The development of a fully reusable vertical-takeoff-and-vertical-landing (VTVL) rocket is indispensable for reducing space transportation costs. However, there are many technical issues associated with such vehicles, such as turnover maneuvers during return flight where the pitching moment plays a key role. It is known that aerodynamic characteristics can be controlled by installing aerodynamic devices, but the relationship between the aerodynamic characteristics and the flowfields has not been explored. To clarify this relationship using computational fluid dynamics (CFD), we investigated these flowfields and aerodynamic characteristics, in the case where we install such devices (fins) in the nose part of a reusable rocket. We found that vortices form downstream of the aerodynamic devices. For angles of attack between 0 and 90 degrees (in which the fins are located in the upstream portion), these vortices significantly affect the surface pressure on the rocket and increase the pitching moment. On the other hand, for AOAs between 90 to 180 degrees (in which the fins are in the downstream portion), the effect of these vortices on the on-surface pressure is negligible, and only vortices formed near the surface of the fins increase the pitching moment.

Most of flight vehicles have various protuberant devices on their surfaces, but asymmetry in their positioning with respect to the body axis can affect aerodynamic characteristics of vehicles, particularly roll moment. Thus, it is important in rocket development to clarify the effects of the protuberances on the vehicle aerodynamic characteristics. In this study, as a basic research, we systematically investigated such effects using CFD, by changing the positions of a protuberance. As a result, the roll moment increased nearly linearly with angle of attack (=α), but its trend was different in protuberance locations, particularly when arranged near the center-of-gravity. In positioning there at α = 20 °, the wake vortex center moved farther away from protuberance compared with α = 15 °, then the pressure decline at its wake side was suppressed, and thus, the pressure difference between its upstream and downstream sides became smaller. As a consequence, the roll moment did not arise linearly, but decreased at α = 20 °.

Space Robotics Laboratory (SRL) of Tohoku University is currently developing a 50-kg-class Earth observation satellite "RISESAT". In this paper, we propose a new ground-target tracking control algorithm and attitude control system for RISESAT. The primal mission of RISESAT is multi-spectral observation of the Earth surface using a 5 m GSD High-Precision-Telescope (HPT) with liquid crystal tunable filters. During an observation, the HPT takes a dozen images of a ground target by changing observation bands. General Earth observation satellites using push-broom cameras require high attitude stability to obtain continuous images. However, RISESAT has to track a ground target with an accuracy of 0.1° and 0.008°/s attitude stability because imaging and wavelength switching take few seconds in total. Therefore, we propose a new ground-target tracking algorithm which can keep orientation of ground-target image to a constant direction which satellite operator requires. Evaluations of the algorithms were carried out by hardware-in-the-loop simulator "MEVIμS" which is a satellite system verification environment developed by SRL. We confirm that the algorithm satisfies the mission requirements under the influence of sensor noise and computation time limits.

A small-scale supersonic flight experiment vehicle (OWASHI) is being developed at Muroran Institute of Technology as a flying testbed for verification of innovative technologies for high speed atmospheric flights which are essential to next-generation aerospace transportation systems. The second-generation configuration M2011 of the vehicle with a single Air Turbo Ramjet Gas-generator-cycle (ATR-GG) engine has been proposed. Its transonic thrust margin has been predicted to be insufficient, therefore drag reduction in the transonic regime is quite crucial for attainability of supersonic flights. In this study, we propose configuration modifications for drag reduction on the basis of the so-called area rule, and assess their effects through wave drag analysis, wind tunnel tests, and CFD analysis. As a result, the area-rule-based configurations have less drag than the baseline configuration M2011. However, the effects of the proposed bottleneck on the fuselage below the main wing are smaller than predicted. It would be caused by the drag due to separation and shocks around the bottleneck. It is necessary to redesign the area-rule-based bottleneck to be smoother.

In this study, an uncertainty quantification technique is applied to the thermal design of a pseudo-small satellite. An emulator that uses Gaussian process regression and the least absolute shrinkage and selection operator is employed to efficiently estimate the thermal margin based on the given uncertainties, and sensitivity analysis based on multiple regression analysis is employed to effectively reduce risk in the thermal design of the satellite. The uncertainty quantification results show that the emulator can reduce the uncertainty quantification cost compared to an ordinary simulator, and that the sensitivity analysis can clarify that only two factors are dominant in pseudo-small satellite thermal design uncertainty.

This paper proposes and investigates the use of Adaptive Coding and Modulation (ACM) to maintain high frequency utilization efficiency in accordance with the change of power limits and frequency bandwidth channel limits for the Ka-band high through-put satellite (HTS) system. Improvements in frequency utilization efficiency of the Ka-band satellite channel model have been observed when carrying out ACM using the DVB-S2X modulation scheme and error correcting codes.

