This paper describes a separation experiment conducted at extremely high spin rates for OMOTENASHI, being developed as the world's smallest Moon lander. The experiment validated the separation method used in the landing sequence and the tipoff impulse was estimated from the motion of the separated object. The study showed that the tipoff impulse increases superlinearly in the high spin rate region. This information is indispensable in determining the operational spin rate of OMOTENASHI in orbit. The method described in this paper can be applied widely to separation experiments for small spinning satellites and rocket stages.

The stability of spacecraft motion in the vicinity of a second body in the elliptic restricted three-body problem (ER3BP) is discussed. This problem is approximated as the Tschauner-Hempel problem perturbed by the gravitational force of the second body (Tschauner-Hempel three-body problem, TH3BP). Under the condition that the eccentricity is small enough, an orbital stability analysis is performed with the initial in-plane amplitude and out-of-plane amplitude of the spacecraft as two parameters. The system's unstable regions in the parameter plane are classified into three patterns: the first instability is caused by the average motion, the second is caused by the non-average motion, and the third is caused by the difference between the periods of the in-plane and out-of-plane motion. These mechanisms and the parameter conditions of orbit instability in each region are analyzed.

The rover for future Mars exploration will be expected to have higher rough terrain traveling performance compared with conventional rovers. In this paper, we propose a rover with a weight of about 30 kg that can be transported to Mars by a Japanese rocket. The purpose of this study is to propose and develop a hinge-type movable center-of-gravity mechanism wherein the bending of the rover's body with a motor shows high rough terrain-driving performance with minimum degrees of freedom. The proposed mechanism's performance is evaluated by conducting travel experiments on a assumed terrain, that is, a hard rock with a grade inclination of 45 deg and a height of 0.15 m above the sand. The effectiveness of the proposed mechanism in getting over a step was confirmed by conducting experiments using a prototype model.

The OKEANOS mission utilizing the Solar Power Sail is one of the candidates of the strategic middle-class space exploration to the outer Solar System lead by the JAXA. The mission is planning to be launched in 2030's, and rendezvous for spectral observations and landing for in-situ measurements of light isotopes and organic molecules to a D or P type Jupitar Trojan asteroid in 2040's. The flagship instrument on board is a high-resolution mass spectrometry (HRMS) system together with suits of remote-sensing instruments. Through in-depth scientific observations, the OKEANOS mission will provide critical input to the key questions of (1) constraining planet formation/migration theories, and (2) inventory and distribution of volatiles in the Solar System. We have conducted experimental tests of a sample canister sealing by a metal seal knife-edge, sample canister-MULTUM test, gas chromatograph-MULTUM coupling test for organics, and H and N isotopic measurements in the atmospheric air. Current performances are (1) high mass resolution was 30,000 at m/z = 20, (2) sample canister system with a knife-edge metal seal kept 90 % of a released gas in 1 hour for Cu or Au gaskets with/without regolith, (3) error meets the required precision and accuracy for nitrogen isotopic measurement but for hydrogen. These experiment tests need to continue for our scientific proposes on the asteroids.

Space Robotics Laboratory (SRL) of Tohoku University has been developing micro-satellite technologies and successfully demonstrating micro-satellite and CubeSat technologies in orbit in the past 10 years. The main computers of these satellites are all custom-made computers based on commercial off-the-shelf (COTS) components. SRL specifies on-board computers' specification and develops both software for the processors and the hardware logic for the FPGAs, supported by industrial manufacturing companies. The specification and radiation tolerance levels of these computers vary based on what types of functionalities they are responsible for, such as power control, attitude control, signal processing, etc. Through the experiences of ground development and orbit demonstration, SRL has accumulated design strategy of on-board computers for micro-satellites by utilizing a combination of FPGAs, and CPUs. These also include small scale microcomputers, System-on-Chip softcore-based computers, purely FPGA-based software-free computers, and very complex triple module redundant multi-processor computers. Functionalities of all these computers were evaluated in orbit. This paper will describe the design strategies applied for the past missions at the SRL, system reliability assessments of the designed computer architectures, and discuss about on-board computer design approaches for future micro space systems.

