The microwave discharge ion thruster μ10 contributed to the success of the Japanese asteroid exploration missions Hayabusa and Hayabusa2. During these space missions, the performance degradation, i.e., the decrease of propellant utilization efficiency, was observed to be approximately 16% after an operation time of 9,000 h. In contrast, no degradation was observed in the ground test. Previous work has indicated that the main cause of the degradation is carbon contamination from ion sputtering in the accelerator grid. This work also suggested that a facility effect occurred, i.e., back-sputtered materials at the wall of the vacuum chamber canceled out the performance degradation. In this study, we propose a time-saving method by increasing the accelerator current, for use in efforts to reproduce the carbon contamination without the facility effect. The approach was successful in reducing the operation time by a factor of 90–200, while the degradation observed during space operation was reproduced. Finally, we attempted to recover the thruster performance by offsetting the microwave frequency by −0.06 to +0.01 GHz from the conventional microwave frequency of 4.25 GHz. The experimental results indicate that the performance degradation can be mitigated using a lower microwave frequency, which provide some ideas to suppress performance degradation in space missions.

This study developed a novel thermal control system to cool detectors of the General AntiParticle Spectrometer (GAPS) before its flights. GAPS is a balloon-borne cosmic-ray observation experiment. In its payload, GAPS contains over 1000 silicon detectors that must be cooled below −40℃. All detectors are thermally coupled to a unique heat-pipe system (HPS) that transfers heat from the detectors to a radiator. The radiator is designed to be cooled below −50℃ during the flight by exposure to space. The pre-flight state of the detectors is checked on the ground at 1 atm and ambient room temperature, but the radiator cannot be similarly cooled. The authors have developed a ground cooling system (GCS) to chill the detectors for ground testing. The GCS consists of a cold plate, a chiller, and insulating foam. The cold plate is designed to be attached to the radiator and cooled by a coolant pumped by the chiller. The payload configuration, including the HPS, can be the same as that of the flight. The GCS design was validated by thermal tests using a scale model. The GCS design is simple and provides a practical guideline, including a simple estimation of appropriate thermal insulation thickness, which can be easily adapted to other applications.

Artificial neural networks are used to predict discharge current, a series of the performance of Hall thrusters. To reduce the calculation cost of predictions, we used an Echo State Network (ESN), a model of reservoir computing, for the prediction of the one-step-ahead and ten-step-ahead discharge current dependency on magnetic field strength using operation parameters and plume images as an input data set. The mean absolute percentage error was 1.70% and its calculation time was about 1.0 second, which was about 1/20,000 in comparison with our previous study using Feedforward Neural Network and Convolutional Neural Network.

A magnetic thrust chamber is a component of a propulsion system which obtain thrust from a plasma in using a magnetic field. A laser fusion rocket has been assumed to use this system and will be used for interplanetary flight. To realize this rocket, we need to investigate the thrust performance of the magnetic thrust chamber. Previous work has investigated the plasma behaviour and thrust using simulations and experiments. However, quantitative comparison has not been much. In this work, we quantitatively compared computational and experimental results of plasma behaviour and performance. It was observed that the impulse bit increases as the magnetic field strength however, results from the experiment were approximately 3 times larger than that of simulation.

Combustion behaviors of non-metallized and metallized gel droplets were investigated using a single-droplet experiment method. Ethanol–methylcellulose gel was used to represent non-metallized gel. Mixed ethanol–methylcellulose gel with 10% aluminum powder was used to represent metallized gel. The gel droplet combustion was conducted at an oxygen concentration of 21% (air), 30%, 40% (enriched air) under an initial convection temperature of approximately 700 (±20) °C. The initial droplet diameter was 2.27 ± 0.19 mm in combustion tests. Furthermore, combustion residuals of metallized gel droplets were analyzed using scanning electron microscopy (SEM) and X-ray diffractometer (XRD). Observations showed that non-metallized and metallized gel droplets led to different combustion behaviors. The combustion of metallized gel droplets showed a particular phenomenon recognized as Al agglomerate period before the end of combustion, which could contribute to the release of more energy of gel fuel. Effects of oxygen concentration on combustion were also evaluated. The results suggest that increased oxygen concentration might improve the combustion efficiency of both non-metallized and metallized gel fuel. Furthermore, for metallized gel droplets, the stage of Al reaction is advanced.

In recent years, planetary exploration has received increasing interest from both the research community and the private sector, with ambitious missions being planned to target the Moon, Mars, and beyond. In the scope of the Artemis Project, multiple space agencies and industry partners are working together to safely send humans back to the Moon, intending to eventually create sustainable on-site settlements. Lunar caves are good candidates for long-term habitats thanks to their almost constant temperature and protection from space radiation. These caves are accessible mainly through skylights that are hypothesized to have been created by the collapse of cave ceilings, making the terrain at the bottom of the skylight particularly rough, with a mix of large debris and finer regolith. Exploring these entrances represents a significant challenge for current robotic platforms due to the roughness of their environment and the limited solar exposure at the bottom of the skylight. This study investigates a disruptive solution to safely transport large payloads together with robotic units, referred to as Base Units (BUs), from the skylight's rim to the bottom of the cave by leveraging the innovative concept of Robotic Exoskeletons (REs). In this work, a lunar cave with approximated dimensions for the skylight of 100m in diameter and 100m deep is considered as a representative test environment. BUs are approximated as compliant spheres of 20cm in diameter and a mass of 400g. This paper explores the design and evaluation of a fully passive Multi-Unit RE weighing less than 10kg and able to transport 6 BUs and a payload in a free-fall scenario from the skylight’s rim to the bottom of the cave. The energy of the free-fall is reused to propel 3 of the BUs further into the cave to act as scouting units, while the 3 others will remain at the point of impact to unpack the delivered payload. Our current non-space-ready design relies on flexible carbon fiber blades and shockabsorbing honeycomb structures. Our tests show promising qualitative results in terms of energy dissipation and BUs propulsion.

