Currently, various microgravity environments are being developed to study new materials and the analysis of physical phenomena. However, existing microgravity generation methods have two typical problems. One is high cost and the other is that experiments cannot be conducted anywhere. Therefore, a system is proposed that solves the aforementioned problems. In the proposed method, the system is dropped from high altitude and accelerated by ducted fan so as to compensate the influence of air resistance force and attitude control is simultaneously provided using mounted actuators. In this study, a planar model is introduced for basic study of the control system. The planar model is formulated by multibody dynamics and control torque is derived for stabilizing the attitude of the system applying the equation obtained and Lyapunov stability theory. In addition, the effectiveness of the model and the control system are verified by conducting numerical analyses and experiments.

Owing to certain limitations of wind tunnel experiments, it is difficult to obtain the real flight flow field of a vehicle entering the Martian atmosphere using wind tunnel data. Consequently, it is necessary to establish a correlation between simulation of the entry of a vehicle into the Martian atmosphere using a wind tunnel and the actual entry of a vehicle into the Martian atmosphere. In this investigation, a numerical method and a comparative analysis are utilized to explore the application of data from wind tunnel experiments to the actual entry of vehicles with representative geometries (excluding the Mars Pathfinder) into the Martian atmosphere. The results of the comparative analysis demonstrate that, if the same double-scale parameters are used for existing reentry vehicles with various geometries, the pressure coefficient and dimensionless pressure near the stagnation point of a low-enthalpy wind tunnel (using air with a composition similar to that of Earth's atmosphere) can be applied to directly extrapolate the experimental parameters of the wind tunnel to the actual aircraft flow field parameters of the Martian atmosphere. In the case of high-enthalpy wind tunnels using air with a composition similar to that of Earth’s atmosphere and those using CO₂, the pressure coefficient and dimensionless pressure near the stagnation point can be used as the extrapolation parameters; however, the Stanton number and dimensionless heat flow cannot be directly used as extrapolation parameters.

Hybrid rockets have the technical problem of a low fuel regression rate, which thereby causes a low fuel mass flow rate. Owing to recent developments in additive manufacturing technologies, using a solid fuel grain with a complicated geometrical port is one method for improving the flow rate. Star fractal geometry was employed as a complicated geometry. Combustion experiments using the grain with a star fractal port were performed. A conventional circular port was also tested for comparison. Local fuel regression rate, axial-direction-averaged local fuel regression rate, thrust, specific impulse and c* efficiency were evaluated. The local regression rate was high near the injector as a result of the impingement of injected oxidizer flow on the surface. While the axial-direction-averaged local regression rate of the star fractal port was slightly less than or almost comparable to that of the circular port, the thrust of the star fractal port was higher than that of the circular port because of the higher fuel mass flow rate. In addition, there was little difference in the specific impulse and c* efficiency when comparing the star fractal and circular ports. Therefore, star fractal geometry is superior to the circular port as the port geometry for hybrid rocket fuel grain.

Flash LiDAR is a relative navigation sensor that irradiates a target with pulsed laser light and then detects the reflected light to measure intensity, direction, and distance as a time of flight. We are developing “Flash LiDAR” for a space rendezvous. It employs a 3D image sensor that has a multi-pixel photon counter with photon-counting sensitivity and a readout IC with sub-nanosecond time of flight resolution. We conducted a feasibility study with a breadboard model to confirm it could satisfy functional and performance requirements of the HTV-X berthing mission, under limited configurations entailing a short range, a narrow field of view, and no relative position/attitude rate. We then obtained excellent performance results that satisfy the requirements of the HTV-X berthing mission. In this paper, we describe an overview and the features of our Flash LiDAR, the results of preliminary design and breadboard model tests, and the applicability to future space exploration missions.

