The impact of the AC voltage frequency and the distance between the exposed electrodes (gap length) on the thrust characteristics of the trielectrode (TED) plasma actuator have been investigated experimentally. The induced body force of the serrated TED plasma actuator is investigated on quiescent air. With the experimental results, the physical factors that affect the growth of thrust is considered by estimating the total momentum of ions from the drift velocity. As a result, we found that the characteristics of the TED plasma actuator can be predicted from two physical parameters, the rate of the AC voltage change and the drift velocity of ions, which are both affected by the gap length and the AC voltage frequency.

Effect of the electromagnetic flow control system on the controllability of the aerodynamic characteristics of the reentry vehicles was investigated by 3-dimensional CFD analyses. Numerical results showed that the applied magnetic field increased the drag force and decreased the lift force at higher angle-of-attack. This suggests that the Lift-to-Drag ratio can be controlled efficiently by the applied magnetic field. Additionally, numerical results showed that as the aspect ratio increases the effect of the applied magnetic field on the controllability of the Lift-to-Drag ratio becomes weak.

The Small BAseline Subset (SBAS) technique allows the line-of-sight displacement image and the analysis of space-time characteristics using the differential interferometric Synthetic Aperture Radar (InSAR) algorithm. In this study, we used InSAR processing software GMSTAR to apply the SBAS approach to eighteen ALOS/PALSAR datasets in order to evaluate spatial surface deformation on the Pacific side of Chiba and to explore the accuracy of the method. To proceed SBAS analysis on this study, twenty-nine InSAR pairs from the eighteen ALOS/PALSAR datasets were accordingly selected. In this SBAS analysis, the alignment of all SAR images must have subpixel accuracy in the radar coordinates of azimuth and range in order to stack all interferograms derived from each InSAR pair. Also, interferograms derived from each original InSAR pair were reviewed to improve the initial SBAS analysis. The phase unwrapping software Statistical-Cost, Network-Flow Algorithm for Phase Unwrapping (SNAPHU) was used to unwrap phase map images, which are needed to produce the line-of-sight displacement image. Subsequently, six InSAR pairs having appropriate phase map images from original InSAR pair were selected to retry the SBAS analysis. Our results indicate that SBAS analysis can be a useful approach to evaluate surface deformation at a mm/year scale. In addition, the subsidence of seashore area in Chiba Prefecture could be evaluated visually by changes in the SBAS analysis.

Debris mitigation measures such as post-mission disposal (PMD) were set by considering the effects thereof using a debris evolutionary model. Many small satellites have recently been deployed in orbit and various plans are proposed for a so-called mega-constellation consisting of thousands of satellites in Low Earth Orbit (LEO). New systems such as an electric propulsion system and air drag augmentation devices are also proposed for satellite de-orbit. Therefore, the impact of them must be discussed in order to preserve the space environment. This study evaluates the debris mitigation measures taken by using the Near-Earth Orbital Debris Environment Evolutionary Model (NEODEEM) that was jointly developed by Kyushu University and JAXA. It revealed that mega-constellations have much impacts on the debris environment, and that the future environment will be affected by the PMD compliance rate and how PMD is achieved.

In 2010, a bare electro-dynamic tape-tether technology experiment (T-Rex) was carried out on board the sounding rocket S520-25th in order to demonstrate tape tether deployment and the orbital motion limit (OML) theorem. In this mission, the tape tether was stored in and deployed from a storage container by ejecting a daughter component using compressed springs. The designed deployment length of the tape tether was 300 m. However, analysis of the flight data revealed that only 130 m was deployed, which was much less than expected. The main reason for incomplete deployment of the tape tether may be friction between the tape tether and the storage container. In order to design an appropriate storage container and select the spring constant, it is very important to estimate the total energy dispersion of the deployment speed due to friction between the tether and the storage container. In this study, the tape tether is modeled as a multi-particle system, and collisions and friction between the tether and the storage container are considered in estimating the kinetic energy dispersion. The important parameters, such as the coefficients of restitution and friction, are chosen by referring to web data and experimental results, so that we can emulate the behavior of tape tether deployment from the storage container.

This paper presents the radiation testing results on several non-space-ready commercial-off-the-shelf (COTS) devices to validate the space readiness of a lunar rover flight model. These COTS candidates are bombarded by a proton beam at a high energy level, simulating the lunar surface radiation conditions under solar flares. We have determined that the tested devices experience a minimal number of minor single events, when each hardware receives beyond the actual mission dosage. Furthermore, the system recovery from a single event is confirmed throughout the test. The result culminates a step further to the qualification of a launch-ready vehicle.

This paper describes the methodology used for the implementation of a time-of-flight (ToF) camera in a lunar exploration microrover as a hazard detection and avoidance (HDA) sensor. Each frame from the ToF camera provides a 3D point cloud of the environment. Such data have unique advantages for autonomous hazard detection. First, we investigate the durability of the ToF camera system to survive the launch and space environment and report the results of qualification tests: a vibration test, thermal vacuum test, and radiation test. Second, we focus on the critical optical performance under lunar surface illumination conditions. We present the rationale and implementation of the first improvement made to the commercial version of the sensor to better suit strong illumination conditions. We conducted optical testing to verify that the quality of the 3D point cloud is sufficiently reliable for input into the HDA algorithm. Finally, we implemented a proof-of-concept algorithm and performed field testing of the integrated detection, analysis and command chain. Through this research, a terrestrial sensor is qualified and customized to provide the data quality capable of autonomous hazard detection and avoidance on the lunar surface.

