Crystal growth of alloy semiconductor has been under investigation at the International Space Station (ISS) to investigate growth kinetics at the solid-liquid interface, because microgravity can suppress the natural convection. In this study, we focused on InGaSb which is one of the promising ternary alloy semiconductors. As a preliminary experiment, wetting angle between the InGaSb and the materials of the ampoule and cartridge, such as quartz, carbon sheet, BN, and C-103 alloy, were measured to check their affinity of the configuration. The InGaSb exhibited much higher wetting ability on the C-103 substrate than that on the other substrates (quartz, BN, and graphite), and this result suggested C-103 can prevent leakage of the InGaSb samples in case the ampoule collapses in the cartridge. On the other hand, BN and graphite exhibited low wetting abilities with InGaSb; therefore, they are suitable materials for the ampoule because they do not significantly affect the InGaSb crystal growth. In another preliminary experiment, concentration of Te dopant, which was added to make striation in the grown crystal, was measured using inductively coupled plasma mass spectrometry (ICP-MS). It is confirmed that Te concentrations were relatively higher at the striation area.

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We developed a measurement method of Soret coefficient S_T using a two-wavelength Mach-Zehnder interferometer. Temperature gradient was applied to pure molten salol or the salol - 6 mol% tert-butyl alcohol solution in a quartz cell. The direction of the temperature gradient was opposite to the gravity vector so as to suppress the convection. The movement of the interference fringes in solution side was evaluated by the number of the dark lines crossing a measuring point from right to left. The changes of temperature and concentration were measured using the number of the dark lines. Wavy noise found in measured number of the dark lines could be removed by subtracting one in the cell wall. The temperature change agreed with one measured using thermocouples. The results show that the measurement method in this study enables the contactless S_T measurement.

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In satellite communications using Wideband InterNetworking Engineering Test and Demonstration Satellite (WINDS), the throughput does not increase because of factors such as presence of a long delay environment and attenuation of signals by rain. One of the solutions for long delay environment is to enlarge the TCP window size. However, packet loss occurs when a signal is attenuated, and the throughput decreases with the retransmission time in TCP communications. Furthermore, the transmission efficiency worsens because of the increase in the number of retransmission packets when the TCP window size increases. In this paper, from the result of throughput measurement by changing TCP window size in the WINDS regenerative mode, the TCP window size which was suitable in purpose to maximize a throughput is presented.

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The ranging link could be a bottleneck and constrain a mission design because it suffers from a degradation of its two-way propagation. JAXA has introduced an onboard ranging signal regeneration scheme to solve this problem. It is new in the point that we do not apply correlation processes to a detection of ranging codes to be regenerated, and that a signal improvement by integration processes alone contributes to the regeneration. We call this technology as synchronous integration. We have developed both an onboard instrument and ground equipment in parallel. The preliminary test using a breadboard model of the X-band transponder and a prototype of our future ground system successfully proved that our regenerative ranging was effective. Based on this result, this new ranging scheme was adopted for our Venus exploration mission, PLANET-C. PLANET-C was launched in May 2010. During her transfer orbit to Venus, the comparison test between the regenerative and transparent ranging was conducted. This was the world first achievement of the regenerative ranging based on the synchronous integration scheme. We gave a detailed analysis to this experimental result. Our regenerative ranging performs better than our transparent one when a received uplink ranging signal to noise power density ratio is worse than 60 dBHz. At the ratio less than 40 dBHz, the regenerative ranging still keeps functioning while the transparent one needs an extended integration time for its operation. Compared with our current transparent ranging, our regenerative ranging has a 14-dB recovery gain for the uplink. It enables us five times as long a ranging link distance as before.

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This paper investigates the gains and losses in terms of power, area, reliability, and speed when applying time redundancy fault tolerance techniques on single core designs compared to space redundancy fault tolerance techniques applied to multi-core designs. The system is developed on the virtex5 FPGA from Xilinx, it uses 65nm technology with a relatively moderate to high static power consumption. The system consists of two design alternatives. The first is a single core embedded processing system that applies time redundancy fault tolerance through execution repetition to perform self-check pointing through consensus. The second system is built from 3 soft IP core processors which perform a space redundancy approach through Triple-Modular-Redundancy (TMR) with feedback among the processors. The performance of both systems is evaluated in terms of the execution speed and latency due to fault tolerance techniques compared to the non-fault tolerant system.

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Active Phased Array Antennas have attracted attention for use as hopping spot beam antennas that might be used as electronically controllable onboard satellite antennas. It is difficult to monitor the output of each amplifier at its output port directly, so the rotating element electric field vector method which has been already adopted as a means of determining the health of the transmitter and receiver modules in the onboard synthetic aperture radar is also effective for Active Phase Array Antennas. However, the difference between relative amplitude and phase values calculated using measurement data and those predicted values often becomes large due to the radio propagation characteristics change, signal level change in the satellite, satellite position and attitude determination errors and so on. This paper describes health check experiment of antenna elements on WINDS featuring two stations simultaneous measurement to improve the predicted relative values. It also describes that the phase difference of the RX antenna was greatly decreased, and the number of RX antenna elements whose phase difference is within 5.625 degrees was increased at two stations simultaneous measurement as compared with the results at one station measurement.