Thailand's agriculture is economically remarkable but most of agriculturists and farmers, on the contrary, have significantly low profit. Nowadays, technology is used to solve many agricultural problems, especially analyze area to find the most suitable crops and area for efficient growth rate of crops. Agricultural Exploration Assistant Satellite or AEASat is the CanSat prototype that participated in Thailand CanSat Competition 2018 and received The First Place Winner and The Best Presentation Awards. The goals of AEASat are to study environmental growth factors (temperature, relative humidity, average rainfall, carbon dioxide intensity, red and blue light intensity) of six Thai economic crops (rice, cassava, maize, sugarcane, rubber tree, palm) and find the most suitable area for many agricultural actions. The first mission of AEASat takes place in Chai Badan, Lop Buri Province, Thailand. Electronic structure of AEASat is categorized into four floors respect to their functions: sensors floor, main controller floor, data logger and wireless communication floor, and power management floor. Another part of the mission is the ground unit. The ground unit also collects environmental factors including light intensity which AEASat does not collect. The last part of the mission is ground station. The function of the ground station is receiving data from AEASat with wireless 488-MHz-frequency LoRa module. The ground station is also used to track AEASat from deployment to a touchdown. Qualitative data and quantitative data are analyzed with different methods. The qualitative data is evaluated to find the growing suitability in the area by using Geographic Information System (GIS) software. Aerial photograph, a part of qualitative data, is analyzed by visual interpretation. The quantitative data is analyzed by calculating average of the correlation coefficients of six factors by using MATLAB. It is calculated in order to find the most suitable crop out of six to grow in the area. As a result, sugarcane is the most suitable crop to grow in Chai Badan area. Agriculturists and farmers can have access to website which the results, detail, and description of the exploration are uploaded. In conclusion, if agriculturists and farmers grow the crops as advised, they can expect more products quantity.

Asteroid Explorer Hayabusa2 arrived at asteroid Ryugu in June of 2018. During its stay around Ryugu, Hayabusa2 will undertake several challenging missions such as collecting soil samples and deploying rovers, a lander, and an impactor. Before the spacecraft attempts these missions, our initial and most important work is to reveal the asteroid’s characteristics and properties. Science equipment is used to obtain data for use in making an asteroid model and for the landing site selection (LSS) of sample soil collection and rover/lander deployment. The LSS requires surface images taken from an altitude of around 5 km. Middle altitude observation was the operation needed to meet this requirement. This paper presents an overview of the operation, design, and study of hovering at a 5-km altitude from the surface, and specifically describes the RHC hovering method and Monte Carlo analysis of hovering trajectory and dV. It also describes the actual operation results.

Deployable structures facilitate the reduction of the stowed volume in CubeSat. In this investigation, an initially deflected panel called a “convex-panel” is proposed. The convex-panel generates a self-deployment force during deployment and has a higher bending stiffness after deployment with a reduced thickness. This results in a reduction in the stowed volume of the deployable flat-panel structures. To demonstrate the feasibility of the convex-panel, deployment experiments were conducted with conceptual models. Based on these experiments, higher self-deployment velocities were observed in the beginning and the last deployment stages, and complete deployment was achieved via appropriate arrangements of the convex-panels. These results demonstrated the feasibility of the convex-panel for the self-deployable structures used in CubeSat.

This paper presents the design methodology and results of a low-energy lunar transfer trajectory with specific landing conditions. Future space exploration missions require a precise landing technique that employs autonomous navigation based on terrain pictures obtained by an onboard camera to estimate relative states to the target landing point. Picture-based autonomous navigation requires an appropriate sunlight condition during the landing sequence, and thus determines the relationship between the landing orbital plane and sun direction. To achieve low-energy transfer, solar tidal force is utilized for perigee raising, thereby making sun direction during coasting prior to lunar orbit insertion (LOI) a key parameter for low-energy transfer trajectory design. This study employs the bi-circular restricted four-body problem to model the spacecraft dynamics. The spacecraft trajectory is solved by the combination of forward propagation and backward propagation. Forward trajectory begins from trans lunar injection (TLI), and is propagated for the specified time period along with a lunar flyby. Backward trajectory begins from LOI and targets the end point of the forward trajectory. The required delta-v for the deep space maneuver is calculated so as to match the end states of both trajectories in velocity. This study identifies three types of transfer trajectory characterized by lunar flyby conditions and Jacobi integral. When the optimized result is compared to a basic Hohmann transfer trajectory, the total delta-v from TLI to LOI is reduced by 124.8 m/s, 175.9 m/s, and 158.6 m/s, but the time of flight increases by 65.2 days, 82.3 days, and 64.6 days, respectively. All types of trajectory are feasible in terms of mission design to meet the required landing conditions, and effectively reduce the required total delta-v with an acceptable increase in the total time of flight.

