and these do not have enough flood events in their archive data for training. To mitigate these constraints, authors have proposed to utilise one large dataset for the flooded area which is extracted from free SAR satellite data and then transfer learning has been used to apply on other satellite images. Transfer learning in deep neural networks is the concept of training a network on a source dataset that is much bigger and then transfer the learned knowledge or network weights on a target dataset which is, in general, smaller in comparison to the source. In this paper, the authors will present the methodology to apply transfer learning to the SAR images. The source training dataset is made from Sentinel-1 images and the target dataset is Advanced Land Observing Satellite-2 (ALOS-2) images. The same methodology has been implemented with the ALOS-2 images to different test areas in different countries namely Japan and Thailand. The result shows the notable improvement over the networks trained from scratch, also the problem of overfitting has been avoided successfully.

JAXA, CNES and DLR have decided to cooperate to develop and fly a scaled reusable VTVL rocket stage called CALLISTO (Cooperative Action Leading to Launcher Innovation in Stage Toss - back Operations). This vehicle is paving the way for future reusable launch vehicles in Europe and in Japan. During phase B important progress in term of methods and operation philosophy specific to RLV have been made. Amongst other progresses, that will ease the development of future operational VTVL, in the domain of aerodynamic modelling, GNC landing leg deployment but also flight domain definitions are presented. These are concrete results which can at least partly be useful for other RLV projects.

Small/micro satellite developments continues to increase, and the demands of propulsion system specialized in small/microsatellites are also getting bigger. Our research group has proposed Microsatellite-Friendly Multi-Purpose Propulsion System (MFMP-PROP) with dual mode thruster of monopropellant and bipropellant. In this paper, to aim extension of catalyst life of the dual mode thruster, research and development status of new catalyst are discussed. Our current catalyst is platinum (Pt)/aluminum oxide (Al₂O₃) on metal honeycomb. The objective of this study is to evaluate titanium oxide (TiO₂) capability instead of conventional wash coat such as Al₂O₃. Base material capability tests with TiO₂ sintering and Pt supporting processes are evaluated. Four different types of base materials, three stainless steels and one aluminum alloy, were tested. It is considered that SUS316L shows better surface conditions than others even though discoloration of thin TiO₂ layer is confirmed. From field emission scanning electron microscopy (FE-SEM) and energy dispersive X-ray spectroscopy (EDS) scanning results, TiO₂ layer helps Pt distribution evenly. Without TiO₂ layer, Pt agglutination is confirmed. Propellant reactivity tests are also successfully conducted with Pt/TiO₂ catalyst. From the results of these studies, it is considered that Pt/TiO₂ catalyst are a possible replacement for our conventional catalyst.

We present an automated system for environmental testing of on-board devices used as a common bus equipment in Japan Aerospace Exploration Agency (JAXA)’s balloon-borne experiments. For each experiment, we provide users with such devices. A telemetry tracking and command system in the balloon consists of a transmitter, receiver, and main/sub balloon operator (MBO/SBO) that controls the other devices according to a set of commands and creates telemetry frames from housekeeping data. We perform environmental tests of the MBO/SBO to confirm their functionality before the assembly of the bus equipment. So far, it took 2 h/30 min to finish one cycle of the MBO or SBO, respectively, during which time an operator must be present to inspect the data. Since we repeat the cycle under some circumstances by changing the pressure and/or temperature, approximately 2 man-months were required to complete the tests. Therefore, we developed an automated test method using an IoT Edge computer, option I/O modules, and a newly developed software suite. The time to finish the cycle is reduced by ~50%:1 h for MBO/20 min for SBO, and the user can simply check the recorded test results, rather than be present. The system reduces manpower, time, and cost, avoiding unexpected human error.

A spacecraft liquid propulsion system, which is the primary actuator for guidance, navigation, and control, is critical to mission success and safety. However, this system has potential of several undesirable phenomena and component faults that lead to compromised performance or even failure. This study focused on using analytical and physical models to assess and treat risk in the system. To clarify our model-based approach, this paper proposes a framework that covers the modeling strategy and methodology based on the objective and scope of each risk management process. The framework also updates and integrates the model step by step manner through the life cycle of the system. This feature emphasizes the advantages of the model-based approach, such as inheritance and reusability, comparing conventional approaches that use different models for each domain and phase.