This paper shows the outline of an automated image processing module for the microsatellite Diwata-1 images. The downloaded images from Diwata-1 are automatically processed using thresholding and feature detection algorithms to determine the cloud cover percentage, land area percentage, and bodies of water percentage for each image. From these percentages, an image is classified as useable or not. The list of all useable images tagged by the automated image processing module is compiled and compared to the list of all usable images tagged by a manual operator. Results show that the automated process has an accuracy of 87% to 97% depending on the payload used for capturing the image.

In recent years, small satellite docking methods have been proposed for satellite missions involving multiple dockings and separations. Conventional docking systems utilize a robot arm or thrusters for close-range guidance in the final phase of a docking operation. However, because of the limitations of size and mass, these systems cannot be applied to small satellites. Therefore, in this study, we propose a novel guidance system that uses only simple electromagnetic devices, such as electromagnetic coils and permanent magnets. In particular, we propose a guidance method that uses only an electromagnetic force based on relative orbital dynamics. For close-range guidance, electromagnetic impulse force is generated to guide a chaser satellite onto a rendezvous trajectory toward a target satellite from a stationary orbit. To compensate for impulse force error and orbit disturbances, an active orbit controller based on the sliding mode control method is formulated to maintain the reference rendezvous trajectory. Based on our simulation results, we confirm that the proposed method can effectively realize close-range guidance for small satellite docking with relatively smaller power consumption. Thus, our proposed system can be effectively used for small satellites in missions involving multiple dockings.

Most hybrid rocket propellants comprise solid fuel and liquid oxidizer. The attractive characteristics of hybrid rockets are their high safety and controllable thrust. However, despite such benefits, hybrid rockets are not yet put into use because of their low regression rate and combustion efficiency. Therefore, many studies specifically examine improvement of the regression rate and combustion efficiency. Early efforts started in the 1960s; then both ammonium perchlorate and ammonium nitrate were employed as fuel additives. This study is conducted to evaluate ignition characteristics, theoretical calculation, and pyrolysis of a microcrystalline-wax based solid fuel adding ammonium nitrate as a solid oxidizer. We report whether the improvement can expect by the addition ammonium nitrate in this paper.

Asteroid Explorer Hayabusa2 has successfully touched-down on Ryugu and carried out the explosive-packed SCI (Small Carry-on Impactor) experiment in the 1st quarter of 2019. The spacecraft (S/C) is planned to be inserted into the return orbit from the asteroid at the end of 2019 and to return to the vicinity of the Earth at the end of 2020. After successive TCM (trajectory correction maneuvre) being carried out, the SRC (sample return capsule) is separated from the S/C and enters the earth atmosphere with a velocity of about 12 km/s. Passing through the severe aerodynamic heating corridor, SRC will jettison both the forebody and aftbody heatshields and deploy a parachute at an altitude of about 10 km. On deployment of the parachute, SRC will start to transmit a beacon signal, based on which the ground direction finding stations (DFS) will localize SRC. The present paper summarizes and reports the return and recovery operation of Hayabusa2 SRC, its instrumentation, and the onboard subsystems concerned with the return and recovery.

This paper describes the design and performance of 100 kW class Magneto Plasma Dynamics Thruster (MPDT) using water as a liquid propellant. In general, MPDTs have a large thrust-to-power ratio and high specific impulse among the electric propulsion devices and use ammonium and hydrogen as a propellant. However, hydrogen propellant requires a high-pressure tank and damages materials, and ammonium is reactive and toxic. Hence, we propose to use water as an MPDT propellant, focusing on the storability and non-toxicity. The use of water propellant will simplify the propulsion systems and reduce size and weight because water necessitates no cryogenic devices nor chemically resistant materials. A 100-kW class prototype was designed to show that water MPDT yielded 1-kA class discharge current, which is necessary for thrust production in self-magnetic field types.

