The degradation of materials in extreme environments is a critical issue affecting the reliability of mechanical systems in space. A low earth orbit (LEO) environment in particular has various factors that strongly affect materials. A number of materials for space applications were exposed to an LEO space environment by the Japan Experimental Module / Space Environment Exposure Device (JEM/SEED) experiment aboard the International Space Station (ISS) to evaluate how they would be affected by a real space flight environment. A solid lubricant coating was one of the experimental materials evaluated in the JEM/SEED experiment. In addition to the orbital evaluation, the same type of coating was irradiated with atomic oxygen (AO), and ultraviolet rays (UV) on the ground. The fluences of AO and UV irradiation corresponded to the exposure in LEO environment around the ISS during the SEED experiment. Samples from LEO exposure, ground irradiation with AO and UV, and a reference (non-flight and non-irradiated) sample were then subjected to friction tests in a vacuum and surface analyses. The results were compared to elucidate the effects of the various factors on the characteristics of the solid lubricant film.

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We have been building up a brand new ablation analysis code which is intended to predict simultaneously thermo-chemical and thermo-mechanical response of the Silica-Phenolic (SiFRP) ablator. In this paper, the model is applied to a ground firing test of a liquid rocket engine. Although the predicted thermo-chemical behavior of SiFRP shows good agreement to the measured data (temperature and char-penetration depth), predicted thermo-mechanical behavior (strain) shows limited success. The present tentative model gives better agreement to the strain data that are measured during combustion at a firing test by intentionally changing the thermal expansion data of SiFRP for perpendicular to the ply-direction in the low temperature region. It is obvious that we are still in short of validation studies. Therefore to attain further refinement, the more extensive validation studies are planned and conducted in the future.

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A non contact whole field shape measurement using the grating projection method was performed to visualize formation and dissipation processes of the wrinkling phenomena on thin membranes, and their nonlinear mechanical properties were discussed. The wrinkling phenomena in a rectangular polyimide film, which held fixed at the bottom side and subjected to loadings at the upper side, was treated. The measurement results showed that the nonlinear wrinkling behavior such as the snap through phenomena appeared in the formation and dissipation processes of wrinkles was successfully captured. Through the discussion on the dissipation processes of wrinkling phenomena, the conditions of the applied tensions to reduce the distortion of the membrane surface feature are presented.

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This paper considers a spacecraft attitude maneuver that is performed by two or three firings of four impulsive thrusters. The minimum number of thruster firings is two when a rest-to-rest attitude maneuver of a spacecraft is performed by impulsive thrusters. However, when the maneuver is performed using three thruster firings, the total fuel consumption can be reduced. Therefore, we derive a method to obtain the maneuvering solutions for two or three thruster firings. First, we decompose the eigenaxis rotation into two successive rotations, and derive the approximate solutions for three thruster firings. Then, we derive the accurate solutions that satisfy the attitude dynamics by using a numerical iteration method. The efficiency of using three thruster firings, which are found to reduce the total fuel consumption, is verified by numerical simulations.

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The Japan Aerospace Exploration Agency (JAXA) is now planning the next-generation infrared astronomical mission - the Space Infrared Telescope for Cosmology and Astrophysics (SPICA). SPICA is the first Japanese libration point mission that utilizes a halo orbit around L2 in the Sun-Earth system. This paper describes a study conducted on the stationkeeping strategy for that libration point mission. The main algorithm is structured with reference to a paper written by K. C. Howell et al. [1993]. In that paper, the attitude constraints of SPICA and frequent disturbances caused by unloading operation of the reaction wheels are considered in applying the stationkeeping algorithm to SPICA. Throughout the analysis, the amount of corrective delta-V for stationkeeping was revealed as being heavily influenced by thruster allocation when considering the strict attitude constraints of SPICA. The total delta-V for one year of stationkeeping varies from 0.1 to 7.6 m/s, depending on onboard thruster allocation. Based on the analysis results, this paper suggests an optimal thruster allocation for SPICA-like missions.