Satellites, especially communication satellites, require considerable electrical energy to operate their bus systems and mission payloads. The lithium-ion battery has become a key component of the power supplies of modern satellites, especially during eclipse mode, thanks to its light weight and long operating life. However, traditional application in LEO is limited because their depth of discharge (DOD) has been restricted to 25% (DOD25%) and discharge rate to 1 C. To overcome these limitations, JAXA is evaluating the high-rate discharge operation of batteries in outer space and has achieved results as high as DOD40%. This paper reports the status of JAXA's high-rate discharge and high-DOD batteries.

Disasters caused by local heavy rainfall have occurred frequently during the past decades in Japan. After World War II, the Tokyo metropolitan area experienced repeated flooding, including that caused by the bursting of the Tone River by Typhoon Kathleen in 1947. Corresponding to increasing urbanization, flood control measures were implemented locally in each basin. Simultaneous occurrence of multiple floods is a matter of concern and requires adequate countermeasures. In this study, the need for new measures involving the monitoring of retarding basins, regulated areas, and floodways after heavy rain , was investigated using satellite images, and the importance of a broad-scale simultaneous disaster response was highlighted.

The High Mach Integrated Control Experiment, HIMICO, is a hypersonic flight experiment using a ramjet engine. One of the characteristics of the ramjet engine is the self-excited oscillation of the terminal shock wave called buzz. This study investigated the effects of intake throat height on the buzz characteristics in the ramjet engine. The wind tunnel test was conducted with a freestream Mach number 3.4. The experimental results show a decrease in the buzz frequency at small intake throat height. The measured unsteady static pressure inside the engine suggests that this frequency reduction is caused by the longer time for the terminal shock wave to move under the intake ramps. This forward movement of the terminal shock wave is caused by the expansion of the high-pressure region from the rear side of the engine. Therefore, one reason for the decreased moving speed of the terminal shock wave is the reduced inflow rate while the terminal shock wave is kept inside the engine. Another is the change in the area ratio of the intake throat to the diffuser entrance.

The research and development of space activities such as Moon exploration have been fields of concentration recently. One key technology for successful missions is lightweight, large-capacity energy storage. Knowledge of the high energy density Li-ion battery is indispensable to realize an overnight stay on the Moon and exploration in shadowed areas for the unmanned first step in exploration. This paper presents the status of developing high-energy density Li-ion cells for lunar exploration.

While Japanese Venus Climate Orbiter Akatsuki has made progress in studying Venusian atmospheric dynamics, the chemistry relevant to clouds and the atmospheric trace elements have not been fully elucidated. The key to solving these problems is ultraviolet (UV) absorbers, which are responsible for the characteristic albedo pattern in UV wavelengths at the cloud top altitude. Since these chemistries are coupled with solar radiation heating, it is important to identify the UV absorber(s). We are currently developing the FUJIN-2 balloon-borne telescope as a new platform for observing the solar system planets from the Earth's stratosphere with a wide wavelength range. FUJIN-2 is equipped with a precise three-stage attitude control and pointing system. A high spatial resolution comparable to a satellite telescope can be attained by even a 400-mm small telescope. The present study describes the status of FUJIN-2 development and introduces its observation plan of Venus by UV spectroscopy and high spatial-temporal imaging.

Microsatellites have smaller masses, sizes, and power resources, and shorter development periods than medium and large satellites. A different thermal design method is required for microsatellites. Utilization of the entire heat capacity of microsatellites for the reduction of temperature fluctuation is a feasible thermal design concept proposed previously. A material with high thermal conductivity and specific heat is preferable. Considering the aforementioned design concept, in this study, the thermal effect of using an aluminum–copper–aluminum clad material with a high conductivity is investigated by comparison with those of aluminum alloys A1050 and A5052, which is generally used in satellites. The microsatellite is 30-cm cubic and has two mass conditions of 10 kg (Case 1) and 30 kg (Case 2) as on-board equipment. The orbit of the microsatellite is a sun-synchronous circular orbit at the altitude of 500 km with the descending node at 11:00 am local time. The microsatellite is Earth-pointing. Thermal analyses and thermal vacuum tests are performed to confirm the temperature history of the microsatellite. The hemispherical emissivity and its temperature dependence of the black aluminum tape used in the experiments are also measured because they were unknown. It is found that the averages of the temperature fluctuation range of the clad material are 1.6 ℃ and 2.8 ℃ smaller than those of A1050 in Case 1 and Case 2, respectively, and 5.2 ℃ and 8.1 ℃ smaller than those of A5052 in Case 1 and Case 2, respectively.