In this paper, a fuel optimal rendezvous problem is tackled in the Hill-Clohessy-Wiltshire framework with several operational constraints as bounds on the thrust, non linear non convex and disjunctive operational constraints (on-off profile of the thrusters, minimum elapsed time between two consecutive firings...). An indirect method and a decomposition technique have already been combined in order to solve this kind of optimal control problem with such constraints. Due to a great number of parameters to tune, satisfactory results are hard to obtain and are sensitive to the initial condition. Assuming that no singular arc exists, it can be shown that the optimal control exhibits a bang-bang structure whose optimal switching times are to be found. Noticing that a system with a bang-bang control profile can be considered as two subsystems switching from one with control on to with control off, and vice-versa, a technique coming from the switching systems theory is used in order to optimise the switching times.

The landing gear for planetary exploration is required to achieve a secure touchdown on rough and inclined terrains for future missions. The conventional methods such as a honeycomb crash or an airbag were employed in previous missions. These methods have the disadvantages of no reusability and no proactive function to prevent tipping over. To solve these problems, a Linear-Rotary-Energy-Conversion Mechanism (LRECM) has proposed. It could prevent a spacecraft lander from tipping over through the counter torque generated by conversion of the kinetic energy of the spacecraft into the rotational energy of the reaction wheels. The effectiveness of this mechanism has already confirmed through the two-dimensional experiment. However, an actual landing has three degrees of freedom of motion and the performance of the proposed mechanism has to be confirmed through the three-dimensional experiment to apply the real world problem. Therefore, we extend the two-dimensional LRECM into a three-dimensional LRECM (3D-LRECM), and its effectiveness is validated by performing landing experiments of a prototype lander with several angles against slope gradient conditions. The experimental results show 3D-LRECM has satisfactory effectiveness and fault tolerance against failure of a reaction wheel.

This study addresses the control method of the azimuth direction of a stratospheric balloon gondola with simple and low-cost system. Many azimuthal control methods of the gondola have been developed mainly to achieve accurate pointing of astronomical observation instruments. As progress of technology, such systems have become complicated and expensive. On the other hand, as variety of balloon mission increasing, demands for azimuthal control system with rough accuracy, simple and small system, and low cost are also rising recently. This study proposes an azimuth control method with only a motored decoupler. The motored decoupler is an attitude control device to twist and untwist the suspension not to transfer torsion torque of the suspension and the rotation of the balloon envelope. However, the mechanism can also be used to generate torsion torque for azimuthal control of the gondola. This study aims to achieve the azimuthal control capability with the motored decoupler. The control law with observable parameters was derived by linear quadrant regulator (LQR) method. The control law was evaluated by some numerical simulations. As the result, the proposed method is concluded useful for actual balloon missions.

To achieve the wireless power feeding through the metal barrier without penetration of the wall by the wire feed through, it is necessary to avoid the eddy current generated on the metal surface. This paper presents a proposal for a pseudo-resonator method based on magnetic resonance coupling that uses the metallic structure to resemble the power relay resonator. To demonstrate the principle of the pseudo-resonator method, the four coils wireless power transfer system via magnetic resonance coupling was used in the 0.1 and 5.0-mm-thick stainless steel hollow tube. Results show that the efficiency of power transfer for the pseudo-resonator method was constant at 83% and 64% with the axial distance between the TX and RX resonator. Results revealed that the pseudo-resonator method has an advantage in wireless power transfer through the metal structure.

Construction of a high tower to the stratosphere was first proposed by Canadian Thoth Company. Because of its enormous scale, people are hesitant to implement the project immediately. This paper proposes several next steps to its realization. Fundamental aspects of the structure are presented. Based on the models, treated are a long bridge over Tsugaru Straight, medium height towers for an inter-city rope way or a suspended rocket launcher in addition to the stratospheric tower. For a light weight structure, some inflatable models are tested to verify its effectiveness to the aerospace field of application.

In recent years, because development of space technology has been increasing for the purpose of improving social infrastructure, the expansion of space transportation system based on low-cost and high-frequency rockets is important. Due to the compactness, inexpensiveness, and easy-handling properties of solid propellants used in solid-fuel rockets, numerous studies on solid propellants have been conducted. However, solid propellants are highly viscous slurries and highly explosive. As there is no device capable of continuously and safely transporting the solid propellant, the process of manufacturing the solid propellant is a batch process. We focused on the movement of human intestines that knead and transport with a small force, as part of the development process. In this paper, we developed a peristaltic pump, Mk. III, for kneading a solid propellant. The pump was comprised of a heating system, an input device for the powder and fluid, and a rapid exhaust valve. An investigation into the amount of input of the raw materials was undertaken, and the tendency of kneading at the point of introduction of the powder and highly viscous fluid was determined.