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In this paper, a new driving method for a stepper motor is proposed. The proposed method can compensate a detent torque of the stepper motor and reduce the disturbance without angle sensors. First, this paper describes studies on conventional driving methods for the stepper motor, full step and micro step driving method. The performance of reducing the disturbance by these conventional methods is investigated through numerical studies. Second, the details of the proposed method are described. The compensation input for a variation of the detent torque is derived. The performance of the proposed method is compared with the conventional methods. The proposed method can drastically reduce the disturbance. Finally, the effects of the sensor less drive of the proposed method on the performance are discussed.

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This paper discusses the dynamics and control of an extendable arm for tether-based mobile robots on the International Space Station (ISS). The Japan Aerospace Exploration Agency (JAXA) is developing the robotic system which has an extendable arm and several tethers to support extravehicular activities (EVAs) by astronauts. The robotic experiment, called REX-J, will be conducted on the Japanese Experiment Module (JEM) of the ISS in 2012.The robot used in REX-J has an extendable and retractable arm that deploys tethered-grippers and attaches them to holding points. The arm is highly flexible due to its light-weight design and slender structure. This flexibility causes vibrations that make operation inefficient and difficult. To cope with this issue, we propose a dynamic model and a vibration control method for the arm, and verify their effectiveness by numerical simulation and experiment. The dynamic model of the arm is introduced as a planar two link model with a virtual passive joint. The proposed control method is capable of suppressing the vibrations of the arm using the inertial coupling between the arm and the wrist attached to the arm tip. The results of the simulation and experiment show the validity of the proposed model and control method.

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Interplanetary spacecraft require precise predictions of their trajectory for achieving their mission targeting requirements. For the purpose of precise navigation, the detailed modeling of all small forces affecting the spacecraft’s orbital motion is necessary. The dominating small forces on an interplanetary trajectory are planetary gravity perturbations and Solar Radiation Pressure. In addition to these traditional perturbations, the Thermal Radiation Pressure induced by the spacecraft surfaces can be an appreciable small force. This study focuses on the precise thermal analysis of the external surfaces by using a numerical Finite Element Method on three actual ESA interplanetary spacecraft, i.e. Rosetta, Mars Express, and Venus Express. This paper also presents the evaluations of the accelerations due to the Thermal Radiation Pressure.

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A novel crater detection method for moon landing system is proposed. This proposal assumes that it is applied to SLIM (Smart Lander for Investigating Moon) project aiming at the pin-point landing to the moon surface. The point where the lander should land is judged by the position relations of the craters obtained via camera, so the real-time image processing becomes important element. Besides, in the SLIM project, 400kg-class lander is assumed, therefore, high-performance computers for image processing cannot be equipped. The image processing is going to realize high-speed processing by using the FPGA circuit for parallel computation. Furthermore, as the crater detection method, we try to apply an object recognition method using Haar-Like feature. This Haar-Like feature is a technique used for the face detection with the digital camera, and is suitable for rapid object detection based on distribution of the light and shade of the image. Generally the surface of the moon image is only the light and shade, therefore, this method is suitable for the high-speed detection of the crater. In this paper, the precision and rapidity of the crater detection using Haar-Like feature is shown by computer simulation.

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New attitude and attitude rate estimation algorithms are presented for a spacecraft using magnetometers. An attitude rate estimation algorithm solves an optimization problem to identify the unknown angular rate from the short period observations of the earth magnetic field vectors, and then, estimates the angular velocity vector of the spacecraft. An attitude estimation algorithm solves another optimization problem to identify the unknown angle from the long period observation of the earth magnetic field vector and the estimated attitude rate, and then, estimates the spacecraft attitude in terms of Euler parameters. The effectiveness of the proposed algorithms is evaluated by numerical simulations.

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This paper proposes a design method of optimal DVDFB (direct velocity and displacement feedback) controller for attitude control of flexible spacecraft whose flexible solar paddles rotate at the orbital rate. The DVDFB controller with collocated sensors and actuators is known to be promising from the robust stability viewpoint even for such a linear parameter varying system. However, the optimal design method is not still well-developed. For this problem, we propose to design an optimal DVDFB controller gain based on LMI (linear matrix inequality) in the H_∞ control framework using GKYP (generalized Kalman-Yakubovich-Popov) lemma. Some numerical simulations are performed to demonstrate the ability.