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This research proposes an X-band high efficiency onboard SSPA (solid-state power amplifier) for deep space missions by focusing on GaN (gallium nitride) HEMT (high electron mobility transistor) whose remarkable material properties, such as high thermal conductivity, wide band gap, and high breakdown voltage, are suitable for high power and high efficiency applications. Developing a high efficiency onboard SSPA is one of the great issues when we consider some missions toward Mars, Jupiter, and much farther planets because of the requirements of both ultra-long distance communication and low power consumption. As a first step toward developing a SSPA for deep space, a breadboard model is fabricated based on preliminary design. It consists of a buffer amplifier, a driver amplifier unit, a high power amplifier unit, an automatic level control unit, a variable attenuator, DC/DC convertors, and an over current protection unit. Here, GaN HEMT is used in both driver amplifier and high power amplifier units. RF (radio frequency) characteristics of these amplifier units are evaluated in experiments. The driver amplifier unit achieves output power of 31.5 dBm with power gain of 33.5 dB and less than -26 dBc of IM3 (third order intermodulation distortion) at P1dB (1dB compression point) at 8.425 GHz. Moreover, the maximum efficiency is up to 35.2%. On the other hand, the high power amplifier unit achieves 42.3 dBm of output power with 46.1% of PAE (power added efficiency) at P3dB (3dB compression point) at 8.40 GHz. In addition, the integrated GaN SSPA bread board model achieves the maximum output power of 41.9 dBm and the maximum total efficiency of 31.0% at 8.40 GHz. At least more than 5% total efficiency improvement can be seen compared to the previous onboard SSPAs. Moreover, space applicability of GaAs (gallium arsenide) MMIC (monolithic microwave integrated circuit), GaN HEMT and DC/DC convertor that are expected to be used in the SSPA are confirmed in total ionizing dose test.

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An aerial deployment technique is required for the Mars airplane to get a large wing area and compactness. A multibody dynamics simulation program was developed to assess the safety of the dynamic behaviour during aerial deployment. An effect of an initial deployment angle was revealed through the simulation. A pitching moment was generated due to the drag and pitching moment acting on the right and left wings.

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Low-density wind tunnel tests are carried out using a 1/5-scale model to evaluate the aerodynamic performance of a Mars airplane at low Reynolds numbers (Re = 10,000-33,000). Particularly, Reynolds number dependency on the aerodynamic characteristics and effects of the elevator are investigated. There is little change of the lift curve and the pitching moment in the Reynolds number range from 10,000 to 33,000. The effectiveness of the elevator is not enough for responsive pitch control. It is found that the low effectiveness of the elevator is due to a non-linearity of the lift curve of the horizontal-tail airfoil and a thinner tail-airfoil having a large lift-curve slope can improve the pitch control.

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The internal structure of the Moon is one of the most important clues to the evolution and origin of the Moon. A lunar gravimetry by differential VLBI observation, named dVLBI, is proposed for the SELENE-2 to improve the internal structure model, especially for the core. As a recent review of the dVLBI mission, two topics are shown. First, new operation modes are introduced for the purpose of saving the transmitter’s electric power. The power consumption is reduced from 9.6 W to 2.1 W in the case of the dual frequency and intermittent transmission modes. Error estimation of the lunar potential Love number k2 shows that the desired accuracy can be achieved in the case of the intermittent transmission mode. Second, the S-band survival antenna that will be installed on the lander is designed. A computational simulation using an analytic model shows that the required performances, which are gain of -5 dBi, beamwidth of 60 degrees, and bandwidth of 140 MHz, are achieved at a temperature within a range of -200 to +120 degrees.

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The objective of this study is to clarify the aerodynamic characteristics of the triangular shaped airfoil at very low Reynolds number corresponding to the Mars airplane flying in very low atmospheric density on Mars. Two-dimensional aerodynamic forces and pitching moments of the thin angular airfoils and the triangular airfoils with various thicknesses were studied through wind tunnel tests conducted in very low Reynolds number range of 3,000 - 20,000. Although the maximum lift coefficient and lift-to-drag ratio were large even for the triangular airfoil composed of a line pattern at the present Reynolds number, these aerodynamic components were affected in the airfoil thicknesses by changing the Reynolds number. The non-linear lift and pitching moment curves were characterized in low Reynolds number flows. It was found that the non-linearity of the aerodynamic curves was caused by the existence of the leading-edge separation bubble by measuring the pressure distributions of the airfoil at various Reynolds numbers.

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We are developing a silica-aerogel-based cosmic dust collector for use in the Tanpopo experiment to be conducted on the International Space Station. The mass production of simple two-layer hydrophobic aerogels was undertaken in a contamination-controlled environment, yielding more than 100 undamaged products. The collector, comprising an aerogel tile and holder panel, was designed to resist launch vibration and to conform to an exposure attachment. To this end, a box-framing aerogel with inner and outer densities of 0.01 and 0.03 g/cm³, respectively, was fabricated. The aerogel mounted in the panel passed random vibration tests at the levels of the acceptance and qualification tests for launch. It also withstood the pressure changes expected in the airlock on the International Space Station.