An ionic liquid electrospray thruster (ILEST) is expected as a propulsion system to control nano-spacecraft attitude and orbit in space. To produce high thrust, we propose a two-stage electrodes thruster which has an accelerator in addition to an extractor. We fabricated emitters on a silicon wafer and the two-stage electrodes on a glass substrate with through holes using microfabrication technologies. We confirmed that the fabricated ILEST emitted ions or droplets from the silicon emitters at near 2.0 kV. The maximum current in single-stage operation was -570 µA at 3.0 kV. With two-stage operation using the accelerator, ion emission happens at high voltage without conduction between the extractor and the accelerator. The results show the advantage of the fabricated two-stage electrodes. Currently, extracted ions were captured mostly at the extractor because of the poor alignment of the emitter silicon chip to the jig.

We plan to conduct an automated docking demonstration using the HTV-X on-orbit demonstration platform. This is a preliminary technology demonstration to contribute to resupply to Gateway, the manned lunar orbit base. Since docking could cause a catastrophic hazard such as ISS collision in case of failure, comprehensive safety analysis is necessary. Since the HTV-X and the docking system are both complex systems, and they must work properly together, we have adopted STPA, which can comprehensively analyze each other's behavior with guide words. In addition, we also proposed an STPA extensible framework for Multiple Controller problem based on this case study.

The General Antiparticle Spectrometer (GAPS) is a balloon-borne experiment that aims to measure low-energy cosmicray antiparticles. GAPS has developed a new antiparticle identification technique based on exotic atom formation caused by incident particles, which is achieved by ten layers of Si(Li) detector tracker in GAPS. The conventional analysis uses the physical quantities of the reconstructed incident and secondary particles. In parallel with this, we have developed a complementary approach based on deep neural networks. This paper presents a new convolutional neural network (CNN) technique. A three-dimensional CNN takes energy depositions as three-dimensional inputs and learns to identify their positional/energy correlations. The combination of the physical quantities and the CNN technique is also investigated. The findings show that the new technique outperforms existing machine learning-based methods in particle identification.

JAXA is focusing on SLR as a method to grasp the orbit and attitude motion of space debris. If the target object is equipped with an SLR reflector, even after it becomes debris, SLR enables orbit determination with high accuracy and spin rate estimation. Because conventional SLR reflectors are expensive, heavy, and large, we developed a small, lightweight, and inexpensive reflector, named Mt.FUJI, which is demonstrated on the experiment on the HTV-X. While most of the previous studies on attitude motion estimation by SLR have only estimated angular velocity by frequency analysis of SLR data, we propose a method to estimate both attitude and angular velocity using global optimization. Specifically, we construct an evaluation function including a penalty term considering the reflection condition of the reflector and apply a genetic algorithm to search for a solution. Through numerical experiments, it has been found that even when observations are polluted by noise and bias, the proposed method can provide a good solution that can be used as initial values for filter theory and batch processing and without a priori of attitude state. This paper describes the overview of Mt.FUJI and the experiment, the estimation methodology, and the simulation results.

Carbon Fiber Reinforced Plastic (CFRP) is often used for space telescopes because of its high specific stiffness and low coefficient of thermal expansion. CFRP structures with complex 3D shapes are expected to be used in the near future, and thus the thermal deformation characteristics of CFRP structures with an arbitrary shape is needed to be studied. In this research, the out-of-plane thermal deformation of CFRP laminated curved plates was evaluated using experiments and numerical analysis. Two L-shaped curved parts with different curvature radii were considered. The experimental results showed that heating increases the curve angle, and the difference in curvature radius does not affect the curve angle change. The result of numerical analysis using solid elements shows the qualitative agreement with the experimental result. However, there is an approximately -30 % difference to be solved between the calculated and measured curve angle change.

The granular flow around a cylindrical corer having a forward-facing cavity on its front surface to catch the sample of the regolith layer over a celestial body was numerically analyzed using the compressible and non-expanding (CNE) fluid model. The equation of state is composed of the irreversible compression curve and the elastic unloading/recompression curves. The speed of sound for the elastic process is much higher than that in the irreversible compression, resulting in almost monotonic increase in the density through the processes. The quasi-static compression experiments by the hydraulic press for glass beads and silica sand showed that their response to the compression is appropriately described by the CNE fluid model. In the result of the one-dimensional shock tube problem with the CNE fluid model for the glass beads, the formation of the precursor wave that corresponds to the elastic process was observed in front of the shock wave that corresponds to the irreversible compression. Finally, the regolith flow around the sampling corer was numerically analyzed solving the axi-symmetric Euler equations on the moving coordinates. The characteristic features of the granular flow including the filling process of the sample were discussed.

A ballute is an inflatable device to enhance the drag force during reentry flight. The device was inflated by high-pressure gas. Thus, the device has a flexible structure. This means the ballute shape can change depending on the pressure on the body surface, and the drag force by the ballute will also change. Therefore, it is essential to estimate the effect of flexibility on the deformation of the ballute. In this study, we performed a fluid-structure interaction (FSI) simulation to estimate the deformation of the ballute using ANSYS AIM. In addition, we compared the experimental results by ISAS supersonic wind tunnel. The numerical results showed that the FSI simulation could reproduce how the aerodynamic force deformed the ballute.