Space technology has been grown stagnant because of the issue of the development cost, environment problem and any other problems. A reusable launch vehicle, called RLV, was developed in order to solve these problems. This paper treats the spaceplane that is one of the RLV that has a fixed-wing and three moving surfaces. We must make land the spaceplane safely because the spaceplane is reused as the next launch vehicle. Many constraints exist while its landing. This paper investigates the automatic landing system that takes the constraints into account by using the Dynamic Window Approach (DWA). The DWA was developed for running a wheeled vehicle in 1996. It has some problems when applying the DWA to the spaceplane. The first problem is increasing complexity because a spaceplane is represented by six-degree of freedom flight dynamics. The DWA can consider the restriction of the velocity, angular velocity, acceleration, and angular acceleration. However, a number of constraints should be considered for safety landing of the spaceplane. The proposed DWA makes land the spaceplane safely. The numerical simulation results demonstrate that the proposed method is useful for landing of the spaceplane by comparing with the conventional method.

A microgravity experiment is going to be held in 2022 by using sounding rocket to clarify a cool flame dynamics of multiple n-decane droplets such as ignition delay time, ignition location, cool flame propagation, and so on. To predict the cool flame dynamics near ignition limit, 2D axisymmetric unsteady numerical calculation was carried out. A single droplet, droplet pairs whose inter-droplet distances are 8 mm and 16 mm, and droplet array with inter-droplet distance of 8 mm are simulated. Ambient temperature, pressure, and initial droplet diameter are 550 K, 1.0 atm, and 1.0 mm respectively. To acquire the mixture reactivity for the identification of the cool flame ignition start point, 0D reaction calculation is also carried out. It is found that the highly reactive mixture distributes in the outside of the droplets where temperature is high and mixture fraction is low. Temperature rise caused by the cool flame generation starts from the outside of the droplet pair, and the cool flame surrounds the droplet pair with small inter-droplet distance. In the case of the droplet array, the cool flame is generated at the outside of the array. As tracing the highly reactive mixture region, it propagates inward with average propagation velocity of 7.3 cm/s, and the cool flame propagation velocity accelerated just before burned out.

This paper proposes a direct method for the minimization of the l¹ norm of the control inputs of the trajectory design in the three-body problem. This method is based on the sparse optimal control, which considers minimizing the thrust arcs. The minimum thrust arc appears as a result of the minimization of the l¹ norm in the formulation of the optimal trajectory problem. The sparse optimal control can be reduced to a convex optimization because the objective function and constraints represent convex sets by assuming linear dynamics. However, the equations of motion of the three-body problem are nonlinear and highly unstable so that it is difficult to find the optimal trajectory by a standard convex optimization problem. To satisfy the nonlinearity condition, this paper constructs the successive sparse optimal control based on the sparse optimal control. This method derives the l¹ minimum solution that satisfies the nonlinearity condition by solving the sparse optimal control iteratively. As an application, the transfer between Lyapunov orbits in the Sun-Earth system is solved and it is verified that the proposed method can perform even for the unstable dynamics. This example in the three-body problem also demonstrates the instability of the dynamics can be suppressed by efficiently utilizing the stable and unstable manifold of underlying dynamics. Then the solution is compared with the optimal solution obtained by nonlinear programming.

This paper describes a robust trajectory design and guidance method to GEO for all-electric propulsion satellites. The method splits the transfer time into multiple steps. The length of the steps is in multiples of the orbital period. Each step optimizes gains that generate steering directions of a thrust vector to minimize the weighted sum of residual errors between the orbital elements at the end of the step and those of the specified target orbit. The method extends the transfer with additional steps until the satellite reaches the target orbit. Guidance is based on this multistep optimization. Because the optimizer handles a few variables and no constraints, this method provides the quasi-optimal trajectory acquired with low computational cost. The method is robust to disturbances because it sequentially calculates the gains considering any disturbances. Numerical examples for the GEO orbit-raising transfer problem show that trajectory is appropriately generated to reach GEO by optimized gains against disturbances.

