Fast tracking algorithm discussed in this paper is applied to star trackers for improving the performance of star identification. Space Robotics Laboratory (SRL) in Tohoku University has developed star trackers for micro-satellites so far. Since it has only a lost-in-space algorithm for star identification, the attitude update rate is limited up to 1 Hz. It was implemented to the Philippines' 50 kg-class micro-satellite “DIWATA-1” released from the International Space Station on April 2016. Although on-orbit evaluation showed good results enough to output attitude autonomously, the performance of continuous attitude determination was worse than expected. Since quite a high access frequency of star catalog is required, timeout of the process for attitude calculation occurs frequently even if update rate is 1 Hz. Insufficient ground evaluation before launch is also one of the causes of operation failure. Tracking algorithm helps to calculate latest attitude faster than conventional methods by feeding back the previous attitude information. This algorithm includes two additional processes. First, future star positions on the image frame can be predicted according to the previous attitude and pre-identified star information. The sensor can find corresponding latest centroids compared to predicted star positions. Second, un-identified stars on the camera field of view (FOV) can be detected by referring to the star neighborhood catalog, which includes the list of some adjacent star IDs against each reference star. PC simulation shows that continuous attitude determination works effectively by keeping low catalog access frequency. The proposed algorithm is implemented to the real hardware. Then, ground evaluation is conducted using star simulator environment and satellite dynamics simulator. The result demonstrates that the processing speed in real situation becomes about 70 times faster compared to the previous method and it is successful to obtain much more stable 1 Hz attitude output.

The “Compressible and Non-Expanding (CNE)” fluid model was applied for simulation of a high-speed impact on a regolith-like target material, in which irreversible compression occurs. The equation of state for the CNE fluid is composed of the irreversible compression curve and the reversible elastic curves. For numerical simulation with the finite volume method, we developed the Riemann solver, whose fundamental solutions are the combination of the shock waves, the elastic waves, the contact discontinuities and the boundary between the regolith and the vacuum. To track the boundary by the finite volume method, the volumetric fraction function and the cell pattern identification method were introduced. To demonstrate that the present CNE fluid model is promising for the simulation of high-speed impact, the crater formation on the regolith-like target was simulated by the two-dimensional analysis and the result was qualitatively reasonable. The effects of the size (depth) of the sand box and the non-uniformity of the target material were investigated, assuming the one-dimensional approximation on the centerline of the impact.

Japan Aerospace Exploration Agency (JAXA) will launch an Earth observing satellite for climate studies named “Global Change Observation Mission – Climate (GCOM-C)” in 2017 which carries a multi-spectral optical sensor named Second-Generation Global Imager (SGLI). The GCOM-C satellite will observe various geophysical variables such as vegetation, land surface temperature, aerosol, clouds, ocean color, sea surface temperature, snow cover extent, snow grain size and so on. The objectives of the SGLI observations are to elucidate the roles of the geophysical variables in the recent changing Earth's climate system and to establish long-term satellite data record of the variables. Three kinds of target accuracies to be achieved for the SGLI observations are defined for evaluating the success of the GCOM-C satellite mission. First one is the minimum thresholds to be achieved for the first SGLI data release at one year after the launch of GCOM-C. Second and third thresholds are the standard and goal accuracies to be achieved at five years after the launch for evaluating the full and extra success of the mission. The target accuracies for individual SGLI products are available at the GCOM-C web site (http://suzaku.eorc.jaxa.jp/GCOM_C/index.html). The quality and accuracies of SGLI products are planned to be evaluated and maintained through validation activities organized by the SGLI validation team consisting of JAXA and SGLI principal investigators (PIs). Uncertainties of the SGLI data products will be characterized through the comparison with in-situ observations, similar products derived from other satellites, climatological data, and/or numerical model simulations. This paper summarizes the overall validation plan for the SGLI geophysical variable products.

The Japan Meteorological Agency successfully launched Himawari-8 in October 2014, and began operations in July 2015. A new sensor system, Advanced Himawari Imager (AHI), is observing the surface of the Japan area frequently (every 2.5 min) and 24 images can be observed sequentially each hour. This study examined the feasibility of multi-frame image processing using Himawari-8/AHI data obtained over one hour. Slight distortion of images was adjusted by dividing one pixel into ten subpixels and by applying automatic registration, which facilitated multi-frame image processing. As a result, Himawari-8/AHI data is improved effectively to allow closer observation of the details of an image.

In order to evaluate the accuracy of footprint measurement for a satellite large scale deployable antenna, we proposed a method of approximating antenna pattern shape with a quadratic function and determining its parameters by least squares method. Moreover, to obtain the required value of the measurement accuracy at the receiving station, the distribution of the boresight position when an error is added to the reception level was calculated. This method was applied to the experimental value of ETS-VIII. As a result, when the variation of each receiving station is within 0.5 dB, it is possible to estimate the boresight position within 1/10 of the beam width.