Purpose: Microgravity leads to muscle atrophy, especially in antigravity muscles including the psoas muscle (PM). The hybrid training system (HTS) is a method that uses the electrically stimulated eccentric contraction of the antagonist muscle as the exercise resistance of the agonist muscle. The purpose of this study was to examine the effects of HTS on PM. Methods: We investigated the HTS (9 legs) and control (CON) (9 legs) sides. Stimulation electrodes were attached to the gluteus maximus muscle of the HTS side, and electrical stimulation was applied during hip flexion. The CON side performed the same exercise without electrical stimulation. The exercise task was hip flexion from a stationary standing position,10 sets of 10 repetitions per set, 3 times/week for 6 weeks. The comparison of PM volume and muscle strength before and after training was made by using the Wilcoxon signed-rank test. Results: The volume of PM on the HTS side significantly increased (pre: 352.73 ± 33.84, post: 365.90 ± 40.07 cm³) (P<0.01). The hip flexor strength on the HTS side significantly increased (pre: 203.89 ± 41.98, post: 230.22 ± 25.67 N) (P<0.05). There was no significant increase on CON side. Conclusions: HTS increased the strength and muscle volume of the agonist when the antagonist electrically was stimulated.

Au wire bonding has not been yet included in the quality assurance of the standard documents for components in space applications. we extracted technical issues and examined the possible method for verifying them to establish the evaluation test for Au wire bonding. we show the specific issues of Au wire bonding and the results of evaluation of it.

The Martian Moon Exploration (MMX) mission plans to explore Martian moons and return samples from Phobos to Earth. The assessment of potential microbial contamination of Martian moons is important for the mission from the view of planetary protection, and hypervelocity impact sterilization and radiation sterilization are important processes to determine the microorganism probability on Martian moons. In this study, we develop experimental approach for hypervelocity impact sterilization and discuss the development for technical issues on hypervelocity impact sterilization tests. Finally, we demonstrate hypervelocity impact sterilization tests using Bacillus atrophaeus spores.

The integrated fuel cell, which was developed by Honda, was evaluated for the usage in spacecraft. The integrated fuel cell has a new structure that has electrical conductive sections to the horizontal direction along the electrolyte membrane to create higher voltage within a single cell¹⁾. The cooling mechanism was updated with use in outer space in mind, and a fuel cell stack that has continuous holes to allow layering for use in space was developed. The fuel cell stack for use in outer space was subjected to fundamental evaluation of mechanical, vacuum, radiational and low temperature tolerance. A prototype of the fuel cell system was also fabricated in order to confirm the property of redundancy, which could be a distinctive characteristic of the fuel cell stack for use in outer space. This report describes the results from the confirmation of tolerance using the fuel cell stack for use in outer space and presents an overview of the fuel cell system that was fabricated.

Hayabusa 2 spacecraft revealed that a small carbonaceous asteroid 162173 Ryugu is a rubble pile and that its surface is covered with various sizes of boulders. A few percent of surface boulders have cracks. They are classified into straight cracks, sinuous cracks, arrested (incomplete) cracks, and complex (branched) cracks. We analyzed 538 boulders (777 cracks) and found 60% of their cracks have the meridional direction (±15deg from N-S) except complex cracks. Impacts or surface mass movement cannot have produced this preferred crack orientation. Thermal stress from solar heating would have grown cracks to meridional direction. Another feature of thermal stress – exfoliation – is also observed on the surface of small boulders of Ryugu as of Bennu.

DESTINY⁺ is a small-class science program led by ISAS of JAXA and is under development for launch in FY2024−2026 with both engineering and science goals. The engineering goal is to acquire technologies that will enable future high-frequency exploration of the solar system. Ion engine operation technology in planetary orbit and high-speed flyby technology of small bodies will be demonstrated. The scientific goal is to elucidate the characteristics of dust that may have transported organic matter to the primordial earth. In-situ analysis of dust at Earth's orbital position and a flyby observation of the asteroid (3200) Phaethon, the parent body of the Geminid meteor shower, will be conducted. The spacecraft system consists mainly of the proven onboard equipment of the small scientific satellite standard bus and Hayabusa2, and the scope of new development was focused on the minimum necessary to achieve low-cost and short-term development. The science instruments are a dust analyzer for in-situ dust analysis during the entire mission, a telescopic camera and a multiband camera for observing Phaethon's surface. The telescopic camera is equipped with a rotating mirror to track Phaethon during the high-speed flyby with a relative velocity as high as 36 km/s.