Atmospheric gravity waves transport momentum in the atmosphere and play an important role in determining temperature and wind distributions through driving the meridional circulation in the middle atmosphere. However, they have wide spatial and temporal scales, which make it difficult to capture the whole feature of gravity waves with any of the latest observations and models. The first Mesosphere-Stratosphere-Troposphere (MST)/ Incoherent Scatter (IS) radar in the Antarctic, PANSY, which was installed at Syowa Station (69.0S, 39.6E) in 2011, can directly estimate the momentum flux of gravity waves in all frequency bands by observing 3-dimensional winds with high precision and high resolution. On the other hand, the super pressure (SP) balloon observation can also estimate momentum transport due to gravity waves in all frequency bands, and its horizontal distribution is also clarified. In order to carry out this SP balloon observation in the Antarctic where the observational constraint on the momentum transport due to gravity waves is especially insufficient, we proposed the LOng-Duration balloon Experiment of gravity WAVE over Antarctica (LODEWAVE). By combining it with the PANSY observation, observational constraints on the momentum transport due to gravity waves in climate models are acquired, which contributes to the improvement of the future prediction by the climate models.

The evaluation guidelines for COTS parts are highly demanded by space craft manufacturers because COTS parts are expected to apply to the space craft due to its advance performance and low cost. The guideline for COTS AD/DA converter which was made based on the results of radiation irradiation experiment of actual COTS AD/DA converters is introduced in this paper. According to the results of experiments, characteristic phenomenon which related to the internal register of recent COTS AD/DA converter was observed during the single event latch-up and transient test. Our guideline suggests the appropriate evaluation methods which consider the effects of such characteristic phenomenon. Following the AD/DA converter, SSD and MEMS are planned to make guideline based on the experimental results of actual device. Also, these will be evaluated focusing on some specific points such as radiation effects or mechanical stress tolerance.

The asteroid explorer Hayabusa2 returned to the earth after 6 years of interplanetary voyage in Dec, 2020. After separated from the spacecraft, the sample return capsule (SRC) entered the earth atmosphere and passed through excessively severe aerodynamic heating at the velocity of about 12 km/s. Finally SRC successfully landed on the ground in the Woomera Prohibited Area (WPA), South Australia on Dec. 5th (UTC). The present paper overviews the reentry operation and the post-flight analysis of the Hayabusa2 SRC. The post-flight analysis program is currently still underway and is mainly comprised of the following three activities; 1) Reproduction of the best estimated trajectory (BET) based on the reentry orbit determination, ground observations, the SRC landing point and the atmospheric properties on the reentry day. 2) Analysis of flight parameters regarding vehicle dynamics during reentry and aerodynamic heating by using flight data recorded by the reentry flight measurement module (REMM) installed in the SRC. 3) Nonintrusive inspection of the recovered heatshields through X-ray CT-scanning and 3-dimensional laser-scanning. These activities will be finally integrated and studied for the SRC design validation purposes and for better understanding of the flight results.

In order to use laser ignition systems for solid rocket motors operating in deep space environments, it is necessary to elucidate the laser ignition characteristics of the ignition charge in low-temperature environments. This study aims to design an experimental system that can confirm the ignition threshold, ignition delay, and ignition temperature by irradiating an ignition charge with a diode laser in a low-temperature environment. Ignition experiments at room temperature were conducted. The data were evaluated statistically to obtain an ignition threshold with the maximum likelihood method. The relationship between the laser irradiation duration and the laser power with respect to the ignition threshold was obtained. The target value of the low-temperature environment temperature was determined as -50 °C. We examined the requirements of the experimental system and conceptually designed the system to simulate the low-temperature environment. It was confirmed that the constructed experimental system cooled the ignition charge to -50 °C. In the ignition experiment, the ignition charge was successfully ignited at room temperature and at low temperature. The ignition delay, the ignition temperature, and the high-speed image were obtained. Eventually, the validity of the experimental system was confirmed through the function tests.

In lunar exploration, many mechanisms require a solid lubricant coating with high wear resistance. The lunar surface is covered in dust particles that cause aggressive wear when they get between sliding surfaces. Conventional wear-resistant coatings for tools have high friction coefficients in vacuum, and they have inadequate wear resistance to lunar dust. This study focused on diamond coatings. A diamond coating deposited by CVD has excellent wear resistance against dust particles in vacuum, although its coefficient of friction is not as low as other solid lubricants. Also, it is difficult to coat metal surface with diamond because of the difference in their coefficients of thermal expansion. Our study approached the solution to the problem of applying a diamond coating to sliding mechanical surface on the Moon. One approach to finding a suitable coating for metal and ceramic substrates was to change the deposition method and conditions. Another approach was to deposit an intermediate layer between the diamond coating and the substrate to mitigate the internal stress of the coating. Our study showed that a hard diamond coating could be deposited on a metal surface with an intermediate layer if its coefficient of thermal expansion and compatibility with diamond coating were optimized. A diamond coating could be deposited at 250 ºC lower temperature than normal method, and the coating had almost the same tribological characteristics as coating at normal temperature. Low wear was observed for the coatings under low load in vacuum.

Earth observation by nano/microsatellites has been developing rapidly over the past decade and its huge amount of data is commercially available. Although nano/microsatellites developed by universities also produce Earth observation data, their data has been manually processed and poorly managed. In this study, a satellite data management system called satellite data manager (SDM) was developed for automated processing and effective management of the Earth observation data from microsatellites. In SDM, the data processing is automated using open-source libraries and open data satellite imagery, and the data visualization is implemented by an interactive map. The feasibility of the system is demonstrated by Earth observation data acquired by the sensors on the RISESAT microsatellite.