This paper presents an aeroelastic analysis for the high altitude propeller designed for the stratospheric airship KFG-series. Aerodynamic analysis is carried out in Fluent by using sliding mesh method and a loosely coupled method is employed to solve the aeroelastic response. The real composite propeller structure is modeled in Ansys Acp module and Mechanical APDL enable the data exchange between CFD and CSD solver. Spring-based smoothing method is adopted for updating the mesh deformation. Firstly a comparison between experimental and numerical results is employed for validating the accuracy of the numerical model. Then in fluid-structure analysis, aeroelastic response of the blade tip leading edge shows a limit cycle oscillation and the blade averaged deformation shows a very small deflection due to the dynamic stiffness effect. In the final, equivalent stress of the blade material and a comparative study of the aerodynamic performance for rigid and deformed propeller at the design condition are presented and it is found that there is a lower thrust performance for the deformed propeller due to the small pitch variation.

The validation of deformation caused by the thermal expansion of structures is an important issue for the success of a mission that requires the large-scale and precise positioning of instruments. Therefore, an accurate alignment monitor system was installed in the X-ray astronomy satellite Hitomi to correct observation data. We compared the analysis results of a finite element model with the flight telemetry data of the alignment monitor installed on Hitomi to validate the finite element model. As a result, it was shown that the finite element analysis was able to reproduce the global and slow trends of thermal deformation on orbit. However, the analysis could not reproduce the rapid changes in thermal deformation. To realize more precise and large structures in the future, we need to identify these rapid changes and correctly improve the FE model of thermal deformation through ground tests prior to launch.

We propose an equalization technique that uses the characteristics of the satellite transponder to compensate for its non-linear distortion. In our previous study, the effects of this technique for 16APSK and 32APSK were assessed in simulations. In this study, we revised the configuration of the equalization system and developed a receiver based on this revision. We examined the reception performance of 16APSK and 32APSK in transmission experiments.

A satellite in an elongated elliptical orbit may be operated while switching the configuration for communications according to changes in altitude. Although such operation is generally complicated as compared with a satellite in low Earth orbit, there is still a common desire to maximize the amount of downlink data. To this end, the operation plan for communications configuration needs to be optimized in detail according to momentary changes, so that more data can be downlinked. Conversely, switching the configuration many times leads to an increase in the operational load for satellite operators. We thus want to control the number of configuration changes. This problem of creating such an operation plan is too complex to be dealt with solely by human operators. In particular, as there are more options in configuration, the number of combinations increases exponentially. Therefore, by representing the operation plan graphically, this paper reduces the planning problem to a shortest path problem, and solves it to obtain an operation plan for maximum downlink. We also introduce parameters to adjust configuration changes, and obtain the correlation between the amount of downlink data and the switching frequency.

Additive manufacturing is recently attracting attention as an innovative production technology of aerospace components to drastically reduce manufacturing time and cost, and improve design flexibility and productivity. In this paper, the feasibility of applying the direct metal laser sintering (DMLS) method to a regeneratively cooled combustion chamber inner liner made of copper alloy and that of applying the laser metal deposition (LMD) method to a combustion chamber outer shell made of Ni-based alloy were investigated. The DMLS and the LMD parameters, such as laser power, beam scan speed, and so on were investigated to establish the fabrication conditions. Heat treatment conditions such as, hot isostatic pressing (HIP), solution treatment, and aging were also investigated to improve the thermal conductivity and the tensile strength of the copper alloy. Interfacial bonding strengths between the copper alloy substrate and the Ni-based alloy outer layer were also evaluated. The 30 kN class LOX/LCH₄ combustion chamber inner liner in which coolant passages were located circumferentially in the liner wall was successfully fabricated by DMLS, and the molding of the outer shell around the inner liner by LMD is now being prepared.

In this study, we developed a high-performance and high-thrust hybrid rocket motor using low-melting-point thermoplastic (LT) fuel and swirling oxidizer flow. LT fuel has excellent mechanical and adhesive properties, as well as a high regression rate compared to conventional hybrid rocket fuel. In this study, we conducted several firing tests using swirling oxidizer flow to obtain the fuel regression rate and evaluate its effects on the geometric swirl number (S_g). We determined that the average regression rate of the LT fuel with S_g = 37.3 was ~2.9 times larger than the axial flow test value. The LT fuel was more susceptible to swirling flow than polypropylene, presumably due to the different physical properties of the fuels. In the swirl flow experiment, we confirmed that the local fuel regression rate behind the fuel is uniform, and it differs from the regression rate seen in the axial flow experiment. For the range of oxygen mass flux values G_oxlo = 30–72, ṙ_loave was fitted to a conventional formula. The results of this fit suggested that the local regression rate at the head region of low-melting-point fuel, such as the LT fuel, cannot be represented only by chemical reactions; therefore, the fluid dynamics of liquefied fuel must be included in the model.

