Cosmic gamma-ray observation provides much information on high energy objects and phenomena in space. Sub-GeV/GeV gamma-ray astronomy has been developed by projects such as EGRET, AGILE, and Fermi; however, many problems have yet to be solved. Not only high statistics observation but also high quality observation are required to advance gamma-ray astronomy. Therefore, we promote the GRAINE project, of which the aim is precise and polarization observations with an emulsion gamma-ray telescope. The balloon flight was performed in Australia during 2015. We aimed for the detection of the Vela pulsar and verification of the overall performance of the telescope. In this paper, we report the analysis status of the balloon flight during 2015.

Typically, it is important for high-altitude balloons for scientific observations to fly for a long-duration to obtain sufficient data. A super pressure balloon facilitates long-duration flight by using many ropes between the poles of the balloon to decrease the local curvature radius of the balloon film and thereby increasing the pressure capacity. However, the increase in the number of the ropes also increases the fabrication cost of the balloon. Hence, this study proposed a method in which a balloon was covered with a diamond-shaped net. The development of the new super-pressure balloon commenced in 2010 with ground inflation tests, numerical simulations, and flight demonstrations in which the size of the balloon was gradually enlarged. The goal of the study included flying a 1,000 kg payload at an altitude exceeding 37 km for a period of several months by using a 300,000 m³ balloon. Thus, this study introduces the development of a new super-pressure balloon.

NIRS measurement was carried out in a confinement experiment for the first time. A tendency for NIRS integrated values to decrease was observed. Problems such as the learning effect needed to be resolved. SOC had an upward trend. Regarding POMS, the first day of confinement showed the highest score, after which no obvious deterioration was seen.

On-orbit service technology is important to develop future orbital operations such as refueling, fixing a broken satellite, and replacing an old part with an upgraded one. Basically, the chaser has to capture an orbital target in each orbital operation. However, a docking interface consisting of the docking mechanism and a target marker for on-orbit service has yet to be standardized. This paper focuses on the target marker, especially a circle-type of retro-reflective marker (CR-Marker). First, the marker requirements are clarified by conducting a survey and the trade-off between existing visual markers is described. Based on these requirements, the proposed marker is designed and produced experimentally. The estimation accuracy of the proposed marker is evaluated by on-ground experiments.

Lately, space developments have accelerated day by day, but space is still an unfamiliar territory for the general public. Therefore, a better method that stimulates a strong interest in people is needed. In response, a new program that combines an aerial shooting with a virtual reality (VR) technique has been carried out by Nagoya University's student team, Nagoya University Aerospace and Flight Technologies (NAFT). A VR form aerial video is created by launching and retrieving a space balloon equipped with two wide-angle cameras. After editing, the VR aerial movie is exhibited at various events, and has a strong impact on audiences. The events confirm the validity of the VR movie as a method of space education.

Spray combustion is widely used in gas turbines, diesel engines, jet engines, boiler and oil-fired heaters. It is indispensable for life and industry. In order to improve understanding of the flame-spread mechanism in the fuel sprays, many researchers have conducted microgravity experiments on the flame-spread of fuel-droplet arrays and droplet-cloud elements. The effect of ambient pressure on the flame-spread phenomena has also been investigated. This research involves microgravity experiments being conducted on the flame spread over droplet-cloud elements at high-pressure condition. The droplet-cloud element consists of three droplets, Droplets B, A and L. Droplets B and A are interactive droplets. Droplet L is the droplet to investigate the characteristics of flame spread from Droplet A. The results suggest that interactive combustion between Droplets B and A increases the flame-spread-limit distance from Droplet A to Droplet L at 200 kPa. The 1200 K position of the thermal layer around Droplet A also increases with the interactive combustion. This research also conducted the flame spread to Droplet L in the perpendicular direction, y-direction, to the direction from Droplet B to Droplet A, x direction. The results suggest that interactive combustion between Droplets B and A increases the y-direction flame-spread-limit distance to Droplet L more than the x-direction flame-spread-limit distance.

In the operation of a solar power sail IKAROS, curved thin-film solar cells were assumed to make the membrane deformed. The deformation caused propellant consumption to counteract windmill torque. For a next solar power sail in the Japan Aerospace Exploration Agency, we conducted experiments and finite element simulations of a simplified model of a part of the solar power sail to understand the effect of a small curved thin-film on wrinkles in a thin membrane under uniaxial tension force. We compared shapes of the membrane with the curved thin-film solar cell under varying tension force. Main findings are that wrinkle lines near both lateral sides of a small curved thin film are bended by curvature of the film and occur under varying tension force.