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The Active X-ray Spectrometer (AXS) for the Japanese SELENE-2 rover has been proposed for elemental analysis on the lunar surface to measure the major elements: Mg, Al, Si, Ca, Ti, and Fe; the minor elements, Na, K, P, S, Cl, Cr, and Mn and the trace element Ni, all depending on their concentrations at a landing site. The elemental data of the AXS allow us to not only classification but also quantification of surface rocks on the Moon. The AXS is a compact low-weight instrument for elemental analysis based on the principle of X-ray fluorescence spectrometry using an X-ray spectrometer and two (four) pyroelectric crystals as X-Ray Generators (XRG). This paper introduces the current status of the pre-project to develop an AXS for the SELENE-2 Rover including the investigations on the generation of X-ray flux of the XRG, required surface roughness for the XRS measurement, and a thermal design of the AXS.

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Japan Aerospace Exploration Agency (JAXA) are planning an outer solar system exploration and sample return mission from a Jovian Trojan asteroid using a 3000 m² solar power sail. The difficulty of this mission is a severe restriction on the weight; only 300 kg is allocated for sampling and returning to Earth. In this weight, fuel for trajectory and attitude control, sampling mechanism, re-entry capsule, and other systems required to return to the Earth are included. In this paper, a preliminary analysis of this sample return mission is conducted. Three scenarios for sampling are proposed; sampling with a 3000 m² solar power sail, with a detachable small solar power sail using electric propulsion systems and with a small probe using chemical propulsion systems. The mission analysis shows that the most feasible configuration is to conduct the sampling with the 3000 m² solar power sail using an extension mast.

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For providing further support toward manned space activity, robots are expected to conduct various and also far dexterous tasks than ever. Such robot, Astronaut Supporting Robot, has to handle different shape, size and mass at various works. One solution is to develop multiple numbers of modularized robotic end effector (EE) that is exchangeable so as to obtain dexterity and reliability at once. In order to utilize modularized EE, robotic arm and EE need to possess interface mechanism (wrist interface) which can connect EE structurally and electrically with robotic arm. In this paper, author presents Detachable Wrist Interface Mechanism (DWIM) as a wrist interface suitable for astronaut supporting robot. First, requests and requirements for wrist interface are clarified, and then the concept and design of DWIM is proposed. Based on the design, prototype model was fabricated, and the results from function verification tests are reported.

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The aim of this study is to find the optimal airfoil for Mars exploration aircraft, which requires high-lift-to-drag ratio. However, existing airfoils for flying in the Earth’s atmosphere do not have a high enough lift-to-drag ratio in Mars flight condition. The airfoil studied here was designed using a Genetic Algorithm (GA) and evaluated using two-dimensional Computational Fluid Dynamics (CFD) without turbulence model (laminar). The objectives in this optimization include the maximization of lift and minimization of drag coefficients at only angle of attack of 6 °. The Reynolds number is 2.3 × 10⁴ under the aircraft cruising condition. B-spline curves that connect neighboring control points express the upper and lower surfaces of the airfoil. The results show that some typical types of airfoils excel in aerodynamic performance. Most optimal airfoils have a large upper surface curvature or a strong curvature at the center of the lower surface. The former feature generates a separation bubble that leads to a high negative pressure, and the latter character makes a high positive pressure. Both phenomena generate lift force, and yield higher lift coefficient and high lift-to-drag ratio. Furthermore, most airfoils on the Pareto front have a thickness less than 10 % of the chord length, which is suitable for the wing structure design of the Mars aircraft.

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A propeller for small scale unmanned aerial vehicle (UAV) is used in low-Reynolds number conditions. In order to obtain high thrust and efficiency under low-Reynolds number conditions, the blade aspect ratio becomes small. The flow around the propeller with low-aspect ratio becomes quite three-dimensional. In the present study, experimental and numerical investigations are carried out to clarify effects of the three-dimensional flow on the propeller performance focusing on vortices associating with the low-Reynolds number flows.

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A new mission to investigate the Martian surface and interior structures using multiple penetrator probes is proposed. As a decelerator and a heat shield during the Martian atmospheric entry, a flare-type thin membrane aeroshell sustained by an inflatable torus frame will be applied to attain a compact and low-weight space vehicle. After separation from a carrier spacecraft, four identical penetrators will be deployed at intervals of a few hundreds kilometers from each other and penetrate into surface material up to a depth of 2 to 3 meters. Each penetrator installed seismometer system and a heat-flow probe will operate on the potential active regions in volcanism or seismic fault zones, associated with magmatic tectonics, crustal structure and current thermal state of Mars. A meteorological sensor package embedded in a backside of the aeroshell will observe the atmospheric structure during the entry, and it will possibly keep monitoring the Martian climate change on the surface. This paper describes the Martian penetrator design, the sequence of its deployment phase, onboard instruments, and their operational strategy.

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The realization of a sustainable manned Mars habitation mission cannot be achieved using conventional methods because the mission conditions, such as available resources, hazards, and journey time, differ significantly from those of previous manned space missions. Therefore, construction of a sustainable interplanetary transportation network, a Mars-based resource management system, and hazard management system should be considered. In addition, the development of key advanced technologies is very important. Hence, the logistics system constructed in this study considers not only mission requirements such as human transfer and stay times but also the available technology and required safety levels. Using simulations of the logistics system, we calculated the optimal stay cycle to minimize the initial mass in low Earth orbit. As a result, the effects of the limitation on the stay time and the superiority of stop-over and cycler trajectories with respect to the requirements of the stay cycle are quantitatively declared.