The flowfields of a self-field magnetoplasmadynamic (MPD) thruster using hydrogen propellant were numerically simulated with a physical model incorporating the ion-slip effect. Thrust performance was investigated for two anode configurations, namely, straight anode and flared anode at discharge currents between 5 to 8 kA. Simulation results show that thrust efficiency increases with increased discharge current for the straight anode, while for the flared anode, thrust efficiency tends to decrease; this opposite trend is caused by the ion-slip effect. When comparing thrust characteristics, thrust for the flared anode was found to be larger than that for the straight anode, but the advantage of the flared anode diminishes at higher discharge currents due to strong pinching and consequent pressure depletion in the vicinity of the flared anode surface. This pressure depletion leads to large electric power consumption owing to the ion-slip heating. That is, at lower pressures, the ion-slip effect becomes more significant because collisions between ions and neutral atoms are not frequent.

This study investigated effects of aluminum powder in a melting experiment and ignition experiment using an electric furnace for microcrystalline wax fuel with 30 μm aluminum powder added. Melting experiment results show that the melting time of the solid fuel decreases. Therefore, the aluminum powder promotes the phase change of WAX from solid to liquid near the surface of the solid fuel in the boundary layer combustion. Ignition experiment results show that adding aluminum powder reduced the time to ignition. By adding aluminum powder to microcrystalline wax, solid fuel ignites easily due to the ignition heat quantity per unit mass decreased. However, the aluminum powder was not ignited. Nevertheless, by adding aluminum powder to microcrystalline wax, the thermal conductivity increases, which promotes solid fuel melting and ignition.

Optical communication technologies are regarded as an important breakthrough for future satellite communications. Tohoku University has been developing Adjustable beam angle Small Optical TrAnsponder (ASOTA) for micro-satellites. The design concept of ASOTA is miniaturization and low power consumption. It is necessary for space optical communication to direct the laser to the ground station with a high precision. This research describes the development of a MEMS-Mirror-based Fine Pointing Mechanism (FPM). The MEMS mirror is suitable for micro-satellites because it is smaller and consumes less power than 2-axis gimbal mechanisms. FPM is a feedback control system consisting of a MEMS mirror and a Quadrant Detector. The characteristics and mathematical model of each device are described. Then numerical simulation and ground evaluation results of FPM with a simple optical system are summarized.

We evaluated the ignition probability of high-energy ionic-liquid droplet due to breakdown from visible pulse laser. The high-speed camera result, pressure profile, temperature trace from an R-type thermal couple, and radical measurement results using an image-intensifier charged-coupled device camera revealed the ignitability of this system. In addition, we simulated a simple thermal-diffusion model and investigated the temperature distribution for comparison with the experimental results. As a result, the probability of ignition due to a single pulse is low; however, potential for gasification and atomization of the ionic-liquid droplet is probable.

The Life Detection Microscope (LDM) is the instrument to take microscopic images of powder samples on Mars surface. The LDM is designed to use fluorescence pigments to detect organic compounds and those surrounded by membrane, and then to test the esterase catalytic activity to detect metabolic reaction, as well as to take microscopic images of regolith and dust particles. The LDM has high resolution to resolve terrestrial prokaryotes (primitive microorganism), 1 μm/pixel, and high sensitivity, 10⁴ “cells” per gram powder, by inspecting sufficient, 1 mm³, volume of powder sample. We propose to conduct extant, not only past, life search program on Mars surface.

Power supply in an electrically closed system such as a spacecraft is strictly constrained. Conventionally, constrained power in spacecraft is supplied to internal subsystems based on their expected demands. When designing, a spacecraft designer assumes those subsystems as static system even in cases that some of the subsystems behave dynamically. In that way, however, power margins are often overly estimated to avoid scarce power supply. This paper proposes a novel power method using hybrid control under multi-agent structure in order to fully utilize constrained power of an electrically closed system. The proposed system structure owns agents that have two modes; static mode and dynamic mode. In static mode, power request from an agent is simply based on its referenced status under steady state. In dynamic mode, an agent requests power based on optimized input pattern that is calculated under transient state. Among agents with the two modes, ones with dynamic mode receive more power than static ones because optimally-calculated input patterns in dynamic mode are delicate while behaviors in static modes are stable even when input patterns slightly change. This paper demonstrates the effectiveness of the proposed method by simulation compared to results with central computing system.

Hybrid rocket propulsion offers inherent safety during handling and launch operations at low cost. This makes it not only attractive for space tourism applications like SpaceShipTwo but also for sounding rockets and for educational activities with students. This has been successfully demonstrated by the HyEnD student project from the University of Stuttgart: On November 8^th, 2016 at 10:30 a.m. the hybrid sounding rocket HEROS 3 was launched from the Esrange Space Center to an apogee altitude of 32,300 m (106,000 ft). This set a new altitude record for European student and amateur rocketry. Furthermore, this is a world altitude record for hybrid rockets built by students. It was successfully recovered with the drogue and main parachute being released. The rocket was powered by a 10 kN design thrust hybrid rocket engine with a paraffin-based fuel and Nitrous Oxide (N₂O) as the oxidizer. The combustion efficiency was verified to be above 97% in ground tests. Liquid burn time in the flight was for 15 s with an additional combustion of gaseous N₂O in the blow-down mode of about 10 s. The rocket was 7.5 m long with an empty mass of about 75 kg. The rocket structure was made completely from light-weight carbon fibre and glass fibre reinforced plastic.