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A radar for navigation in future Japanese lunar/planetary landing missions is being developed in the Japan Aerospace Exploration Agency (JAXA). The C-band pulse radar provides not only altitude information but also relative velocity against the surface. In this paper, detailed design of a Doppler simulator for the landing radar is described. The simulation accompanied with Digital Terrain Models (DTM) is quite effective because it can analyze accuracy of the velocity measurements while taking account of terrain effects. The validity of the simulation is quantified by comparing with the results of actual field experiments. Furthermore, the performance of the landing radar over lunar terrains is evaluated by the simulation with the KAGUYA DTM products of lunar surfaces.

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When a spacecraft lands, the large shock load can lead to undesirable responses, such as rebound and trip. The authors have previously discussed the problem of controlling these shock responses using momentum exchange impact dampers (MEIDs). However, the optimal design parameters of MEIDs for spacecraft landing have not yet been addressed. These parameters are crucial for MEID applications. This paper discusses the parameters of Passive-MEID (PMEID) for a single-axis falling-type problem, which is the most fundamental problem. It is found that the rebound height is proportional to the mechanical energy of the spacecraft. Thus, the optimal design parameters of the PMEID correspond to the parameters that minimize the mechanical energy. A PMEID with the optimal design parameters is called optimal PMEID in this paper. In order to improve the performance of the optimal PMEID, this paper proposes a novel MEID — HMEID (active/passive-hybrid-MEID). The HMEID combines actuators with passive elements such as contact springs. Based on the optimal design results for the MEIDs, this paper applies a stiffness control to the HMEID in order to suppress the mechanical energy further. Simulation studies reveal that the HMEID can effectively reduce the influence of shock responses. The robustness of the HMEID against the landing ground is shown. The feasibility of the HMEID is also discussed. The HMEID is superior to a PMEID, even if the actuator has a dynamics with a large electric time constant.

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It is essential to perform dynamic simulation in order to verify the safety of capture operation at time of contact between the grapple fixture of the H-II Transfer Vehicle (HTV) and the Latching End Effector (LEE) of the Space Station Remote Manipulator System. This paper presents the contact dynamics model for a snare wire inside the LEE in capturing the grapple fixture of the HTV. An explicit method for modeling contact dynamics is chosen so as to achieve a simple form and low computational cost in dynamic simulation. A contact model is developed using a rigid bar (grapple fixture) and a rigid wire. Wire stiffness is derived from the simplified equation of motion of the wire. The proposed stiffness model enables one to predict contact force during capture operation. Experimental measurement of the wire stiffness in static conditions verifies the proposed stiffness model. For the verification of the proposed model in dynamic conditions, a contact dynamics experiment is conducted using an air-floating test bed, which is capable of emulating two-dimensional planar micro-gravity environment. The experiment and corresponding dynamic simulation confirm the validity and usability of the proposed contact dynamics model.

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This paper presents a new adaptive flight control system using the backstepping method for a future space transportation system. A space transportation system using the backstepping method can be constructed with an expansion of a control system using feedback linearization with time-scale separation. It is difficult to obtain prior knowledge in detail, such as that regarding the aerodynamic forces and moments, and wind disturbances. Furthermore, the necessity of the derivative of pseudo-input is considered an issue in control systems using the backstepping method. Therefore, in this study, making use of a disturbance observer, we propose a method which can estimate dynamic properties of a vehicle and the derivative used as pseudo-input. Moreover, we attempt to guarantee the stability of the entire system, including the estimation mechanism, controller, and vehicle dynamics. We conducted numerical simulations to verify the effectiveness and robustness of the proposed system.

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The three-dimensional reaction wheel (3DRW) is a three-axis free rotor reaction wheel, which can achieve a compact and high integrity attitude control systems. In the system we propose, a spherical rotor is levitated without mechanical contact by gas bearings and rotated by the torque generated by magnetic fields. To realize the precise control of the rotor, it is important to understand the characteristics of spherical gas bearings such as pressure distribution, load capacity, stiffness and flow rate. Some previous works clarified them but few researches are there relating the use of gas bearings in the space. What is required for the space gas bearing is having the minimum robust characteristic against various small disturbances. It is more appropriate to adopt small-light air compressor which can provide the minimum air pressure and flow than big-heavy powerful one. In this study, the authors have developed a spherical gas bearing with a circular slot restrictor and analyzed its performance by experiments and numerical calculations. The result shows that this type of bearing can generate necessary floating force with low supply pressure and low flow rate, and that leaves the possibility to realize 3DRW system operated by a compact pump such as piezoelectric type.