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A capsule-type vehicle is promising for the manned return vehicle from the Moon, Mars or from further bodies from the stand point of the weight budget regardless of its drawback of relatively low lift-to-drag (L/D) aerodynamic characteristics. The present paper describes the 6 degree of freedom (6DoF) simulation of the skip entry of the 4-ton class manned space return vehicle and discussed the issues related to the RCS bank control : the characteristics of the undesirable attitude motion due to leakage of the control force. Moreover, a fault tolerance algorithm based on extemded Kalman-filter (EKF) and the Sequential Monte Calro (SMC) has been introduced and applied to the state estimation in case that the roll rate signal and GPS signals are lost in the GPS/IMU integrated navigation system. The numerical simulation has showed that EKF+SMC fault tolerance algorithm successfully achieved state estimation of vehicle flight positioning by integrating IMU signal even under the condition that the roll rate and GPS signals are lost.

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Sound measurements were carried out using rocket-based propagation diagnostics in the middle and upper atmosphere via infrasonic/acoustic waves to investigate their frequency dependence. The S-310-41 sounding rocket was used, equipped with one main microphone, two sub-microphones, and a loudspeaker, to examine the sound propagation in the payload section of the rocket as a function of the ambient atmospheric pressure. The output from the loudspeaker showed a clear trend of gradual attenuation with decreasing atmospheric pressure at almost all frequencies. Acoustic wave propagation was measured, and was in good agreement with the predictions of attenuation of the sound strength of the mass spectrometer incoherent scatter model (NRLMSISE-00).

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The Normalized Difference Snow Index (NDSI) has aroused more and more interests as a snow detecting method, and its feasibility has been proven in many studies. However, the accuracy of NDSI alone has failed to meet the expected requirement. The detection of snow cover is not only affected by the atmosphere, but also influenced by the regional terrain and the underlying surface covered by snow. Due to this reason, in the present work, the snow detection in Akita Prefecture located in the Tohoku region of Japan was conducted through the combination of NDSI and NDVI (Normalized Difference Vegetation Index). Because of the large area of mountains and forests in the rugged Akita Prefecture region, the surface reflectance was retrieved from the top of vegetation after atmospheric and topographic corrections. Furthermore, in order to reduce the effect of the misclassification of snow and vegetation cover, a Normalized Difference Vegetation Index (NDVI) model was used to discriminate the snow and forest pixels. Compared to the MOD10_L2 data and NDSI alone results, the combination of NDSI and NDVI showed high accuracy in snow cover detection.

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In this paper, we evaluate the polarimetric sensitivity of multi-frequency synthetic aperture radars (SARs) for detecting landslide areas in forest-covered mountains. We tested L-band airborne and X-band spaceborne SARs, i.e. the airborne Polarimetric Interferometric SAR in L-band (Pi-SAR-L2), Terra SAR X, and Cosmo Skymed, at the Totsukawa-mura test site in Nara, Japan. We found that three parameters—the coherence of HH and VV, polarimetric entropy, and the power ratio of HH/HV—are very effective, especially with L-band SAR, for detecting land cover changes from a forest to a landslide. Results show that X-band SARs are less sensitive to landslide areas because the X-band penetrates less through a forest compared to the L-band.

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The Medium Earth Orbit (MEO) offers a greater width of satellite view than the Low Earth Orbit, and a greater proximity to the Earth’s surface than the Geostationary Earth Orbit. Global positioning systems and ionospheric sounding satellites have already operated in this orbit, and the MEO will appeal to satellite missions undertaking Earth observations, communications, positioning, and scientific explorations. The Sun-Synchronous Medium Earth Orbits (SS-MEOs) used in this study enable clear and daily observations of the Earth’s surface. However, problems such as severe radiation from the Van Allen radiation belt, and the inefficiency to launch a satellite in a southwest direction, need to be resolved. This paper presents the results of early studies on SS-MEO utilization: orbital characteristics and environments, candidates for Earth observation missions, and the feasibility of the development and launch of a satellite for operation within this orbit.

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Mangrove forests are one of the main forest types in Vietnam. However, during the last decades, a lot of mangrove forest area has been converted to other land-use types. The objectives of this study are to assess the changes that occurred in the mangrove forests from 1990 to 2010 and to determine a method for carbon-stock calculating for the forests that can be used as the input data for a Reference Emission Levels (REL) calculation. The study was conducted in two different research areas; one was the Quang Ninh province in the north of Vietnam, and the other took place in the Ca Mau and Kien Giang provinces in the south. In this research, the maps of mangrove forests for the periods of 1990, 2000, and 2010 were created by using optical satellite-imagery data. The change matrix was then calculated by overlaying maps of different periods in the forests for analyzing the changes and assessment. Field trips were also taken to each study area in order to verify/adjust the latest classified mangrove forest maps and interview/identify the main causes of land-use changes taking place during these periods. During the fieldwork, a number of sample points were surveyed to collect the number of parameters of mangrove trees. These data are then used for analyzing the relationship between tree height and the HH and HV values of ALOS PALSAR (Advanced Land Observing Satellite - Phased Array type L-band Synthetic Aperture Radar) data. An allometric relationship between biomass and tree height was also applied for investigating the relationship between aboveground biomass and ALOS PALSAR polarimetric data. Finally, the change in the carbon stock of the forests through all periods was estimated for living biomass, fuel-wood gathering, and other losses by applying the stock-change method. The results of this study were the mangrove forest maps and biomass maps for the years 1990, 2000, and 2010; the changes analysis showed that the forest area is decreasing and that the main cause is the conversion of the forests to aquaculture land, according to the equations/models of biomass estimation for biomass mapping based on ALOS PALSAR data.