In this study, a reusable educational motor that can burn different types of materials in the chamber was developed. Specifically, five candidate sweets were selected as fuel and burnt. The combustion performances of these sweets were compared, and the soft candy was selected as the solid fuel to be used in the motor, because it had the highest total impulse. The SOUKI Systems Co. Ltd. designed a small candy hybrid rocket based on these results. The candy rocket was launched successfully using soft candy at the Kada Cosmo Park. The motor carried the candy rocket to 319 m altitude and was fully functional under the high-acceleration environment.

The Martian Moons eXplorer, the objective of which is to obtain a sample from a Martian moon, was presented by the Institute of Space and Astronautical Science and the Japan Aerospace Exploration Agency. For its trajectory design, some configurations of a propulsion system were considered using a comparison of the fuel efficiency and total mission time. The use of chemical and electric propulsion for its return trajectory was introduced in order to realize a fast and low-energy escape from Mars. In this configuration, the spacecraft departs from Mars using three-body dynamics with chemical propulsion and using electric propulsion for a v-infinity leveraging maneuver (VILM). In this paper, we first propose the modified all-three-body method, which constructs the escape and VILM trajectories using three-body dynamics to obtain the chemical–electric hybrid-propulsion return trajectory. We then propose the integrated three-and-two-body method, which uses three-body dynamics at the edge of the sphere of influence and integrates it into two-body dynamics in order to use a two-body-based low-thrust optimizer. A comparison of the integrated three-and-two-body methods and modified all-three-body method is performed. Furthermore, the patched conics method is used to compare the optimized solution obtained using the two-body and three-body dynamics.

In recent years, a number of sample return mission and planetary exploration probes have been discussed and proposed. Our group has developed a new atmospheric re-entry vehicle with a membrane aeroshell to increase the variety of these missions. However, there are still several important technical problems to be addressed to apply the membrane aeroshell to an actual mission. One of them is evaluation of the thermal durability of the inflatable structure. The thermal durability of the inflatable structure was evaluated using a 10 kW class inductively coupled plasma (ICP) heater. This ICP heater can produce a plasma flow with a high enthalpy and relatively low heat flux of about 120 kW/m², which is a suitable condition for the heating test of the membrane aeroshell. The tests proved that the inflatable structure, made of polyimide film, silicon rubber adhesive, ZYLON textile, and alumina felt, maintains the gas tight in the plasma flow with a heat flux of 120 kW/m² in 300 s. This layering structure is proposed as a potential candidate for use in actual flight vehicles.

Quantum Key Distribution utilises fundamental quantum properties and mechanics to generate a cryptographic key, ensuring the highest security regardless of advances to computer processing. AMMEQ-1 is a 3U cubesat being developed by The University of Sydney, Australia hosting a primary payload provided by Centre for Quantum Technologies, Singapore. Its main purpose is to show USYD's capability of developing a satellite using COTS products and to demonstrate various technologies to support future QKD missions. This study describes the mission and overall system design of the cubesat.

The School of Aerospace, Mechanical and Mechatronics Engineering (AMME) at The University of Sydney is proposing the design of a commercially viable 100 kg class microsatellite, HISA - a Hyperspectral Imaging Satellite for Australia. The primary mission is centred around the application of hyperspectral imagery to address compelling agricultural and environmental problems specific to Australia at a low cost. This study presents the technical mission requirements for maximum utility and data throughput. It also presents a description of system design and the subsystems which have been carefully selected after discussions with Australian Government departments.

We have worked to develop new optical surface shape measuring methods enable to grasp surface shape of large space structures with high precision and high speed on orbit for future space antenna and telescope. These methods are applicable to testing such structures on ground. In these methods, we analyze phase values of projected or painted grating patterns on the structures and perform calibration using a reference plane. It is hard to project such patterns on large structural surface on orbit, however, we must paint some grating patterns on the structures. In that case, high contrast images of grating patterns, for example white and black grating painting, are needed for precise measurement. Nevertheless, high contrast white and black patterns on surface make thermo-optical features of the structure more complex, then there is a possibility of interfering with thermal design of spacecraft. Therefore, to widen the application range of our method, we propose surface shape measuring method based on grating patterns using ultra-violet range. We use two different kinds of painting materials and cameras having sensitivity to light of ultra-violet range. Both painting materials are photographed as white in the visible range, however, one is white and one is black in the ultra-violet range. In this method, we can get high contrast grating images on the surface only in the ultra-violet range. In this study, we provide some feasibility study using commercially available ultra-violet cameras and painting materials such as titanium oxide.