GALA (GAnymede Laser Altimeter) is one of the payload instruments of the JUICE (JUpiter ICy moons Explorer) project to be launched in 2022 to the Jovian icy moons Ganymede, Europa, and Callisto. GALA is developed through an international collaboration between Germany, Japan, Switzerland, and Spain. With the performance model of GALA, we have sought to create the interface conditions that satisfy the science requirements on the probability of false detection (PFD) and the range accuracy. The science requirements on GALA performance can be summarized as involving the following four criteria: [1] for Europa fly-by, PFD is less than 0.2 from an altitude of 1300 km or lower, [2] under the worst observation condition for albedo and surface slope of GCO500 (Ganymede Circular Orbit whose height is 500 km), the accuracy of ranging is less than 10 m and PFD is less than 0.2, [3] under the nominal observation condition of GCO500, the accuracy of ranging is less than 2 m and PFD is less than 0.1, and [4] under the best observation condition of GCO500, the accuracy of ranging is less than 1 m and PFD is less than 0.1. For the assessment, however, we had used literature data as the characteristics of the laser detector of GALA, avalanche photodiode (APD), which should be degraded due to the severe radiation environment around Jupiter. Then we carried out a more realistic model simulation of GALA by incorporating these degradation effects of APD. Characteristics of APD, such as gain, quantum efficiency, excess noise index, surface dark current, and bulk dark current, were re-evaluated through radiation tests using the data of dark and photo current of the APD irradiated with 2-MeV-electron and 50-MeV-proton beams, which are the radiation conditions assumed for JUICEGALA around Jupiter. These degraded characteristics of APD by radiation were introduced to our performance model of GALA. As a result, our performance simulation of GALA showed again that the science requirements are satisfied even after taking into account the degraded characteristics of APD. The remaining matter is the effect of noise or digitization in the Analog Electronics Module (AEM), which must be taken into account for the final specifications of GALA.

Today, high-power electric propulsion devices are attracted for the next generation spacecraft; an arcjet thruster is regarded as a promising option because it produces higher thrust to power ratio among electric propulsion. Moreover, the next generation spacecraft would have 400-V bus, and hence, the 400-V bus direct drive of arcjet thrusters would simplify the thruster systems by eliminating voltage converters, and using lower-current cables. However, high-power arcjet operation possibly aggravates cathode erosion, to shorten lifetime. Hence, we propose to use dimethyl ether (DME) as life-extending agent. DME has various preferable characteristics for additives: DME is not toxic nor reactive, and through the thermal decomposition, generates carbon, which would form the layer on the cathode so as to prevent erosion. In this study, we prototyped a 3-kW class water-cooled arcjet thruster to evaluate the effect of DME addition. Cathode erosion rate per power were 0.210 μg/kJ, 0.192 μg/kJ, 0.154 μg/kJ, at mass ratio of DME of 0%, 3%, and 5%, respectively, and hence, DME addition reduced cathode erosion. However, the arc discharge was more frequently unstabilized with increasing DME mass ratio in the range of 0–5%.

A new space transportation system with an expendable solid-fuel booster and a reusable liquid-fuel orbiter is under consideration as part of activities in JAXA to construct a fully reusable space transportation system in the future. This paper shows this new system's conceptual study results, the system specifications, the new technology to be applied, the requirements to the subsystems, and the prospects.

To avoid the severe thermal load within the embeddedly configured Rocket-Based Combined Cycle engine, we arranged the rocket engine location in the present study, that is, the rocket engine was separately mounted for the scramjet flowpath. The rocket engine was taken out from the scramjet flowpath and located on the ramp wall of the scramjet external nozzle. In this case, the ramp wall acted as the additional nozzle in the spike nozzle manner for the rocket internal nozzle. One-dimensional analysis showed that there was an optimal expansion ratio favorable for the scramjet external nozzle. Subsequently, more complex analysis was partially performed using Method-of-Characteristics based two-dimensional wave model with some novel modifications, which can express the pressure mismatching between the exhaust and the ambient flows. The two-dimensional analysis showed that thrust production within the scramjet external nozzle could become much lower due to impingement of the cowl lip expansion waves resulting from the pressure mismatching. This pressure change also sizably reduced the thrust performance of the rocket spike nozzle. It was, however, also demonstrated that presently analyzed nozzle system has great potential to drastically improve its thrust performance by means of controlling the impingement of the cowl lip expansion waves.