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A wide variety of organic compounds have been found in space, and their relevance to the origin of life is discussed. Interplanetary dust particles (IDPs) are most promising carriers of extraterrestrial organic compounds, but presence of bioorganic compounds are controversial since they are so small and were collected in the terrestrial biosphere. In addition, IDPs are directly exposed to cosmic and solar radiation. Thus, it is important to evaluate the stability of organics in IDPs in space environment. We are planning a novel astrobiology mission named Tanpopo by utilizing the Exposed Facility of Japan Experimental Module (JEM/EF) of the International Space Station (ISS). Two types of experiments will be done: Capture experiments and exposure experiments. In the exposure experiments, organics and microbes will be exposed to the space environments to examine possible alteration of organic compounds and survivability of microbes. Selected targets for the exposure experiments of organic compounds are as follows: Amino acids (glycine and isovaline), their possible precursors (hydantoin and 5-ethyl-5-methyl hydantoin) and complex precursors “CAW” synthesized from a mixture of carbon monoxide, ammonia and water by proton irradiation. In addition to them, powder of the Murchison meteorite will be exposed to examine possible alteration of meteoritic organics in space. We will show the results of preparatory experiments on ground by using a UV lamp, a ⁶⁰Co source, synchrotron facilities, and a heavy ion irradiation facility.

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Heat shock protein 70 is known to inhibit muscle atrophy in in vitro and in vivo models. The goal of this study was to investigate whether a herbal compound celastrol (CEL) promoted HSP70 overexpression via activation of heat shock transcription factor 1 (HSF1) in C2C12 myotubes. The HSP70 expression was increased in a time-dependent manner with 5.8- to 9.1-fold upregulation at ≥ 4 h after 1.5 µM CEL treatment. Nuclear accumulation of phospho-HSF1 was evident at ≥ 2 h over the 6-h CEL treatment. Phosphorylation of Akt1 (the protein anabolic marker) was also increased 3.5- to 6.6- fold at ≥ 2 h while total Akt1 expression was not changed over the CEL treatment. These results imply that CEL has a potency to inhibit muscle atrophy via HSP70 induction and Akt1 activation. The results will be used for development of a protocol for a future ISS/JEM onboard experiment.

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Sub-millimeter debris impact is a cause of mission failure for satellites in low earth orbit. To protect a satellite, important components whose failure can cause critical damage to the satellite should be installed inside the satellite structure. However, this is impossible for some components, e.g., expandable structures and wire harnesses. To protect these components, a flexible debris bumper is proposed in this study. The ballistic limits of fabric bumper shields made from high-strength fibers were investigated through hypervelocity impact experiments. Three types of Kevlar cloth, two types of Beta cloth, and three types of Nextel cloth were tested, and their ballistic limit thicknesses and weights were calculated. The Beta cloth was the thinnest bumper, and the high modulus Kevlar cloth was the lightest bumper. Fabric bumper shields required more thickness than a conventional aluminum bumper. However, they can reduce weight by 30–60% compared to aluminum.

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A sample return mission to Mars has lately been proposed at Japan Aerospace Exploration Agency (JAXA). In the mission, we plan to capture floating Martian dust particles during the Martian atmospheric flight. In order to improve the feasibility of the Mars mission, we have carried out a study of capturing Martian dust particles using aerogel in the Martian atmospheric flight condition. At JAXA, a carbon aerogel-silica aerogel (CASA) two layered sample collector has been developed for the mission, and arcjet heating tests and light gas-gun capture simulations have been conducted to investigate characteristics of the sample collector. It was found that the superiority of the carbon aerogel against heating has been validated, and the sampling time can be longer than 10 seconds for the mission. Also, particles of montmorillonite, the primary constituent of Martian dusts, were successfully captured by the CASA sample collector in the LGG tests.

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As the electron emission yield induced by electron and photon plays a key role in surface potential of spacecraft materials, the ground based degradations including 500 keV electron and 50 keV proton irradiation with 4 different fluences were conducted for the polyimide film separately. Based on the developed measuring systems, the comparative measurements of total electron emission yield and photoelectron emission yield were carried out for the virgin and degraded polyimide samples respectively. The total electron emission yield and photoelectron emission yield tended to have different variation tendency after high energy electron and proton irradiation. The Monte-Carlo analysis software Casino and SRIM were used to analysis the distribution and stopping power of electron and proton respectively. According to the measurement results and analysis, the free radicals caused by irradiation was considered to be the main effect for polyimide films, which can primarily reveal the degradation mechanism of energetic electron and proton on the emission yield of polyimide.