ITF-2 is the second CubeSat developed by University of Tsukuba YUI project. It took over ITF-1's missions (building of a global YUI network, demonstrative experiment of an ultra-small antenna originally developed and a new microcontroller) and its system was designed based on ITF-1. The operation of ITF-2 started right after its release from the Japanese Experiment Module named Kibo attached to the International Space Station on January 16, 2017 and continued until its atmospheric re-entry on January 5, 2019. The Tsukuba ground station conducted more than 1000 operations over one and a half year and acquired considerable data. In addition, we received over 2100 signal reception reports from over 20 countries. In this study, we analyze the data accumulated throughout the operation and evaluate the missions.

Solar neutron observations are very important for understanding of nucleon acceleration mechanism in solar flares, but there are only a few tens of detection since the discovery in 1980. This is because solar neutron observations have been mainly carried out from not space but the ground with insufficient sensitivity. For microsatellite applications, we have designed very compact and high sensitive solar neutron and gamma-ray spectrometer utilizing a novel photo-sensor Silicon photo-multiplier. This paper describes concept, design and performance of our detector for micro/nanosatellite applications.

ALE project is a micro-satellite project with the mission objective of artificial generation of a shooting star. The first satellite of this project (ALE-1) is launched in January of 2019 at an orbit of approximately 500 km of altitude. There are manned missions in the International Space Station at an altitude lower than this, therefore the altitude of ALE-1 needs to be lowered before the actual mission start. There exist mechanical systems that allow a passive lowering of the altitude of satellites by accelerating the natural decay of the orbit with use of a thin polyimide film as a drag sail. Previously such systems were only used to de-orbit completely and decrease space-debris. We propose and design a mechanical system which allows operation of the satellite during de-orbiting, and separation of the drag sail when the desired altitude is reached. We consider sufficient power generation during de-orbiting; passive stabilization of the satellite attitude for shorter de-orbiting time; and mechanical safety and reliability of the system. This paper summarizes the design, development, and ground verification of the proposed module SDOM (Separable De-Orbit Mechanism) along with a projected orbital decay of ALE-1 satellite.

The microsatellite ALE-1 has been jointly developed by ALE Co., Ltd. and the Space Robotics Laboratory (SRL) of Tohoku University in order to demonstrate a novel artificial meteor-generating payload. This mission necessitated the adaptation of SRL's flight heritage microsatellite communications system architecture to follow the Recommended Standards of the Consultative Committee for Space Data Systems (CCSDS). This paper describes and evaluates the new ALE-1 communications system, which implements the CCSDS protocols while maintaining the previous functionality and making minimal modifications to the legacy design. A simplified CCSDS implementation was developed around the integration of a commercial-off-the-shelf (COTS) CCSDS controller into the existing system. The performance of this adapted and simplified implementation is reduced as the maximum real-time data rate is limited and significant overhead is introduced. However, the changes to the size, weight, and power (SWaP) of the system are small, thereby achieving the design objective of making minimal system modifications. This system was tested thoroughly on ground and is currently operating on orbit, facilitating critical operations for the ALE-1 mission and demonstrating the feasibility of CCSDS-compliance on microsatellites and the adoption of the CCSDS standards on legacy platforms.

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

“Two-dimensional grating and two-camera method” has been validated to measure areal relative displacements of curved surface of structural shape. A parabolic antenna of 1.5m in diameter has been measured by only one shot of photograph for the purpose of correction of reflector error, in order to develop a high accurate antenna system. The effectiveness of the proposed measurement method for a large and precise curved surface has been verified.

This paper considers formation flying control of multi-spacecraft in the presence of stochastic uncertainties. While relative position keeping is an important requirement in formation flying control, there practically exist various uncertainties such as solar radiation pressure, atmospheric drag and sensor and communication noises. Therefore, we suppose a stochastic disturbance in the relative motion dynamics, and we model the dynamics as a stochastic port-Hamiltonian system. Specifically, this paper considers stochastic bounded stability as a stability concept. Stochastic bounded stability guarantees that for given bounded region and achieving probability, the sample paths starting from the initial region remain within the bounded region with a probability more than the assigned achieving probability. Finally, the simulation results demonstrate the validity of the proposed method.