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In recent years, LIDAR has been used in remote sensing systems, obstacle avoidance systems on planetary landers, rendezvous docking systems, and formation flight control systems. A wide dynamic range is necessary for LIDAR systems on planetary landers and in rendezvous docking systems. For example, a dynamic range of 60 dB was required for the receiving system used in the Hayabusa mission in order to measure distances between 50 m and 50 km. In addition, an obstacle detection and avoidance system of a planetary lander requires a ranging resolution of better than 10 cm. For planetary landers, ISAS/JAXA is developing a customized integrated circuit (IC) for LIDAR reception. This report introduces the design of the customized IC and reports the results of preliminary experiments evaluating the prototype, LIDARX03.

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A spherical projectile was launched with 110% – 180% of a Chapman-Jouget (C-J) velocity into a detonable mixture, and we investigated the oblique detonation wave (ODW) that stabilized around it. High time-resolution visualizations were conducted using a high-speed camera with 1-μs frame speed to directly confirm the ODW stabilization and to investigate an unsteady phenomenon observed near the stabilizing criticality. In this case, the ODW was a three-dimensional conical wave, and the curvature effect on the conical detonation wave is not negligible near a projectile. We investigated the wave velocity distribution along the wave and revealed that it had a local minimum point at 0.8 – 0.9 times a C-J velocity during the decay process from an overdriven detonation near a projectile to a C-J ODW in the far field. We defined a characteristic wave curvature radius normalized by a cell size on this local minimum point. In this study, the minimum characteristic wave curvature radius of about 18 was needed to stabilize the conical detonation wave around a sphere. Near the stabilizing criticality, we also observed the unsteady ODW stabilization or detonation initiation on a shock-induced combustion. This unsteady regime was characterized by periodical onsets of local explosions that initiate or stabilize an ODW. We investigated the wave velocity distribution along this regime, and our findings revealed that the ODW transition or the detonation initiation following the shock-induced combustion occurred when the wave velocity decayed to 0.5 – 0.6 times a C-J velocity.

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Visualization experiments employing rectangular cross-section curved channels were performed in order to examine the fundamental characteristics of a curved detonation wave propagating stably through an annular channel. A stoichiometric ethylene-oxygen mixture gas and five types of curved channels with different inner radii of curvature were used. The detonation waves propagating in the curved channels were curved due to the expansion waves from the inner walls of the curved channels. The ratio of the inner radius of curved channel (ri) to the normal detonation cell width (λ) was an important factor determining the stability of the curved detonation waves. The detonation propagation mode in the curved channels transitioned from unstable to stable in the range 14 ≤ ri/λ ≤ 26. The normal detonation velocity (Dn) of the curved detonation wave propagating stably in a curved channel was approximately formulated. The approximated Dn given by the formula agreed well with the experimental results. The front shock shape of the curved detonation wave could be reconstructed accurately using the formula. The value of Dn nondimensionalized by the Chapman-Jouguet detonation velocity became a function of the local curvature of the curved detonation wave (κ) nondimensionalized by λ regardless of the shape of curved channel. The front shock shapes of the detonation waves in the stable mode became similar to each other under constant ri/λ conditions.

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The effects of a vortex flow (VF) on the characteristics of deflagration-to-detonation transition (DDT) were examined in order to achieve the shortest distance of DDT for a pulse detonation engine (PDE). The DDT distances in a vortex flow were shortened by 15-47 % than those in a counterflow. The shortening effect becomes remarkable as the rotating velocity increases. Formation of the area of higher energy density in the ignition domain of the tube, and flame acceleration due to rapid flame propagation in the vortex flow and promotion of turbulence near the tube wall by the rotating velocity in the transition domain of the tube are considered to be the governing factors in shortening the DDT distance.