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Developers of small satellites may think of enlarging their effective cross-sectional area to minimize the long-term presence in the low Earth orbit region. However, the enlargement of their effective cross-sectional area causes the temporal increase of cumulative probability of accidental collisions. This paper evaluates the influence of small satellites’ post-mission disposal by enlarging their cross-sectional area on the future population using an orbital debris environment evolutionary model. According to the result of the future projections, the small satellites’ post-mission disposals by enlarging the effective cross-sectional area have the potential risk of increasing the accidental collision. However small satellites are inserted into orbit below 800 km in this paper, the fragments to be generated by collision expected to decay within short term. The effective number of objects below 2,000 km altitude is not different between the case with post-mission disposals and the case without post-mission disposals.

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Unnecessary artificial objects, or so-called space debris, orbiting Earth contaminate the orbital environment and represent a serious problem for sustainable space development and use. One fundamental debris countermeasure is object tracking and cataloging. Such catalogs allow spacecraft mission risk assessment and the conducting of collision-avoidance maneuvers. However, the SpaceTrack object catalog released to the public is insufficient because the catalog includes relatively large objects (> 10 cm), only. We propose using two space-based sensors to address this problem and analyze their observational and tracking capabilities. Initial results suggest that approximately 4.2% of objects in low Earth orbit in size of 5-10 cm could be detected by space-based sensors. We also propose collaborative observations with a network of ground-based observatories. The space-based sensors are for detection and initial orbit determination, and the ground-based observatories are for catalog maintenance under collaboration with the space-based sensors. Initial orbit estimates from space-based sensors enable ground observatories to employ the image stacking method. Simulated differences between predicted and true apparent positions and motions indicate that it is possible to observe a target based on orbital data from space-based sensors.

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The number of electrostatic discharge (ESD) accidents on solar arrays has been increasing. Such accidents can cause serious problems for power generation, such as halting the normal operations of a satellite. To prevent ESD, ground checks should be performed under a normal potential gradient, i.e., when the satellite surface potential is negative with respect to space plasma. In this study, we obtained ESD parameters to establish a method for ground testing. Experiments were performed in a vacuum chamber with an electron gun. Discharge images (propagation length and velocity) were captured by an IR camera. The charge flowing into the discharge point was captured by a current probe. A non-contact surface potential probe measured 2D-potential distributions on test samples. As a result, we found that the threshold voltage of the electrical discharge was −10 kV and the amount of electric charge depended on the propagation area of the discharge plasma. The propagation velocity in a normal gradient potential was 10⁵ m/s.

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We have examined the possibilities of a ground-based optical observation system for monitoring LEO objects. Simulations and a test observation showed that two longitudinally separate observation sites with arrays of optical sensors can detect many LEO objects 30 cm in size and precisely determine their orbit. The proposed optical survey system may complement or replace the current radar observation system for LEO objects monitoring in the near future.

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In the recent times, there is a huge amount of data gathered about the Earth observation in case of disaster, not only from satellite observation but also from various data sources such as ground observation station, social network service (SNS), etc. One of such data collection techniques is Remote Sensing, which is also an essential part of interest for quality research works. There can be various study areas covered under the Space Technology such as Disaster, Astronomy, Remote Sensing, Geographic Information Systems, etc. In this case, data related to space technology alone may not be a sufficient resource every time. For example, the information collected from social data may be required to collaborate with space-based data/metadata to design an informative system. Therefore, there is a high necessity of data interoperability, which incorporates data from wide range of information along with space-based data, such as remote sensing and GPS. This scenario leads to challenges how to integrate geographic information of different forms from various domains along with the problem of heterogeneity. At the same time, the other challenges include standardizing spatial syntax and formalizing spatial semantics. With such excessive information, it is very difficult for the users to identify and extract the effective amount of information. Therefore, this study is an initiative to design a system specially focusing on disaster domain for accessing data from both space technology data sources and other available disaster related web portals (news, blogs, etc.) through text extraction processes implementing ontology for the effective information retrieval and visualization.

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Next generation arcjets should have light-weight design and prolonged lifetime. For the former topic, it is shown that the radiator mass can be drastically reduced by the effective use of propellant as a coolant at the lower temperature region on the radiator. Resulting thruster weight of 2.0 kg including the radiator is possible for 15 kWe arcjet. For the latter topic, replaceable cathode system is proposed and some key issues are mentioned.

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An air data sensor (ADS) system is one of the key components in space transports. It can provide precious information on the pressure and the Mach number data that cannot be obtained from internal motion sensors. When a spaceship or a reentry vehicle is cruising at a very high altitude, the ADS will suffer problem of the time-delay in its pressure signals that propagates through a thin pressure tube. In this study, the time delay of the pressure propagation inside a thin tube under a low-pressure environment is evaluated based on the direct-simulation Monte-Carlo (DSMC) method. A CFD analysis based on axisymmetric Navier-Stokes equations is also conducted to evaluate the contribution of rarefied gas effects to the pressure propagation speed. These results are compared with experimental results. Based on the DSMC and CFD analyses, the time constant for the pressure change is determined to be around 0.013 seconds whereas the experimental results yielded 0.12 seconds when the pressure at the exit of the tube increases for about 20 to 40 Pa starting from 30 Pa. These results suggest that the propagation delay observed in the flight test of the atmospheric-entry capsule descending at an altitude of 79 km with Mach-4.5 flight velocity was partly caused by the rarefied gas effect.