We propose a horizontal-pendulum type thrust stand in which two-point acceleration measurement is applied to the null-balance method. The conventional displacement method was been applied to thrust measurement; thrust is determined by pendulum displacement which is caused by thrust. The null-balance method is also applied to the thrust measurement; a controller adjusts actuator force on the basis of pendulum displacement so as to null the pendulum displacement, and thrust is determined from actuator-driving current or voltage. However, both measurement methods can evaluate thrust variation up to one-third of the natural frequency that is determined by the configuration of the controller and pendulum. Hence, we proposed to apply acceleration measurement to the null-balance method, and showed that thrust variation could be measured up to 80 Hz. However, the proposed methods showed errors beyond 80 Hz; the error was caused by the translational flexibility of the torsional hinges. Then, we propose a new method that uses two-point acceleration measurement and adjustment of the center of gravity position to evaluate thrust variation beyond 80 Hz. The prototype evaluated thrust variation accurately beyond 200 Hz, although at 20–160 Hz, underestimating the amplitudes of reference thrust by 50 %.

Solar electric propulsion (SEP) is the dominant design option for employing low-thrust propulsion on a space mission. Spacecraft solar arrays power the SEP system but are subject to blackout periods during solar eclipse conditions. Discontinuity in power available to the spacecraft must be accounted for in trajectory optimization, but gradient-based methods require a differentiable power model. This work presents a power model that smooths the eclipse transition from total eclipse to total sunlight with a logistic function. Example trajectories are computed with differential dynamic programming, a second-order gradient-based method.

This paper proposes an accurate aerodynamic model for membrane structures in free-molecular flow that allows the calculation of torque and force exerted on them, and which can be used for precise analysis of the attitudes and orbital dynamics of the structures. Recently, various types of deployable thin membrane structures attached to satellites for deorbit purposes have been developed and demonstrated on low-Earth orbits (LEOs), and their attitude dynamics have also been investigated. However, in early studies, deorbit sails were modeled as simple flat plates because there were no precise models for membranes that could express wrinkles, billowing, and asymmetric shapes. In this paper, a calculation method using tensor expressions of aerodynamic disturbance is proposed for the design of deorbit devices using high-fidelity membrane models. The proposed method is tested in the attitude stability analysis of membranes on LEOs, and the results show that the asymmetric shapes of the membranes affect their attitude stability.

This paper proposes to apply probabilistic classification technique to stochastic robust flight control design. Stochastic robust control design is an approach to optimize a flight controller based on Monte Carlo Evaluation. The region in the parameter space satisfying design requirements with high probability is estimated using Bayesian inference and is used to search the optimal design parameters. The approach has the potential to reduce computational load in stochastic optimization. It is applied to a simple example problem of flight control design and the result is promising.

A lanthanum hexaboride (LaB₆) hollow cathode with radiative heater has been developed for high power Hall thruster neutralizer. The influences of orifice and keeper shapes on the performance of the hollow cathode and on its operating characteristics are investigated by measuring the discharge voltage and current. Six different orifice shapes are used in this study: straight, long tapered, and short tapered, each with an orifice diameter of 2 or 3 mm. The straight orifice with the 2-mm diameter has the widest range of spot-mode operation, although no effect of orifice shape is observed for the 3-mm diameter. This is because the high neutral-gas density around the orifice fosters transition to the spot mode from the plume mode with a straight Φ2-mm orifice. Regarding the effect of the keeper, having a large exit diameter and the shortest distance possible between the cathode tube and the keeper gives the widest range of spot-mode operation.

These days, spacecrafts using membrane structures, such as a solar sail, membrane solar panels and so on, have been attracting attention for the use in deep space exploration missions. At the orbit insertion with the aerocapture technique, the membrane structure is expected to work as an efficient decelerator with a large area and small mass. However, its flexible structure causes difficulty in the trajectory analysis and the evaluation of the success rate of the aerocapture because its easily deformed shape as well as the uncertainty in the atmospheric properties significantly degrade the reliability for estimation of the aerodynamic forces, the spacecraft motion and the trajectory. To overcome these problems, the coupled analysis code of the aerodynamics, the structural deformation and the spacecraft motion was developed using the particle based method. In this paper, the method and the validation studies were explained in detail. The simulation results for a solar-sail-type spacecraft flying in the upper atmosphere of Mars for the aerocapture were presented.

To realize innovative transportation systems for intercontinental flight and orbital spaceflight, it is essential to establish the necessary aerodynamic, structural, propulsion, and control technologies for a vehicle to fly at high altitudes and speeds within the atmosphere, and to verify them by using small-scale unmanned supersonic experimental aircraft. This paper describes the control technologies for a 3-kg low-speed model airplane flying autonomously from takeoff, through a circuit, to landing. Guidance and control systems are established for the plane, and flight experiments verify that the control laws work well for all flight modes.