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The thermal characteristics of the low-density ablator are quantified by conducting the heating tests of materials in an arcjet wind tunnel. In the tests, test pieces of the ablator with a diameter of 40 mm and a length of 58 mm are put into arcjet flow. The arcjet wind tunnel is run for two sets of operational condition: a low heating condition for the heat flux of 0.97 MW/m² and the impact pressure of 1.93 kPa, and a high heating one for 1.97 MW/m² and 4.25 kPa, respectively. The surface temperature and in-depth temperature are measured during the testing. The surface shape changes of ablator and the amount of mass loss of ablator are measured after the heating tests. The obtained experimental data are analyzed by using an integrated computational method developed earlier. Agreement between the calculations and the heating tests are excellent in the comparison of the surface shape changes of ablative test pieces. The temporal variations of surface temperature are also well reproduced by the present numerical method within the experimental error. However, the calculation underestimates the in-depth temperatures and the amount of mass loss of ablative test piece. The reason for this discrepancy may come from the lack of the thermal conductivity data especially in high temperature range above 500K. In addition, the present study suggests that the thermal conduction of the ablative test piece used in this study could be anisotropic to some extent.

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A space vehicle which undergoes the atmospheric re-entry or a planetary entry needs the heat shield system to protect inner equipments against severe aerodynamic heating environments. Charring ablator is usually used for the heat shield system. In order to design the heat shield system, it is necessary to predict the thermal behavior under aerodynamic heating by ablation analysis. A computer code for charring ablation and thermal response analysis is newly developed for simulation of one-dimensional transient thermal behavior of charring ablation materials. The mathematical model for the charring ablation including basic equation and computational method of ablation analysis is briefly described. A new ultra light weight phenolic carbon ablator called LATS (Lightweight Ablator series for Transfer vehicle) was recently developed. Arc-heated tests of the LATS ablator were carried out and measured results of the temperature response and surface mass loss are compared with the simulation results of the ablation analysis program. The agreement between the results of simulation and measurement is found to be good. It is also found that the mathematical model used in the ablation code can be applied to the ablation analysis of the low density LATS ablator.

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In order to investigate silicone carbide (SiC) ablations in detail, the test pieces of SiC were ablated using air plasma freejets, and spectroscopic measurements were performed at three points around the ablating SiC. It was found that spectra measured at each point were quite different with each other both in distribution and in intensity. At the point just upstream of the tip of the test piece, the molecular bands of CN, N₂, and the atomic lines of Si were observed, and in the downstream region, that is, in the pale greenish flow region, strong molecular bands of C₂ appeared. It was suggested that the C₂ molecules were generated by recombination reaction of C atoms originated from the ablation of SiC. Experimental spectra of radiation from the heated test piece were fitted with the black body radiation function to estimate its surface temperature. The brightness temperature thus measured was 3,600K and this was considerably higher than the decomposition temperature of SiC.

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In the present research the characteristics of nonequilibrium flow are studied by applying CARS (Coherent Anti-Stokes Raman Spectroscopy) method to hypervelocity shock waves in low-density air. Furthermore, in order to get more detail information on the flow behind the hypervelocity shock wave and to inspect whether the disturbance is seen at the shock wave front during the CARS measurement, radiation images behind the shock waves are obtained simultaneously. Finally, the CARS signals are detected behind hypersonic shock waves with a shock velocity of up to 7km/s. In addition, the radiation distributions are obtained at the same experiment. The vibrational and rotational temperatures are estimated by spectral fitting method. It is found that the rotational temperature exceeds the vibrational temperature when shock wave velocity is 5km/s or more, where both temperatures remains lower than translational temperature.

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In this study, aerodynamic design exploration for reusable launch vehicle (RLV) is conducted using genetic algorithm with Navier-Stokes solver to understand the aerodynamic characteristics for various body configurations and find design information such as tradeoff information among objectives. The multi-objective aerodynamic design optimization for minimizing zero-lift drag at supersonic condition, maximizing maximum lift-to-drag ratio (L/D) at subsonic condition, maximizing maximum L/D at supersonic condition, and maximizing volume of shape is conducted for bi-conical shape RLV based on computational fluid dynamics (CFD). The total number of evaluation in multi-objective optimization is 400, and it is necessary for evaluating one body configuration to conduct 8 CFD runs. In total, 3200 CFD runs are conducted. The analysis of Pareto-optimal solutions shows that there are various trade-off relations among objectives clearly, and the analysis of flow fields shows that the shape for the minimum drag configuration is almost the same as that of the shape for the maximum L/D configuration at supersonic condition. The shape for the maximum L/D at subsonic condition obtains additional lift at the kink compared with the minimum drag configuration. It leads to enhancement of L/D.