We investigated obstacles (craters and boulders) at five long-term sunlit areas of the lunar south pole as candidate landing sites for future JAXA medium-sized landing demonstration missions to the Moon. We used resampled ortho-rectified images and digital elevation models of 1 m/pixel resolution, originally 4096 pixel / deg (7 m / pixel along meridian) resolution, based on data acquired by the Lunar Reconnaissance Orbiter and SELENE (Kaguya). The numbers of craters were found to be mostly 15–25 (> 7 m in crater diameter) in a 50-m-radius-circle, similar to that of the area around the Apollo 16 landing site. Almost no boulders exist at particularly high terrains (from the highest to about 10–20 m down) in each investigated region, where the longest illuminated conditions are expected, probably because the areas are so topographically high that boulders ejected by nearby crater formations did not reach the site.

In microwave discharge ion thrusters, the plasma mode transition and hysteresis are two of the most important phenomena. The plasma mode-transition restricts the maximum thrust, whereas the hysteresis splits two mode of thrust in each flow rate. The previous thruster development and optical measurement indicates that there are two important parameters; the volume of magnetic confinement region where electrons are continuously heated by ECR and confined by magnetic mirrors, and waveguide plasma conditions. In this study, several kinds of thrusters that have different volume of magnetic confinement were tested. In addition, the enhancement of cut-off density inside the waveguide using larger diameter of waveguide was proposed. Experimental results show that the volume of magnetic confinement region is a key parameter for determining the maximum thrust. It was also found that the enhancement of cut-off is effective to eliminate the plasma hysteresis, whereas it does not affect the plasma mode transition. Hence, to further improve the thrust, one possible solution is to enhance the cut-off density at downstream of waveguide outside of magnetic field confinement region.

Perovskite solar cell (PSC) is a novel type of solar cell and can be thin, flexible, and lightweight. Considering these advantages, PSCs have the potential for applications in inflatable structures, such as plastic film-type balloon envelopes. This study developed PSCs to form on the plastic envelopes of stratospheric balloons used in space science researches. The development was performed step-by-step by proceeding with both ground (in the laboratory) and flight (in the actual flight environment) tests iteratively in parallel. As the first step of flight tests, a small test platform system was developed to evaluate small PSC specimens, and their flight tests in 2019 and 2021 were performed by hanging them beneath rubber balloons. In the course of the balloon flights reaching a maximum altitude of around 30 km, the current–voltage curves of PSCs were obtained as expected. The platform housekeeping data was also obtained and it was found that the platform attitude was significantly and continuously disturbed throughout the flight. The obtained attitude data will provide guidelines not only for the current project but also to design other rubber balloon experiments.

Satellite imagery has been widely used to study of monitoring and analysis of coastal areas considered low cost and its ability to collect information over large spatial areas. The aim of this study is to detect the coastline changes corresponding to coastline changes in volumes of coastal sediments in Jembrana Regency, Bali Province by using Sentinel-2 imagery data. Sentinel-2 image has high spatial resolution and high temporal frequency compared with other free-access satellite imagery data with 10 m spatial resolution. Coastline measurements provide information about the coastal loss over the years and the changes in the coastal area. This study is essentially needed in order to identify areas that require further investigation and protection. Results showed during the initial 3-year period from 2016 to 2019 coastal area of Jembrana Regency changed with total accretion of 226,999 m² and erosion of 586,042 m². The shoreline changes from 2016 to 2019 are highly changed due to the morphological changes in the coastal area.

This paper presents the latest GPS receiver (GPSR) developed by JAXA and NEC/NEC Space Technologies, Ltd. for satellites in geosynchronous orbit (GEO). The flight results are also presented. This is the first time a Japanese GPS receiver has used GPS signals above the GPS constellation of satellites. GPS was originally designed for ground users; thus, there are difficulties in receiving signals from GPS in GEO. GEO satellites fly above the GPS constellation orbit. Therefore, they can receive signals from GPS satellites positioned on the other side of the earth. The expected signal strength is weaker in GEO than observed in low Earth orbit (LEO) since the distance from GPS satellites to the receiver is farther than from LEO satellites. With improved sensitivity, the new GPSR was developed to receive these weak signals and determine satellite position using a highly directional antenna optimized for GEO. This receiver was mounted on an optical data relay satellite launched in late 2020. The GPSR successfully activated and determined the satellite's position. This paper introduces the flight results and evaluates the receiver performance in orbit. The results indicated that the GEO GPSR could determine its position within 100 m. This accuracy is far beyond the conventional range and range rate (R&RR) method for determining a satellite's orbit.

JAXA operates scientific balloon campaigns, aiming at obtaining scientific results through safe and reliable balloon flights. The development of the prototype of the flight prediction and control system began more than 20 years ago. It has become a mature system through many years of operation and functional enhancement and modification. The main functions of the system are implemented by a database system, which has been used for at least 82 heavy balloon experiments and 102 light balloon experiments since 2007. The applications used in client computers include more than 180 graphical user interface panels. The system is designed to incorporate redundancy for availability during balloon flight operations. Although various constraints face balloon flights, such as scientific requirements, flight safety, and severe high-altitude wind conditions, the flight prediction and control system enable us to construct a detailed flight plan and to control the flight based on predictions. In addition to the report of the system, flight prediction is explained with an example of boomerang flight control planning.