This paper investigates the orbit maintenance approach for the quasi-periodic orbit about libration points. Because of the highly unstable dynamical environment, a tiny deviation will lead spacecraft to be far away from the mission orbit. Floquet mode analysis incorporates the dynamical systems theory for the controller design. Unfortunately, Floquet mode method is difficult to be applied to a quasi-periodic orbit because the monodromy matrix is not well-defined so that the invariant manifolds cannot be determined to create the stationkeeping controller. To solve the problem, a new orbit maintenance strategy named covariance-based approach is proposed, integrated with an autonomous orbit determination. Based on the property that the maximum orbit uncertainty direction will converge to the direction of the unstable manifold of a local orbit, the controller is reformed by using the eigenstructure of the state covariance, which is isochronously computed by the orbit-determination process, to substitute for the monodromy matrix. Application to the orbit maintenance of a Lissajous orbit about Earth-Moon L₂ verifies the feasibility of our method and also investigates parameters that influence the overall stabilization and consumption cost.

Deep auto-encoder (DAE) is applied to the decomposition analysis on the transition process of intrinsic combustion oscillation. Two dimensional URANS simulation reproduced combustion oscillation in CVRC (continuously variable resonance combustor) including the dynamics of the stable-to-unstable and unstable-to-stable transition. The transition process has hysteresis while decreasing/increasing its oxidizer post length to change the stability. The time evolution of temperature, pressure, and heat release distributions are sampled during the hysteresis process. The DAE consists of an encoder network that outputs two feature variables and a decoder network that outputs the approximation of the input data set. The sampled time series of the distribution data are put into the DAE to train. The two variables characterize the state of the oscillation in a 2D feature space. The weight vector of the neurons in the last layer are trained so that the weighted sum of them approximately compose the simulated fields; and thus, the fields are decomposed into modes. By tracing the trajectory of the transition on the feature space during the hysteresis, it is found that the space successfully classifies the level of combustion instability. The decomposed modes are useful to identify the key phenomena of hysteresis process.

Research was conducted on such coupling characteristics as drag, lift, and pitching moment between the airframe and engine gas exhausts for a wide range of flight Mach numbers from an experimental aspect and a computational aspect. In case of no fuel injection, numerical results were in good agreement with the experimental results. Thus, the effects of fuel injection on the aerodynamic characteristics were numerically investigated, and a thrust margin was computed as a function of rocket conditions to achieve positive net thrust.

The world's first solar sail IKAROS (Interplanetary Kite-craft Accelerated by Radiation of the Sun) was launched along with the Venus Climate Orbiter Akatsuki in May 2010. At the time of writing, IKAROS is still the only solar sail in deep space. IKAROS was full successful of missions that have been planned for first half year. After the Venus flyby in Dec 2010, it carried out an extended mission. IKAROS has succeeded many additional engineering and scientific missions. On the orbit determination, we obtained various skills. IKAROS has been almost 8 years from launch, we have been still maintaining operation that repeating the power generation period and hibernation period cycle for 10 months and more. Even though we cannot be obtained enough observation data for orbit determination that can withstand long-term propagation over several years. But we have been successful in four times search operation, at the wake-up from the hibernation in 2012-2015. As a similar success, Rosetta has been successfully wake-up from hibernation for 31-month form June 2011 to January 2014 due to a power shortage. In that operation, the antenna pointing accuracy was not a problem at all. However, IKAROS has an area to mass ratio which is 50 times that of Rosetta, and the solar distance of IKAROS's orbit is only a fraction of Rosetta's one. So, it has been received several hundred times of Rosetta's solar radiation pressure. Therefore, we needed high precision attitude dynamic model for predicting future position of IKAROS. We have been 3 years from when we got receiving the last observation data in 2015. We studied extension of attitude motion model and optimizations for model parameter. The base of attitude motion model, was constructed using nominal phase of data and modified for the first wake up from hibernation. Then, the attitude motion model was divided 3 sections, by spin rate for search operation after 2015. And the attitude motion model parameters were estimated using one of the devised optimization method. In this paper, we present modified attitude motion model for solar angle-free and the optimized attitude motion model parameter.