This paper reports on our study to develop a magnetic circuit with a discrete outer coil and test its performance in a Hall thruster. The results were compared to those of a Hall thruster on board the Engineering Test Satellite-9 (ETS-9), which used a cylindrical outer coil. At the normalized magnetic field strength of 0.58 for the 6 kW operation using a six poles discrete outer coil, the thrust, discharge current, Isp, thrust-to-power ratio, thrust efficiency, and discharge oscillation were 386.7 mN, 19.0 A, 1890 s, 66.9%, 62.0%, and 13.5% respectively. The Hall thruster with the new magnetic circuit achieved equivalent performance (including discharge stability) to the Hall thruster on board the ETS-9. Further, the study investigated the effect of the number of poles on the performance of the Hall thruster: the performance with the three pole coil was lower than that with the six poles coil. This paper also reports on the effect of pole numbers on utilization efficiency. Mass and current utilization efficiency were more affected by the magnetic field generated by the three poles outer coil.

When a sample return capsule enters the Earth's atmosphere, a strong shock wave is generated in front of the capsule and it receives severe aerodynamic heating. An ablation method is the effective thermal protection one to protect the capsule from the heating. In the future, capsules are expected to be larger and to increase reentry speed. Therefore, the capsules will be exposed to more severe aerodynamic heating. In this experiment, porous carbons with different pore diameters (5 μm, 10 μm, and 25 μm), and porous carbons impregnated with cyanoacrylate were used as test pieces. It was found that there were differences in wear time and wear behavior depending on the test pieces. Furthermore, the effective heat of ablation was calculated by conducting an experiment using an automatic position control system that can detect the tip of the test piece and control it to the target position. The effective heat of ablation of porous carbon impregnated with cyanoacrylate (5 μm) was about 2.8 MJ/kg.

Targeting at medium- to large-scale spacecraft, JAXA is working on the development of next generation's electric propulsion. The milestone mission for this activity is the engineering test satellite-9 (ETS-9) where the first flight experiment of a Japanese Hall thruster is planned. Although this opportunity is limited to about 3,000 hours of full power operation at 6 kW and 200-cycles operation at a lower power of 4 kW, the flight experience in combination with qualification processes, including a full mission duty cycle life test in a ground facility, contributes to establish medium-power Hall thruster technology in Japan. In addition to this first step R&D, improvements of the 1st step Hall thruster are under study as the 2nd step R&D to satisfy the needs for future medium- to large-scale all-electric propulsion Earth orbiting satellites by enabling lightweight thruster body, deep throttling, and higher specific impulse (Isp). Furthermore, as the 3rd step, wide Isp operation are pursuit to provide the optimum Isp for a variety of future missions including space transportation and exploration. These R&Ds enable step-by-step advancement of Japanese Hall thruster technology.

In recent years, there has been a growing trend toward deep space exploration using nano- and micro-satellites, and in the future, planetary exploration such as Mars is expected. In this context, orbital insertion using aerocapture is attracting attention as a method to significantly reduce the mass of propellant and other components. The use of electric propulsion with high specific impulse could be considered as a further weight reduction method, but it has not been studied in detail so far. In this study, we propose an aerocapture method using electric propulsion and drag modulation control using an algorithm called NPC. First, we optimize the control of the thrust direction, which is necessary for the use of electric propulsion, and achieve a significant improvement in efficiency over a previous study. In addition, we compare the results with conventional high thrust propulsion for aerocapture, and show that a significant mass reduction is possible. Subsequent Monte Carlo simulations with various uncertainties showed that the proposed method has a high probability of success and is promising for future planetary exploration.

We are developing an instrument called Circum-Mars Dust Monitor (CMDM), which will be installed on the Mars Moon Explorer (MMX) being developed by JAXA, to observe dust particles in orbit around Mars. CMDM aims at in-situ observation of dust particles orbiting around Mars which has been theoretically predicted to exist ranging between tens of μ m and hundreds of μ m in size with impact speed ranging between about 0.2 km/s and about 0.8 km/s. The instrument must be a real-time dust instrument with high reliability to distinguish actual dust events from noise-induced trigger signals. In addition, a sensitive area of at least 1 m² is required to study the undiscovered Martian dust ring particles to the background flux level of interplanetary dust particles. However, the resources required for the spacecraft system must be kept to a minimum. The sensor unit of the CMDM consists only of a polyimide film and a piezoelectric ceramic element. Therefore, the sensor has a sensitive area of 1 m² but a mass of only a few hundred grams. In this paper, we describe the results of solving the technical problems in sensor development identified during the conceptual design and the results of sensitivity measurements using a sensor BBM made from materials actually used in-flight products.