The crushable structure realizes not only chuteless landing on the planetary bodies with atmosphere but also on airless ones by absorbing landing shock energy and protects the inner instrument modules against the landing shock within a prescribed deceleration level. The lunar semi-hard landing mission OMOTENASHI proposed by JAXA is selected for launch by NASA SLS in 2018. The present study shows state-of-the-art technology development of the crushable shock absorption structure together with a design example for the small lunar semi-hard impact surface probe of OMOTENASHI.

Laser altimetry is a powerful tool for addressing the major objectives of planetary physics and geodesy. Through measurements of distances between a spacecraft and the surface of the planetary bodies, it can be used to determine the global shape and radius: global, regional, and local topography: tidal deformation: and rotational states including physical librations. Laser altimeters have been applied in planetary explorations of the Moon, Mars, Mercury, and the asteroids Eros, and Itokawa. The JUpiter Icy Moons Explorer (JUICE), led by European Space Agency (ESA), has started development to explore the emergence of habitable worlds around gas giants. The Ganymede Laser Altimeter (GALA) will be the first laser altimeter for icy bodies, and will measure the shape and topography of the large icy moons of Jupiter, (globally for Ganymede, and using flyby ground-tracks for Europa and Callisto). Such information is crucial for understanding the formation of surface features and can tremendously improve our understanding of the icy tectonics. In addition, the GALA will infer the presence or absence of a subsurface ocean by measuring the tidal and rotational responses. Furthermore, it also improves the accuracy of gravity field measurements reflecting the interior structure, collaborating with the radio science experiment. In addition to range measurements, the signal strength and the waveform of the laser pulses reflected from the moon's surface contain information about surface reflectance at the laser wavelength and small scale roughness. Therefore we can infer the degrees of chemical and physical alterations, e.g., erosion, space weathering, compaction and deposition of exogenous materials, through GALA measurements without being affected by illumination conditions. JUICE spacecraft carries ten science payloads including GALA. They work closely together in a synergistic way with GALA being one of the key instruments for understanding the evolution of the icy satellites Ganymede, Europa, and Callisto.

For secure operation of spacecraft, automatic or assistive health monitoring systems utilizing telemetry data are important. However, it is difficult to utilize them comprehensively because they consist of myriad heterogeneous variables. Although various monitoring systems focusing on only a few variables or homogeneous variables have been suggested, a definitive method to deal with the relationship among multiple heterogeneous variables has not yet. This paper proposes a new visualization framework that aims to show the correlation rules underlying multiple variables of spacecraft telemetry data. The proposed framework consists of a change-point detection algorithm based on subspace identification, clustering methods using dimensionality reduction, and a visualization method using heatmaps. In experiments conducted with real telemetry data obtained from JAXA spacecraft SDS-4, the proposed framework demonstrated effective visualizations that reflected the correlations among variables expected from mechanical characteristics of the satellite. Despite differences in scales and/or units, this framework succeeded in visualizing dynamic correlations not only among continuous variables but also among continuous and discrete variables. This framework can be utilized as an initial stage of anomaly detection focusing on the relationship among multiple variables, as well as a method to perceive the overall state of the spacecraft at a glance.

ITF-2, which was developed by the University of Tsukuba YUI project, was designed based on ITF-1. ITF-1 was the first CubeSat at the University of Tsukuba; however, signals were never received. Although ITF-2 was developed based on ITF-1, the components that resulted in failure were reconsidered and improved for use in ITF-2. In this paper, the overview, subsystems of ITF-2, and the design improvement from ITF-1 to ITF-2 are presented. In addition, ITF-2 was deployed from “Kibo” on January 16, 2017 and it is still operational as of September 2017. This paper reports the operation results and the achievement of missions: “YUI network”, which is a “participatory” mission using CubeSat, involving participation from people around the world, ultra-small antenna, which is used for two amateur bands by attaching it to a metal structure, and a new microcontroller, which has been reported to be saving energy, high radiation resistant and demonstrated to acquire a new achievement in space. The operation results as of September 2017 indicate that more than 900 reception reports from 19 countries were reported, indicating the success of the outreach.

In recent years, consumer parts have been used in nano satellites or small satellites. Radiation resistance testing is a useful method of evaluating operational stability in the space environment. Since the single event effect (SEE) is a probabilistic phenomenon, it is necessary to monitor the test pieces at all times during the radiation test. If response to a serious error such as SEL is delayed, the test pieces may be damaged or broken, and the accuracy of the test result may be reduced. In this research, a test support system that can instantly detect the occurrence of an SEE and respond automatically was developed in order to improve the efficiency and accuracy of radiation tests. Radiation resistance tests for consumer microcomputers and communication module were conducted using the test support system. From the test results, the effectiveness of the test support system and improvement points were confirmed, and the possibility of space application of consumer microcomputers and communication module were evaluated.