The discharge process in a dielectric-barrier-discharge (DBD) plasma actuator was numerically investigated in order to understand the EHD force production process and seek the key factors to enhance the performance for flow control applications. In this study, a sinusoidal, triangle, and square voltage waveforms were applied with the same amplitude and frequency. The streamer discharges were repetitively observed in the positive-going phase in a voltage cycle, whereas the diffusive structures are obtained in the negative-going phase as reported in previous experimental studies for the three waveforms. The time-averaged EHD force for the sinusoidal waveform case was the smallest in our simulation condition, and the forces for the triangle and square waveform cases were almost the same value. However, the distribution of the EHD force for the square waveform was quite different from the triangle waveform because of the discharge regime difference. For the square waveform case, a lot of streamers propagate in a short time in the positive-going phase, and no discharge is obtained in the plateau phase. The applied voltage waveform has impacts on not only the time-averaged EHD force but also the EHD force distribution. These differences will also affect on the induced-flow structure and the control performance of the DBD plasma actuator.

Assessment of the worst plasma environment for spacecraft surface charging in the geosynchronous Earth orbits (GEO) is important for spacecraft designs and operations, because it could cause spacecraft anomalies due to surface charging with resultant discharging arcs. The differential charging potential of spacecraft surfaces is considered to generate the harmful discharging arcs. There was no common standard of the worst GEO plasma environment for analysis of differential surface charging to prevent and mitigate the anomalies. Therefore in order to determine a new International Organization for Standardization (ISO) document for the purpose, a round-robin simulation was performed using the NASCAP-2k and MUSCAT. We perform surface charging simulations by Spacecraft Plasma Interaction Software (SPIS) as same as the round-robin simulation and evaluate the worst GEO plasma environmental models showing the detailed results. This study will contribute to a revision of the ISO document and also apply spacecraft charging risk estimations as space weather forecast.

Recently, problems involving space debris have become more serious. According to NASA research, the volume of space debris is projected to increase, even if no new satellites are launched. Therefore, debris-removal satellites must be developed immediately. A mandatory function of a debris-removal satellite is to recognize and approach target debris. Thus, visual guidance using image processing is being considered as an effective means of guiding debris-removal satellites toward unresponsive targets. A small satellite is suitable for use as a debris-removal satellite; however, because of weight and/or size limitations, the installation of certain cameras in small satellites is difficult. Thus, we have developed a compact camera system that can perform on-board image processing, by expanding the functionality of an existing camera system to enable it to acquire the multi-direction images required during the satellite-debris rendezvous process. Experiments were conducted using our proposed system on the H-2 Transfer Vehicle (HTV) as part of an electrodynamic-tether experiment. This paper presents a brief report on the results of this HTV flight experiment.

In the development of the recent scientific satellites, requirements of increases in size and improvements in shape accuracy of observation instruments have become more stringent. In order to satisfy those requirements, our research group has examined a pointing control mechanism utilizing artificial thermal expansion as a linear actuator. Our previous research indicated that the pointing control mechanism could satisfy a design requirement of the next generation scientific satellites under a certain orbital environment. It is desirable to conduct performance evaluation under various orbital environments to design a flight model. In order to do that, a detailed thermal mathematical model needs to be built. In this paper, estimated parameters in a thermal mathematical model were estimated by correlating numerical simulation with experiments obtained by thermal vacuum experiments. As a result, the temperature errors between the numerical simulation and the experiments were minimized when the thermal conductance of elastic hinges of the pointing control mechanism was lower than expected. This implies that the estimation of the thermal conductance of the elastic hinge needs to be carefully done.

Ground segment image navigation and registration algorithms and space segment image motion compensation algorithm are presented. The fundamental ground algorithm is based on Kalman filter instead of least squares and does not require ranging or accurate orbit to be provided by flight dynamics. Kalman filter is used to estimate attitude correction angles, orbit position and velocity relative to ideal geostationary orbit, and internal misalignments of imagers with single mirror or two mirrors. Kalman filter measurements consist of landmarks extracted from the imaging instrument level 1A data blocks, maneuver delta V or coarse orbit from flight dynamics, and spacecraft attitude telemetry inserted in the imager wideband data. The ground algorithm is then shown how to be applied to systems with star and landmark measurements, systems with star only measurements with accurate orbit provided by flight dynamics or by the global positioning system, and systems with spacecraft inertial angular rate. Pixel location transformation from imager line of sight reference frame to GOES and CGMS fixed grid frames is also provided for on-ground image registration using resampling.