Hayabusa2, a Japanese asteroid sample return probe, arrived at Ryugu, a C-type asteroid, on June 27, 2018. It departed from Ryugu on November 13, 2019, after completing all the missions in the asteroid proximity phase. The propulsive cruise with its ion thrusters began on December 3, 2019. After the cruising phase, the spacecraft moved to the final precise guidance phase to the Earth, and it released its reentry capsule on December 5, 2020. The reentry capsule was retrieved in Woomera Prohibited Area in Australia, then Hayabusa2 completed its nominal mission. After the capsule release, the Hayabusa2 spacecraft performed a divert maneuver to pass by the Earth and started its extended mission to rendezvous with 1998 KY26, a fast-rotating small near-Earth asteroid. This paper outlines the trajectory design of the extended Hayabusa2 mission. First, the selection of target bodies is described, and then the trajectory design, including an asteroid fly-by and multiple Earth gravity assists, is shown.

Shape memory polymers (SMPs) are smart materials with shape memory effect that can be used in various industries. Recently, deployable structures are widely used for space structures. Using SMP as a material for deployable structures can have the benefit of high reliability, low weight, and cost-effective compared to usual deployable structures. In this study, a self-thickness-growing sandwich panel with multiple wavy core strips crisscrossing each other in orthogonal directions is proposed to have isotropic bending stiffness in orthogonal directions. SMP is used as the upper and lower layers of the structure. The core is made of a polyethylene naphthalate (PEN) sheet and is attached to the SMP sheet with double-sided tapes. The panel can be compressed to a thin layer at high temperature and can keep the shape after cooling down. After being heat up again, the structure can return to its original thickness with high bending stiffness. Through multiple times of deployment experiments, it is shown that the structure is functional and is possible to be used for application after further development for specific purposes. Bending test shows that the panel has reasonable bending stiffness while they have an individual difference in the bending stiffness because all of them are handmade.

In this study, we investigated the deformation of the ballute using the supersonic wind tunnel experiments at ISAS/JAXA. Due to the very high dynamic pressure of the wind tunnel, it was challenging to create the unbreakable flexible test model using the inflatable structure by gas. Thus, we tried to approximate the flexibility of the inflatable test model by using a hollow torus with a flexible material like hard rubber. The test model was the hollow polygon with a 50 mm major radius and 20 mm cross-sectional diameter. The test model was made of a flexible resin using a 3D printer Form 3. The flexibility of the test model was controlled by changing the inner diameter of the cross-section. In the wind tunnel experiments, the deformation of the ballute and the aerodynamic force on the ballute was measured. Experimental results showed that as the inner diameter increased, the deformation of the test model increased. The above result suggests that the current test model can simulate the flexibility of the test model. As for the drag force, as the inner diameter increased, the drag force decreased. This result suggests that the deformation of the ballute significantly impacts the drag force.

The present study numerically investigates the flow fields of the twin- and tandem-jet in crossflow (JIC) under supercritical pressure. An ILES/RANS method is applied. In the present JIC, cryogenic hydrogen jets are vertically injected to warm temperature hydrogen crossflow through circular injector holes. The results show that the penetration of cryogenic jets into the crossflow is more enhanced with the tandem-jet than with the twin-jet. The higher jet penetration with the tandem jets is caused by a downstream jet behavior less influenced by the crossflow. Nevertheless, a mixedness analysis indicates that the mixing state at a downstream position is almost similar between the tandem and twin jets.

A high-density emitter array for ionic liquid electrospray thrusters is fabricated using a fabrication process for field emitter arrays. SU-8 photoresist is employed as the inter-electrode material. The distance between the emitter and the extractor electrodes was about 10 µm, enough for about 1-kV application between the electrodes. Adequate thrust and specific impulse can be obtained without an accelerator electrode, resulting in a simple propulsion system. Optimal parameters for the photolithography of SU-8 are determined. The fabricated emitter array had a tip diameter of about 1 µm and an emitter density 200 times higher than conventional emitter arrays. A preliminary experiment for ion emissions showed a negative current of −79 nA at 800 V using a 4 × 4 emitter array in an area of 100 µm square, resulting in approximately twice as much current density as a conventional type.

In this paper, we introduce an alternative approach for design, development and testing of electronics for space applications based on the reuse of tested and qualified design elements. Electronics for space applications have long development cycles due to mission specific design, part level qualification, functional verification and unit level testing. We try to reduce the recurrent work by reusing and testing design elements on circuit level. Associated with the recurrent work are overlooked design flaws and long test campaigns leading to risks for budget and schedule. To mitigate these shortcomings, we propose an alternative approach for development, implementation and testing based on the reuse of circuit-level layout elements, which we call Snippets. Besides schematic and PCB layout, simulation models, test results and documentation can be added on the same level. Limiting Snippets to a single function allows simple functional verification and maximizes reusability. The advantages of early testing include the elimination of parts unsuitable for the mission, validation of the expected functionality and verification of the PCB layout.