Operational characteristics of an ablation sensor embedded into an ablative test specimen are examined under a time-varying heating profile using an oxy-hydrogen gas torch test set up. The time-varying heating profile is given to replicate temporal variation of surface temperature encountered in an atmospheric entry heating condition of a space vehicle. This study demonstrates the detected performance of the surface recession and char sensors for the case of the OREX flight condition. The result shows that the present ablation sensor can monitor ablation behaviors of the test specimen while detecting the difference of the ablating history in between time-dependent and time-independent heating profiles.

This paper describes the performance of a combustion-controllable solid-propellant microthruster using laser heating. Conventional solid propellant thrusters are relatively compact and reliable because the thrusters require neither tanks nor valves and never present the risk of propellant leakage during storage. On the other hand, interruption and restart of thrust production are difficult because the combustion is autonomously sustained once the propellant is ignited. Therefore, solid propellant thrusters have never been applied to orbit maintenance or attitude control. Hence, we have developed the solid propellants in which combustion is sustained only while the heat was supplied to the burning surface and proposed a throttleable solid propellant microthruster using semiconductor laser as a heat source. In our previous study, in a prototyped thruster, a laser head was moved by a linear traverser such that laser beam followed the regressing burning surface. The thruster interrupted and restarted thrust production by switching laser. However, the use of a linear traverser increased thruster weight. Hence, we redesigned a prototype, which maintained the gap between the burning surface and the laser-introducing window to prevent laser-beam attenuation due to the combustion products. Thrust measurement showed that the prototyped microthruster successfully yielded a thrust 0.06 N, I_sp of 70.3 s, and ignition delay of 0.9 s.

Solar sail IKAROS's membrane has unexpectedly deformed into a pyramid shape. Using finite element analysis, previous studies proposed the curvature of thin-film solar cells as a possible cause for this deformation and showed that a square sail consisting of four trapezoid petals tends to deform into a pyramid shape or a saddle one. The global deformation, however, was not verified experimentally. This study aims to shed light to the effect of the curved thin-film devices on the global deformation using a scaled model of the sail and to experimentally verify the finite element model of the scaled sail. Our results show that the simulation and experimental results for the out-of-plane deformation match for the saddle shapes, but those of the pyramid shapes are different. The disagreement between the simulation and experimental result is likely due to the adhesive layer. The difference in the simulation-experiment agreement between the saddle- and pyramid-shape cases is due to out-of-plane stiffness. It is suggested that the prediction of out-of-plane deformation in the saddle shape is more accurate than that in the pyramid shape because the adjacent petals contact each other and have high tensile stress in the circumferential direction in the saddle shape case.

IHI Aerospace previously developed a small (approximately 0.5 N) and low-cost green monopropellant thruster by using additive manufacturing (AM, 3D printing technologies). Based on this development research and small satellite market research, the development plan of an additively manufactured 4N-class thruster fabricated using flight-proven thruster materials was drafted. This paper reports the development status of an additively manufactured 4N thruster fabricated using Inconel® 625 and hydrazine as fuel. We adopted the binder jetting approach as the AM method instead of laser or electron beam welded methods to reduce the flow resistance, number of thruster parts, and welding process. A thruster firing test and a vibration test were conducted after a prototype thruster was assembled, and we confirmed that the thruster was well designed and possessed fundamental thruster performances.

We propose a new educational use of a nanosatellite for high school and junior high school students, and not making a nanosatellite but using it. We held a workshop in which high school and junior high school students made observation plans using Stars-AO, which is a CubeSat equipped with a high-sensitivity visible-light camera. We evaluated the educational effects of the workshop by analyzing the response writings of the students and their replies to questionnaires. The participants' abilities and interest in science and space engineering increased because of the workshop. In this study, we report the workshop details and evaluation of its educational effects.