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An existing computer code named super charring materials ablation (SCMA) is updated by implementing a coking process, in which a pyrolysis gas cokes within a char layer. The conservation equations for masses for resin, and coke, and energy are given. A coking rate equation is calculated by accounting for mass conservation of carbon deposited within a char layer. The method so upgraded is applied to the post-test sample analysis of the in-depth density profile for a low density ablator heated in an arcjet wind tunnel under one operating condition to evaluate the temperature variation of mass fraction of carbon contained in a pyrolysis gas used in the coking rate equation. The calculated result between with and without coking is compared to discuss the effect of coking on ablation behaviors.

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The application of the Knudsen pump to the thruster for the orbital control system is considered. This is a kind of the “fuel-free” type thruster using the ambient rarefied gas around a spacecraft as the propellant. The thrust is produced due to the thermal transpiration flow in the pump, which is the characteristic flow in the rarefied regime. In this paper, the basic behavior of the Knudsen pump, which operates in a high speed and highly rarefied gas flow, is numerically investigated using the direct simulation Monte Carlo (DSMC) method to confirm its efficiency as a thruster. The dependency of the skin friction on the Knudsen number, the velocity of the inflow and the temperature distribution on the channel wall is clarified. It is found that a tendency of the variation of the drag reduction rate with the increase of the Knudsen number in the case of the inflow velocity at 1 km/s is different from that at 8 km/s.

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In this paper, flexible aeroshell shape optimization method for aerocapture mission is formulated using particle based membrane model and GA with response surface modeling. And the calculation is conducted considering various objective function. As a result, it is shown that proposed shape optimization method for flexible aeroshell vehicle works properly, and thus gives reasonable solutions. In this analysis, three clusters of solution sets are obtained; 1) solutions which minimized fuel mass fraction, 2) solutions which minimized both TPS mass fraction and outer frame mass fraction, 3) solutions which minimized ballistic coefficient. Cluster 1) could reduce the total mass fraction most in three kinds of clusters. It shows that fuel mass is much more sensitive to total mass than any other required mass for assumed entry condition. And also, tradeoff relation ship between minimization of fuel mass fraction and minimization of outer frame mass fraction.

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We analyze the asymmetric vortices in the flow over a slender body at high angle of attack by numerical simulations aiming a proportional control of the side forces generated by vortices with a device such as dielectric barrier discharge (DBD) plasma actuator. With regard to the computational method, Reynolds averaged Navier Stokes/large-eddy simulation hybrid method is adopted with high-order compact spatial difference scheme. The grid convergence analysis is firstly conducted and the results show that the computational grid adopted in this study is fine enough for qualitative discussion. The total number of the grid point is 411 million points. Then, the effects of bump height on flow fields and aerodynamic characteristics are discussed. Note that bump is added near the body apex to simulate the symmetry-breaking imperfection. As a higher bump is adopted, stronger asymmetry is observed in the flow fields. On the other hand, side-force has nonlinearity with the bump height.

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Asymmetric separation vortices over a slender body at a high angle of attack exert a strong side force on the body and lead to the loss of attitude stability. We investigated the active control of the separation flow over a slender body and addressed the proportional control of the side force and the pitching moment. A flow control experiment was conducted in a wind tunnel using a cone-cylinder test body and a Dielectric Barrier Discharge (DBD) plasma actuator as a flow control device. The free-stream velocity was 9 m/s and the Reynolds number was approximately 42000. The side force coefficient was proportionally controlled within approximately ±1.0 using the actuator at the aft body, and the static stability angle of attack was controlled from 25 to 40 degrees and 65 to 85 degrees by controlling the pitching moment when the center of gravity was at the 55% position from the body apex. We estimated that a higher actuator output power is required for the effective control of the aerodynamics in a real flight. In addition, we confirmed that the actuator burst operation mode could reduce the required output power.