The low reliability of nano-satellites has become a problem, with the cause often being defects in the design and manufacturing process. It is required to take sufficient measures to account for those defects through ground tests to reduce on-orbit failures. The model-based fault diagnosis method can prevent oversight of fault candidates and compensate for a lack of knowledge in human-based diagnosis. Since the developing time of nano-satellites is limited, fault diagnosis needs to be carried out efficiently also while being careful not to cause secondary faults. This paper presents a novel method that can find the appropriate command to verify candidates of fault location in a satellite. The proposed method can consider the side effect of a command and the efficiency of narrowing them down by using the information transmission path model. In the case study, the proposed method is applied to an earth-orbiting 6U CubeSat, and this case study reveals that 33% of the failures that occur in ground tests can apply to the proposed method. In addition, the proposed method can identify the fault location of some of these failures, although there are several limitations to implementing the proposed method in the actual satellite development.

This study highlights the team management capabilities of astronauts, working in team-centric environments in space and on Earth, referred to as the Space Flight Resource Management (SFRM). The purpose of this study is to examine the effectiveness of the SFRM-incorporated training, when it is used on an online platform for terrestrial remote workers. The training was designed for teams with four players, working collaboratively to accomplish the assigned missions. To reproduce a remote working environment, each player was isolated and only allowed to communicate with their teammates using a video conferencing tool, without any visual transmission. The participants played a board game designed to impart the SFRM skillset, including skills of communication, situational awareness, decision making, teamwork, leadership, and followership. An experimental training session was conducted with a group of remote workers (n=40), and their SFRM skill level was measured using 39 questions before, immediately after, and one month after the training. The results revealed statistically significant improvement in most SFRM skills. The findings contribute to the establishment of new ways of remote working and learning, under the COVID-19 impacted work situations, especially with regard to team behavioral skills.

An all-in-one unpowered Mars rover functions as an entry, descent, and landing (EDL) system during atmospheric entry, an airbag at landing, and an unpowered rover at Mars surface with an inflated spherical structure. This rover can access attractive but risky areas, where typical rovers cannot access. This study considers the feasibility of this rover in terms of EDL and roving, and demonstrates the required rover configuration that enables both robustness in EDL and mobility on Mars surface.

The Sail-propelled Explorer to Investigate the Make-up of Enceladus Icy plumes (SEIMEI) is a 340 kg satellite to shed new light on the origins of life in the solar system. The SEIMEI concept is proposed by a team of graduate students from five countries at The University of Tokyo. By contrast with conventional, multi-payload, many-objective deep space missions, SEIMEI is a small low-cost satellite with only a few payloads and focused objectives. SEIMEI's primary mission objective is to sample the plumes of Saturn's moon Enceladus via a simple and compact satellite architecture, using two state-of-theart mass spectrometers. The second mission objective is to demonstrate the feasibility of planetary orbit insertion via electric propulsion by solar power sail at Saturn. This paper summarizes the promising results of the SEIMEI concept study. Based on the analysis, SEIMEI will be launched on an HII-A or HIII rocket in 2033, deploy its 600 m² sail, follow a low-thrust trajectory with a gravity assist at Jupiter, and arrive at Saturn in 2044 having consumed around 70 kg of fuel. The paper also presents the online student-led collaborative design methodology adopted in response to the coronavirus pandemic.

The use of small satellites is becoming more and more widespread, not only for technological demonstration but also for actual applications. In flight software development for small satellites, it is becoming more and more important to reduce the cost of implementation and verification and to make the software highly reusable. Therefore, we have improved the C2A (Command-Centric Architecture), flight software architecture, which has been developed in our previous satellite development projects, to enhance its reusability. In addition, we are developing a verification scheme that introduces the concept of Continuous Integration to reduce the verification cost and improve the reliability of the software. We have open-sourced C2A and its related software and tools, and released them on GitHub.