A quasi-transient simulator called Hypersonic Engine Dynamic Simulator (HEDS) was developed for the purpose of analyzing the transient behavior of the scramjet engine system. However, the transient analysis along the flight profile could not be performed with HEDS. In this study, an inlet performance calculation model and a quasi-one-dimensional analytical model for hydrocarbon fueled combustor were introduced into HEDS. In order to confirm the response of the fuel supply system with the improved HEDS, the transient analysis along one flight profile was performed. The engine model used in the transient analysis was based on n-dodecane as fuel and electric pump cycle as engine cycle. In addition, the behavior of the engine system due to the control was also studied by comparing the uncontrolled and sequence-controlled cases. As a result, when hydrocarbon fuels were used, a sudden decrease in fuel flow rate was observed around the critical point. Therefore, it was necessary to control the fuel flow rate so that it does not fall into a state near the critical point. In addition, the fuel flow rate was controlled to increase by two methods: increasing the pump speed and dumping the fuel. These two methods showed different characteristics to the engine system. Using the dynamic simulator to analyze the behavior of the engine system during the flight profile was effective for considering the engine system and developing control methods.

Magnesium has attracted widespread attention as fuel. Previous studies have suggested that the state of the residue oxide changes with the oxidizer. This study investigated the magnesium combustion plume by changing the oxidizer and its pressure. An ignition system was developed to observe the self-ignition of magnesium in a fixed position. We obtained the photographs of the field where the magnesium-based alloy plate, which was 5×5×0.5 mm³, was burnt. When the oxidizer was oxygen or air, most of the oxide dispersed, but in the case of water vapor, much of the oxide remained near the fuel. The mass differences before and after combustion also show that most of the oxide generated remained on the fuel setting prate. The residual oxide size was found to depend on the oxidizer pressure. As the pressure decreased, the size of the residual oxide that was generated became larger. The spectroscopic measurement of the combustion flame did not observe the MgO lines around 500 nm, suggesting that the gaseous magnesium oxide was not generated or immediately condensed.

In the liquid propulsion system for a spacecraft, a dynamic behavior, such as pressure fluctuation caused by valve operation, can change significantly because of gas bubbles that become trapped in pressure conduits or pipe joints. This study has developed a dynamic response model that considers the trapped gas effect to improve the system-level integrated model of a spacecraft liquid propulsion system. The trapped gas effect in the proposed model is expressed as a volume change of gas phase remaining in a branch pipe by the equation of state for a real gas and the energy balance equation based on a gas-liquid accumulator model. The proposed model can reproduce effects caused by gas volume changes due to pressure wave fluctuation and reflection. A fundamental test was conducted in a simplified single-pipe system of a spacecraft propulsion system to validate the proposed model. The test confirmed that the proposed model could reproduce the peak frequency of the pressure history with a relative error of less than 7 % and the water hammer surge pressure with a relative error of less than 9 %.

This paper aims to construct a method to design low-energy transfer trajectories from Jupiter to Europa, one of the Galilean satellites, via ballistic transfer. The proposed method leverages the periapsis Poincaré map and efficiently finds ballistic transfer trajectories from Jupiter’s sphere of influence (SOI) to Europa. For the evaluation of the proposed method, the interplanetary trajectory from Earth to Jupiter’s SOI is designed by using existing methods, the Tisserand map and multiple gravity assists. By solving the phasing problem that considers the ephemeris of the celestial bodies, this map can find interplanetary trajectories with various transfer times depending on the launch date. It is demonstrated that the periapsis Poincaré map can efficiently find all ballistic transfer trajectories in the region, although it depends on the initial conditions. Furthermore, by applying the clustering method to the obtained ballistic transfer trajectories, we succeeded in finding families of ballistic transfer trajectories existing within the region.

In the project CALLISTO (Cooperative Action Leading to Launcher Innovation in Stage Toss-back Operations) DLR, JAXA and CNES are jointly developing and building a demonstrator for a vertical take-off, vertical landing rocket first stage. CALLISTO aims to gather essential knowledge for the development of operational reusable rockets by demonstrating relevant technologies that are enabling recovery and reuse and leading to a sustainable launcher market. The development and design of reusable load-carrying structures is particularly challenging since the components have to withstand a variety of complex maneuvers for multiple times while keeping the maintenance and repair operations between flights to a minimum. This paper focuses on the progress of the design for CALLISTO's fairing, the deployable aerodynamic surfaces and the stowable approach & landing system considering the specific sizing load cases and functions during the mission. The challenges of movable structures, mechanisms and reusability are especially highlighted paying close attention on the detailed and individual design processes that are accompanied by analyses and pre-tests.

Low Earth orbit (LEO) has become a proving ground for advanced solar sailing missions. The combination of solar radiation pressure thrust and active attitude control enables various fuel-free orbital maneuvers. However, it is challenging to accurately predict the orbit of a LEO solar sail during early-stage mission design. The sail is highly sensitive to solar radiation pressure and air drag, both of which are strongly time-dependent in LEO due to eclipses and space weather respectively. In addition, the impact of active attitude control on the orbit should be accounted for. Analytical orbit-attitude modelling offers one attractive, but usually simplistic option, ignoring eclipses and air drag. In response, this paper develops a new analytical orbit-attitude model with a more realistic LEO environment, via the variation of parameters method. As a case study, the model is used to design a LEO space science mission with solar sailing and active attitude control. The orbital predictions are verified against a high-fidelity coupled simulator. Good accuracy is obtained, with a four to six order of magnitude reduction in computation time. The model offers a fast, accurate tool for early-stage mission design of next-generation LEO solar sails.