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The Space Systems Laboratory at Kyushu Institute of Technology is developing a fully reusable sounding rocket named WIRES (WInged REusable Sounding rocket). The winged rocket incorporates many novel technologies, including a full composite structure and a navigation, guidance, and control system. It is also equipped with an innovative hybrid rocket engine named CAMUI (CAscaded MUltistage Impinging-jet). In such a complex rocket, system integration is difficult to achieve and innovation is imperative. The laboratory is therefore also developing a subscale model of the rocket named WIRES#014 to assess the new navigation, guidance, and control system. This paper describes the procedure and results of a trial and error approach, comprising three ground combustion tests, to integrating the systems of the rocket. In the first and second try of the combustion test, the tests had some troubles mainly about ground support system and avionics. Authors eliminated these errors after the cause analysis; the third combustion test was finally succeeded.

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The thermal condition of ASTRO-H under air-cooled environment before launch is investigated with a thermal testing and a computational analysis. The thermal testing shows that the temperatures of devices are confirmed to be within the operating range if the additional fans are used. Moreover, the results of the thermal testing are compared with those of computational results. The computational results of temperature of the devices around the dewar with the additional fans are in good agreement with those of the thermal testing. The good agreement in the condition with the additional fans is because the forced convection, which is a dominant effect, is well captured in the computational analysis. Meanwhile, the computational results of temperature on the side panels are in very good agreement with thermal testing despite the difference in the flow outside satellite by air conditioner: computational analysis models the air flow from the air-conditioner while thermal testing does not. This is because the air-flow is very slow (0.1[m/s] at the side panel locations) and forced convection effects are very small.

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ASTRO-H is the sixth X-ray observation satellite in Japan developed by Institute of Space and Astronautical Science (ISAS/JAXA). ASTRO-H will investigate the physics of the high-energy universe by performing high-resolution and broadband observations. In order to achieve the mission objectives, ASTRO-H will carry six scientific instruments (two telescopes, two imagers, a detector, and a spectrometer) covering very wide energy range from 0.3 to 600 keV. This paper describes thermal design for the Soft Gamma-ray Detector (SGD) and the Hard X-ray Imager (HXI) installed on the ASTRO-H satellite. These instruments consist of silicon and cadmium-telluride semiconductor-based cameras, BGO active shields with avalanche photo diodes, and electronics. In order to achieve highly-sensitive observations, the semiconductor cameras and BGO active shields are required to be maintained at low temperatures for their high signal-to-noise ratios. This paper shows some measures and efforts to keep these devices at low temperatures under varieties of external heat input conditions and high heat dissipations from the semiconductor cameras and electronics. Thermal analyses using thermal mathematical model and some development tests using thermal test model of these instruments are reported as well.

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The authors have been developing CAMUI (Cascaded Multistage Impinging-jet) type hybrid rockets, explosive-flee small rocket motors. This is to downsize the scale of suborbital flight experiments on space related technology development. A key idea is a new fuel grain design to increase gasification rates of a solid fuel. By the new fuel grain design, the combustion gas repeatedly impinges on fuel surfaces to accelerate the heat transfer to the fuel. To demonstrate flight performance of a newly developed 5000 N thrust class motor and accumulate flight data around the sonic speed, a launch test was conducted from a coast to the sea. Basic technologies for the sea recovery are staged braking by parachutes, suspending the fuselage on the ocean, and locating the fuselage by an electric beacon and sea marker. Test results were successful and all of the fuselage was recovered. A typical drag coefficient profile around the sonic speed was obtained.

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New reaction initiation (ignition) system utilizing discharge plasma is proposed for 1N-class reaction control system (RCS) thrusters utilizing green monopropellant; especially for one of hydroxyl ammonium nitrate (HAN) based liquid monopropellants, called SHP163. This reaction initiation system was designed in substitution for conventional catalytic decomposition system. In this study, fundamental reaction initiation characteristics of propellant, based on total amount of energy and propellant mass reduction, were investigated. At the same energy level, higher mass reduction was observed at larger number of discharge attributed from frequency. Additionally, a laboratory model of the reaction initiation system was built, and fundamental characteristics based on power consumption were investigated experimentally. The new reaction initiation system was demonstrated without propellant flow. In order to understand effects of helium mass flow rate and frequency on power consumption and plasma generation characteristics, helium mass flow rate was varied from 3.2 to 27.8 mg/s, while frequency was varied from 100 to 5000 Hz. Power consumption was increased as frequency increased, and the highest power consumption, 34 W, was recorded at 5000 Hz of frequency at 12.9 mg/s of helium mass flow rate. In this study, stable plasma generation was observed at any frequency and helium mass flow rate.

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We have been testing di-methyl ether (DME) as liquid propellant for a coaxial type pulsed plasma thruster (PPT). Liquid-PPTs can set arbitrary amount of ejected propellant at arbitrary timing, avoiding “particulate-emission” and “late-time ablation,” whereas solid propellant PPTs cannot. Additional feeding pressurant would be removed with the help of DME's expedient vapor pressure. We conducted a study to investigate the effect of capacitance on the thrust performance with our prototyped DME-PPT. The result showed that impulse bits increased monotonically with capacitor-stored energy E. In higher range of E, impulse bits I_bit in the case with capacitance C=3 μF was lower than in the case with the other Cs. The maximum I_bit and specific impulse I_sp were 102 μNs and 575 s, respectively, at E=13 J (C=6 μF) and a mass shot Δm=18.4 μg in the range conducted in this study C=3–9 μF.