Epsilon-4, which was the Enhanced Epsilon launch vehicle optional configuration with a multi-satellite mount structure, was successfully launched with seven satellites in January 2019. The seven satellites consist of 200 kg-class small-satellite, three micro-satellites and three CubeSats in the Innovative Satellite Technology Demonstration-1 program. The post flight analysis was conducted. The chamber pressure and thrust of three main motors and chamber pressure of SMSJ and SPM were analyzed to confirm the effectiveness of design and production methodologies. All solid propulsion systems for Epsilon-4 showed a very good behavior during the flight. As a result, the validity of the design and operation of the Enhanced Epsilon launch vehicle was confirmed.

At Japan Aerospace Exploration Agency (JAXA), the Super Low Altitude Test Satellite (SLATS), named “TSUBAME” was successfully launched in 2017. In this work, we investigate rarefied aerodynamic characteristics for the SLATS geometry by carrying out free molecular (FM) and direct simulation Monte Carlo (DSMC) computations, and SLATS aerodynamic databases are developed on the basis of FM and DSMC results. The databases are used to estimate atmospheric density for SLATS, and we finally discuss the SLATS aerodynamic characteristics by comparing preliminary SLATS density data with several conventional atmospheric models.

A flight-ground test comparison program is undergoing at JAXA to reveal so-called facility effects on hypersonic aerodynamics and combustion phenomena. The flight vehicle will mount a combustor duct along its centerline, with so-called alligator type inlet with side-spillage to attain good starting characteristics. The incoming flow condition was estimated with 3D-CFD, and one-dimensional chemical kinetic calculation was conducted to settle design guideline of combustor which could enhance so-called vitiation effects on combustion. The design guideline was further evaluated by 3D-CFD with fine chemical reaction model, showing possible difference in combustion efficiency and resulting pressure level.

This paper presents the concept of the Ring Enclosing Docking System for Spherical Shaped Free-flying Robot. This docking system captures the robot with caging-based capturing and form closure technology. A tendon actuation mechanism with an elastic element is selected to realize light-weighted, faster actuation and contact force reduction. The validity of the proposed docking system’s acceptable range for capturing and the method of contact force reduction are experimentally evaluated.

In this study, the semi-optimal attitude steering logic for low-thrust orbit raising considering both the power generation and the angular momentum constraints is analytically derived to minimize the transfer time. In the proposed method, the attitude steering is realized by the rotation around the sun direction and the successive rotation around the rotation axis of solar array driving mechanism (SADM). These two rotation angles are determined so that the relative attitude error with respect to the ideal attitude is minimized in the presence of the angular momentum and torque constraints. The numerical simulations show that the degradation of the transfer time by the proposed logic is only less than several percent compared to the case without the angular momentum and torque constraints.

This paper describes the performance of a proposed premixer-type bipropellant thruster using nitrous oxide (N₂O) and dimethyl ether (DME). Conventionally, nitrogen tetroxide and hydrazine are used in thrusters for spacecraft. However, they require gas treatment systems in the ground tests owing to toxicity and reactivity to materials. Hence, we proposed to apply a premixer to N₂O/DME bipropellant thruster to develop a compact and eco-friendly thruster. The bipropellant is neither toxic nor reactive to materials, and require no gas treatment systems in the ground tests. Moreover, N₂O and DME, which are liquefied gases, are stored in liquid form, and are readily mixed with a premixer. A propellant supply system is simplified because they have sufficient vapor pressures for supply using self-pressurization. In this study, we prototyped a 0.4-N class N₂O/DME premixer-type prototype bipropellant thruster using a coil-type premixer with three types of injector. The thruster yields stable combustion with C* efficiency of 81.8 % at O/F=3.5, at which the theoretical specific impulse is the maximum.

A highly sensitive laser absorption spectroscopy system was constructed to measure the atomic oxygen number density using a forbidden line at OI 636 nm. First, to detect the small transition probability line, integrated cavity output spectroscopy (ICOS) was developed. The sensitivity of the ICOS system was estimated to be enhanced 822 times based on the ring downtime measurement. Based on microwave plasma diagnostics, the number density was measured to be as low as 1.2×10²¹ m^-3. Next, wavelength modulation spectroscopy (WMS) was combined with ICOS to increase the sensitivity further. For the optimized modulation frequency, sweep frequency, and time constant, the signal-to-noise ratio of WM-ICOS was increased 26 times compared to that of ICOS. Thus, the results indicate that the WM-ICOS system enables the measurement of atomic oxygen number density at a partial pressure exceeding 250 Pa.