This paper proposes a systematic method to calibrate and predict atmospheric densities up to 5 days ahead only from publicly available data. The proposed method consists of calibration of an empirical atmospheric density model and prediction of the calibrated model. The calibration is based on dynamic calibration of atmosphere, which dynamically assimilates the empirical model with orbit-related measurements. The calibrated model is compared with accelerometer-derived densities of the CHAMP satellite, and it is confirmed that errors relative to empirical models decrease by about 17-34 % on a high solar activity day. For the prediction, this paper adopts the Hermitian-Space Dynamic Mode Decomposition with control algorithm, which is a data-driven model discovery approach considering exogenous inputs. Prior studies use time series of density data as input for the algorithm, whereas the proposed method uses time series of density data and its time derivative. The proposed prediction method is validated against accelerometer-derived densities of the CHAMP and GRACE satellites. The proposed prediction method improved errors relative to empirical models by about 26-28 % to forecast densities 5 days ahead. The development of accurate density forecast model in this paper will contribute to the safety of satellite's operations in this congested and contested space environment.

The solar array paddle is an important component that provides power for operating satellites in space. As the complexity of space missions gradually rises, satellites with a high specific power, or power generated per weight (W/kg), with a large number of solar cells on the solar array paddles, are required to meet this rising complexity. Furthermore, the rigidity of the solar array paddle to maintain attitude control in a satellite system requires a natural frequency of 0.1 Hz or higher, as recommended by the Japanese Aerospace Exploration Agency. To meet these design needs, we propose a new concept of solar array structure offering a high specific power and sufficient rigidity. To improve the rigidity, this structure uses flexible thin-film solar cells, which are a few micrometers thick, attached to panels with convex surfaces, and semi-cylindrical tubes attached to both ends of the panels, which are lightweight, can be rolled out/in, and stored. The rigidity of the structure was estimated via finite element analysis and compared with the experimental results. This new concept-based structure is expected to offer an improved specific power and sufficient rigidity over conventional systems.

On July 11th, 2019, the Hayabusa2 spacecraft achieved its second touchdown on the asteroid Ryugu with a 0.6 m accuracy. The landing was accomplished based on the newly implemented Pin-Point Touchdown (PPTD) method. It was extended from the former Touchdown (TD) method established by the Hayabusa spacecraft. When changing the point of view, two TDs of Hayabusa2 are regarded as successful autonomous rendezvous-dockings (RVDs) against the artificial landmark in deep space more than 15 minutes of light time from the Earth. This paper proposes the Deep Space Orbital Transfer Vehicle (DS-OTV) concept inspired and driven by Hayabusa2 technology to economically support various future deep space missions. The DS-OTV usually stays in a parking orbit at the boundary of the Earth system and possesses autonomous RVD functionality against the visiting deep space probes. The OTV provides the following services: (1) refueling of chemical bipropellant to the probe enabling it to transfer into the interplanetary trajectory by itself, (2) insertion of the visiting probe into the high-C3 Earth free-return orbit. This paper introduces the background and the concept of the OTV, and the simulation-based feasibility analysis results applying the Hayabusa2's PPTD method to the DS-OTV mission.

Ammonium diniramide (ADN)-based energetic ionic liquids (EILs) are promising new liquid rocket propellants for small satellites because of their high energy content and safety. This research focused on a mixture of ADN and hydrazide compounds. When simply mixed at room temperature, two solids' 1:1 (mass ratio), ADN (melting point = 92°C), and formic hydrazide (melting point = 56°C) formed a eutectic mixture. Additionally, they ignited at a lower temperature than other ADN-based EILs when a droplet was dropped into a heating furnace. To better understand the thermal decomposition mechanism in the condensed phase, the thermal behavior, and evolved gas during constant rate heating were examined using sealed cell differential scanning calorimetry (SC-DSC) and thermogravimetry-differential thermal analysis-mass spectrometry. Although in SC-DSC, the onset temperature of ADN/formic hydrazide's exothermic reaction was almost the same as that of pure ADN, heat quantity was higher than that of pure ADN. The evolved gas of ADN/formic hydrazide's decomposition seemed to contain a small amount of hydrazine. The ADN/formic hydrazide's low ignition temperature will be realized due to the reaction between hydrazine and nitrogen dioxide or nitric acid formed from the condensed phase reaction.

Small satellites have been used for remote sensing with a spatial resolution of several meters from LEO in recent years. However, it is difficult to increase temporal resolution for LEO remote sensing due to the short orbital period. Therefore, GEO remote sensing which enables observation of high temporal resolution from GEO or its nearby orbit is getting important. In order to obtain enough spatial resolution in GEO remote sensing, an optical system having a diameter of several meters is required because of the diffraction limit. It takes huge cost to realize such a large diameter primary mirror due to manufacturability and required accuracy. To address this problem, we propose a synthetic aperture telescope by small satellites formation flying. The synthetic aperture telescope is composed of several mirror satellites constituting a primary mirror of the telescope and an imaging satellite having a focal plane assembly. By optically synthesizing the light collected by each mirror satellite with the imaging satellite, a virtual large aperture telescope is constructed. In this paper, we showed numerical simulation results of images for urban areas and early-stage fires from GEO in the near-infrared band. As a result, it was shown that the fine structure of the ground could be observed. It was also shown that it is possible to identify the location in high-precision in the image of the initial fire assuming multiple point sources.

The extensible mast which is used as optical benches of space telescopes such as ASTRO-H with a focal length of over 10 meters requires high accuracy in shape. There are several factors that affect the shape accuracy of extensible mast, and deployment repeatability is one of the major factors. The deployment repeatability is affected by design parameters such as the size and physical properties of the extensible mast. In this study, the influence of each design parameter on the deployment repeatability was quantitatively evaluated by using two-dimensional analytical model of extensible mast. The results revealed the following. Deployment repeatability ㏌ the lateral direction of truss is largely influenced by the friction coefficient between the shaft and the hole, the hole diameter, axial stiffness of diagonal rod, and tension. Deployment repeatability ㏌ the longitudinal direction of truss is largely influenced by the friction coefficient, the hole diameter, and the gap betweenshaft and hole. Based on these results, it is expected that the deployment repeatability can be effectively improved by changing the values of parameters that have a large effect on the deployment repeatability. Therefore, this study has provided fundamental design guidelines for improving the deployment repeatability of two-dimensional extensible mast.