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Measurements of plasma parameters and thrust imparted from the compact helicon source having an annular permanent magnet in addition to the previously reported magnet configuration are performed, where two different diameter annular permanent magnets of the 5.4-cm-diam and the 8-cm-diam are tested. The source diameter, the working argon gas pressure, and the 13.56 MHz rf power are chosen as 6.6 cm, 0.85 mTorr, and 500-2000 W, respectively. Both the plasma density in the magnetic nozzle and the measured thrust for the 8-cm-diam annular magnet are higher than that for the 5.4-cm-diam one. The thrust can be increased by the effective rf power, and reaches a maximum of about 9 mN and 6 mN for the 8-cm-diam and the 5.4-cm-diam annular magnet, respectively. As some of the plasma flow is trapped by the closed field lines near the annular magnets, the obtained thrust is lower than the configuration with no annular magnet.

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Magneto Plasma Sail (MPS) is spacecraft propulsion that produces an artificial magnetosphere to block solar wind particles, and thus impart momentum to accelerate a spacecraft. In the present study, we conducted three-dimensional particle-in-cell simulations on small-scale magnetospheres (< ion Larmor radius, 100 km) to investigate thrust characteristics of MPS, in which the magnetosphere is inflated by an additional plasma injection. As a result, we revealed that the finite thrust generation and the increase in thrust are obtained on the small-scale magnetosphere even if the electron kinetics is taken into consideration. The thrust of MPS (1.0 mN) becomes up to 14 times larger than that of the original magnetic sail (0.07 mN). However, it was also revealed that the thrust gain of MPS defined as thrust of MPS / (thrust of magnetic sail + thrust of plasma jet) is approximately unity and thrust-mass ratio and thrust-power ratio are further smaller compared with other existing propulsion system. An extensive improvement of a thrust is required for the realization of MPS.

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A miniature ion propulsion system is currently developed by the University of Tokyo collaborating with the Next Generation Space Technology Research Association (NESTRA) in Japan. This miniature propulsion system is intended for the installation on 50 kg small spacecraft "HODOYOSHI-4" developed by NESTRA under the Japanese government funded project, "New Paradigm of Space Development and Utilization by Nano-satellite". The engineering models of all the components were developed and assembled as an ion propulsion system. A number of operations of the system were conducted using a vacuum chamber. The present specifications of the engineering model was evaluated as the weight of 7.9 kg (dry:6.9 kg), volume of 39 cm×26 cm×15 cm, power consumption of 39 W, and thrust of 300 μN with specific impulse of 1200s.

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To investigate mechanisms of high-specific impulse characteristics, or high speed plasma generation, of a laser-assisted pulsed plasma thruster, speeds of ions of plasma plumes exhausted from the thruster were measured with the time-of-flight (TOF) measurement. In the measurement, temporal variations of ion currents exhausted from the thruster were measured with a Faraday cup. From the results of single Faraday cup measurement, placed at 400 mm away from the thruster, the maximum speed of ions of 169 km/sec at charge voltage of 2000 V (or charge energy of 14 J) was obtained. On the other hand, with double Faraday cup measurement, installed at 400 mm and 600 mm, the maximum ion speed was 220 km/sec at charge voltage of 2000 V.

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Laser absorption spectroscopy (LAS) is a useful technique for sensitive quantitative measurement of plasma parameters. We used LAS to obtain plasma profiles inside the μ1 miniature ion thruster developed by the Institute of Space and Astronautical Science of the Japan Aerospace Exploration Agency; this thruster is intended for installation on small spacecraft weighing up to 50 kg. It operates at low microwave power (1 W) and has been developed as a complete system, including a neutralizer. At present, the thrust efficiency of the μ1 thruster is not as high as that of standard ion thrusters and optimization is required. A detailed profile of plasma inside the discharge chamber is necessary for further improving performance. We developed an experimental LAS setup with spatial resolution of 0.1 mm for a modified μ1 thruster model. The number density distribution of neutral particles in a metastable state was measured and found to be about 10¹⁷ m^-3.

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A laser-ignition micro solid rocket for application as propulsion on a 1-10 kg-class microspacecraft was proposed and designed. The laser-ignition micro solid rocket uses boron/potassium nitrate as propellant, and a number of solid propellant blocks are ignited by irradiating their surface with a 1 W diode laser through a laser-incident window. According to previous experiments, contamination of the laser-incident window can cause ignition failure to as high as 39% (7 thrusters out of 18) as a result of the gas generated at the initial phase of the laser irradiation. To reduce this contamination, the distance between propellant and laser-incident window is the parameter of interest in this study. The ignition probability was measured using a newly designed propellant stand with variable distance between propellant and laser-incident window, and an optimized window-to-propellant distance was applied to the thrusters. As a result, the ignition probability reached 100% (26 thrusters out of 26) while maintaining the specific impulse.

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The speed and number of ions in a laser-electrostatic hybrid thruster were experimentally estimated using a Faraday cup for various electrode configurations to determine the optimum electrode geometry. The results showed that, for a 10-mm-diameter electrode placed in front of a laser ablation target, the speed of ions increased with target voltages. The ion speeds, measured 80 mm away from the target, were 16 km/s and 32 km/s for target voltages of 0 V and 500 V, respectively, for a target–electrode gap of 20 mm.

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