This paper describes a solid propellant microthruster that is throttleable through laser heating. In general, solid propellant thrusters are relatively compact and reliable because the thruster requires neither tanks nor valves, and never induces propellant leakage. However, the start and interrupt of combustion is difficult because 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, wherein combustion is sustained only while external heat was supplied to burning surface, and proposed a throttleable solid propellant microthruster using semiconductor lasers as a heat source. Our previous study showed that a prototype thruster successfully started and interrupted thrust production using near-infrared 45-W laser. However, the prototype yielded ignition delay of 3-5 s. In this study, to reduce the ignition delay, carbon black (C) diameter was reduced from 50 to 10 μm such that laser beams are absorbed in a shorter depth of the propellant. Thrust measurement showed that prototype thruster yielded an ignition delay of 1.6 s, a stable thrust of 0.06 N, I_sp of 127 s at laser power density of 0.93 W/mm² for φ10-μm C.

The electric propulsion system for CubeSat is required to achieve various missions. A miniature water ion thruster for CubeSats has been proposed and tested. To investigate the effect of the dissociation of water on its performance, a direct thrust measurement is required. In this study, a 100 μN-class thrust stand for 10 kg units was designed and integrated. It was calibrated using weights whose mass was measured in advance and achieved the measurement of the thrust of the miniature ion thruster using water. The range of measured thrust was 60–140 μN, and the average of its uncertainty was ±7.6 %. The thrust coefficients were around 0.9, similar to or slightly less than that of xenon. This shows that the dissociation of water has little effect on the thrust.

The reduction of the spacecraft's (SC) asymptotic velocity and the radiation hazard are really main problems for low-Delta V cost Jovian moons missions: orbiters and landers. Algorithm to overcome the "obstruction of solo disturbances" for one-body flybys around some Jovian moon with using full ephemeris with two coupled CR3BP engaging has been implemented. The region where the total received radiation dose (TID) exceeds is skirted along the upper section of Tisserand-Poincare graph. Withal low-cost reduction of the SC asymptotic velocity is required for rendezvous with small body. It became possible to find such scenarios when restricted three body problem is transformed into the two-coupled CR3BP models and full ephemeris model. Advanced Multi-Tisserand coordinates has been exploit for parametric passage into this region. With their help it is shown that the "cross" gravity assists at the early stage of reduction of the orbital period are required. As a result, a reasonable increase in the duration of the mission can be exchanged on a sharp decline TID and found "comfortable" (in TID) rounds scenario in the system (less than 70 krad for standard SC protection 8-10 mm Al). This will provide significant gains in the payload for spacecraft missions in Jovian system and systems of other outer planets and improving the reliability of their scientific instruments.

Disturbances in space environment around the Earth (Geospace disturbances), which are driven by the solar activities, are causes of spacecraft anomalies, radiation hazards of astronauts and aircrews on polar route, problem of navigations and HF communications, and induction current in long-line power cables. To mitigate the risk of geospace disturbances, improvement of operational space weather forecast with high precision based on numerical forecast scheme would be important. However, the poor observation points in geospace prevent us to introduce numerical forecast scheme with data assimilation technique which was common in terrestrial weather forecast. To breakthrough this situation, we are planning to develop a space environment sensor package, which can contribute to safety operation of the micro satellite itself, and to realize low-cost global monitoring of space environment based on constellation of micro satellites. Space weather observations by several tens of micro satellite constellation enables us to narrow down the sensor performance of single satellite. The optimization strategy of sensor performance for further miniaturization and power saving can accelerate mountability of the package. Based on this approach, it is expected that numerical forecast scheme with data assimilation technique will be introduced for operational space weather forecast in the future.

This paper describes the conceptual study of a dust sensor with large sensitive area exploiting multi-layer insulation (MLI). The sensor targets dust particles having a size of 10 μm or more for detection and a hypervelocity possible to penetrate the outermost layer of MLI made of polyimide film (typically about 20 μm in thickness). The sensor of the dust instrument consists of the outermost layer of the MLI and piezoelectric (PE) elements attached on the layer. The PE elements detect elastic stress waves formed by impact pressure generated by a dust particle penetration to the layer. The electronics section shall have a function of ADC (analog-to-digital) sampling that records the signal waveforms output from the PE elements of the sensor section. In this paper, we will describe the study results of the initial concept study of the dust instrument.

The Japan Aerospace Exploration Agency is now planning a Jupiter Trojan asteroid exploration mission using a 40 m-wide solar power sail spacecraft. In this mission, a 100 kg lander is soft-landed on an asteroid in order to perform in-situ analysis of asteroid samples by a mass spectrometer. In this mission, not only surface sampling, but also sub-surface sampling up to a depth of 1 m, will be performed. In the present paper, a sampling scenario including sub-surface sampling is proposed that does not require an anchoring system and that omits contamination from the analyzed sample. In this scenario, newly developed sampling instruments are implemented. In addition, ground experiments have been performed using these sampling instruments, surface sampling instruments, and sub-surface sampling instruments. The results of these experiments indicate that the sample mass required for in-situ analysis can be collected and that the proposed scenario is realistic and practical.