In recent, there has been a growing demand to utilize 6U class CubeSats for advanced missions, which require strict pointing accuracy for 6U class CubeSats. Although the past studies on large satellites suggest the importance of considering microvibrations to realize micro-radian order pointing accuracy, microvibrations effects in 6U class CubeSats have some uncertainties since only a few studies have been reported on their effects on the satellites of the similar size. This paper describes the evaluation process and results of the effects of microvibration on the pointing accuracy of 6U class CubeSat, using our bus system under development as an example. In the evaluation process, the disturbance from the reaction wheels (RWs) is measured to develop a numerical model. Then the pointing error induced from the RW's disturbance is evaluated with a numerical simulation. The evaluation result suggests that the degradation of the pointing accuracy due to the RW's disturbance is sub-micro-radian order if the satellite body is regarded as rigid-body, while it can increase to micro-radian order if the satellite body has structural resonance modes.

Heat transfer characteristics of liquid oxygen under high pressure and high heat flux should be elucidated to design a regenerative cooling nozzle for a hybrid rocket. Previously, we developed an experimental system using Joule heating to a thin tube in which liquid oxygen flows. The objective of this study is the evaluation of the measurement system to obtain the heat transfer characteristic under high pressure and high heat flux. We conducted a cold flow experiment and an experiment of heat transfer characteristic measurement. The friction factor of the tube is estimated with the Colebrook-White equation. The heat transfer coefficient was calculated by solving the heat conduction equation in the tube and the conservation laws of the fluid. The calculated heat transfer coefficient is in the same order as the experimental value. Consequently, we confirmed that experimental and analytical methods are appropriate to obtain heat transfer characteristics.

A composite plate consisting of ultrahigh-molecular-weight polyethylene (UHMwPE) was placed in front of aluminum alloy plates to reduce ejecta with the ultimate goal of reducing space debris. The temperature effects on crater size and ejecta generated from aluminum alloy plates with UHMwPE fiber composite were examined. When a UHMwPE fiber composite plate with a thickness of 1.0 mm was placed in front of an aluminum alloy 6061-T6 thick plate, the crater diameters and depths on the aluminum alloy plate clearly decreased, and the number of ejecta from the aluminum alloy plate with UHMwPE fiber composite clearly decreased. In addition, UHMwPE fiber composites with a thickness of 1.0 mm continuously reduced ejecta, even at high temperatures.

In this paper, we report an achievement of Human Resource Development Program for student of National Institute of Technology (KOSEN). This program is aimed to cultivate an ability of the aerospace engineering in non-space major students. The program consists of three parts of sub-programs: (1) KOSEN Space Academia which is a remote hands-on seminar using video conference system (2) KOSEN Space Camp which is a four-days camp-style workshop held in Niihama city, and (3) KOSEN-1 project which is an actual development of 2U size CubeSat. A total of 263 students and 78 faculty members participated in this program for three years from 2017 to 2019.

This paper proposes an algorithm to systematically search for a good sequence and the number of intermediate orbits in order to design the low-thrust minimum-fuel halo-to-halo transfer. The proposed algorithm applies successive convex optimization for the optimal transfer design, and initial guesses for the optimization are generated by beam search. In the beam search process, intermediate halo orbits that construct the initial guesses are chosen from a family of halo orbits so that they are located between the initial and final halo orbits. By determining the appropriate evaluation index for the intermediate orbit design, the proposed algorithm can collect local optimal solutions with a reasonable computational time. The obtained optimal transfers provide several options for the total △V and time of flight, which helps to find the suitable solution based on mission requirements.

Aiming for an improvement in transmission performance of 12-GHz-band satellite broadcasting, we are researching a digital pre-distortion (DPD) technique to compensate for distortion arising from the radio frequency (RF) component in the satellite transponder. We evaluated the performance of a digital pre-distortion technique on 32APSK for a 12-GHz-band satellite transponder, which has an analog pre-distortion (APD) linearized traveling wave tube amplifier (LTWTA).

The present paper addresses the development and demonstration of a numerical approach to study the characteristics of radio wave scattering by a solid object in space covered by a charged particle or plasma layer. The plasma conditions near the surface of the object are totally different from those of a pristine plasma in space, due to various types of ionization processes occurring at the plasma–body interface. This sometimes results in the formation of a dense plasma layer around the surface, which significantly alters the electromagnetic properties of the object. Such processes are relevant to a broad range of in-space radio applications such as radar observations of space debris, meteors, small dust grains, and re-entry objects. This paper proposes a new numerical approach of applying the particle-in-cell simulations to the scattering analysis. PIC is a well-established first-principles model that provides a kinetic description of a plasma by following the trajectories of an ensemble of charged particles in self-consistent electromagnetic fields. It is advantageous to include plasma kinetic effects as well as nonlinear behavior of the plasma in a straightforward manner. A number of numerical techniques such as improved treatments for conducting object surfaces and the field decomposition capability are implemented for the present study. The proposed method is applied to a simple situation of wave scattering from a conducting object covered by a dense electron layer. The simulation reveals an antenna-like behavior of the electron layer swayed by an incident wave, and its contribution to the differentiated wave scattering properties.