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Liquid Droplet Radiator (LDR) is an important candidate for disposing large quantities of waste heat more than 1 MW which will be handled by large space structures such as Space Power Satellite. The working fluid is heated through a heat exchanger by the waste heat generated in a large structure in space. Then, the working fluid is emitted in space through nozzles of the droplet generator toward a droplet collector as multiple streams of droplets. During the flight in space, the droplets lose thermal energy via radiative heat transfer. After the cooled droplets are captured by the droplet collector, the working fluid is recycled to the heat exchanger by a circulating pump. The automatic control system on the circulation of working fluid in a liquid droplet radiator has been built using a programmable logic controller. The proportional control of flow rate with the term of the variation of the counter flow in the gear pump and the relaxation of the change of an target flow rate has succeeded within 5 percent at the automatic circulation control of the working fluid from 100 ml/min to 200 ml/min and from 200 ml/min to 100 ml/min.

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Recently the activities for the development of micro- and nano- satellites like a cubesat are being increased. A lot of satellites are operated at a ground station located in a university, but public orbit information is used. In this paper, the orbit determination system is constructed using observations of Doppler frequency at a low-cost ground station. The traditional batch state estimation filter is used, and the orbital elements of satellite position and velocity in inertia coordinates are determined from the Doppler frequency of receiving signals. The verification tests are conducted using UHF-band signals (about 430 MHz) received at an amateur radio station and S-band signals (about 2.2 GHz) received at a 2.4-m parabola antenna from real satellites, and the valid performance is conformed for tracking satellites not depending on public orbit information. The error is max. 0.3 degrees in direction and max. 3.9 km in position for UHF-band signals, and max. 1.1 degrees in direction and max. 37 km in position for S-band signals. This method is valid especially for low-earth-orbit satellites with large Doppler effect.

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We are in progress to develop a system for automatic operation of a satellite in order to reduce human load at satellite steady operation phase. The ground station for small satellite REIMEI (INDEX : INnovative-technology Demonstration EXperiment) is used as a test bench for verification of the proposed method. In our new automatic operation system, a scheduler software as a substitutive operator manages all the operations through a unified procedure, including sending command, receiving telemetry, and driving antenna in accordance with an operation time line which is prepared before the operation pass. The scheduler also performs diagnostics of satellite anomaly based upon the received telemetry data and status of the ground station. In case that some anomaly of the satellite is detected, the scheduler initiates an emergency schedule that was prepared depending on the emergency level. The automatic operation system is nearly completed for downlink operations of the data recorder that account for 75% of REIMEI steady operation. This approach is very effective to reduce psychological and physical load of operators.

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As a high-speed manned transportation system in the future, point-to-point (P2P) suborbital rocketplane is currently studied in Japan Aerospace Exploration Agency (JAXA) space transportation mission directorate. The vehicle was designed on the basis of the concept of waverider allowing high L/D in hypersonic regime, which is required for longer flight range and smaller load factor. Compared with an ideal waverider, the designed P2P suborbital rocketplane has outer wings for the improvement of the low-speed aerodynamic performance, finite thickness in the leading edge for the reduction of the aerodynamic heating, and twin vertical tails for directional stability. The aerodynamic performance of the P2P vehicle was investigated through numerical simulation of both subsonic and hypersonic flows, and the baseline aerodynamic shape of the P2P vehicle was discussed. The L/D in the trim condition at hypersonic speed was 2.6.

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Recently, the suborbital plane has attracted attention for space tourism. However, the suborbital plane has some aerodynamic problems such as severe aerodynamic heating and deceleration G from aerodynamic forces. To solve these problems, a high-drag-force device, like a parachute, is thought to be effective, since high-drag-coefficient spacecraft can reduce the velocity at an early stage of atmospheric re-entry to the Earth. In this study, we investigated experimentally the aerodynamic characteristics of the parachute for the suborbital plane as the high-drag-force device by using three kinds of model in a hypersonic wind tunnel. From this experiment, the following results were obtained: 1) the drag force increases drastically when a parachute is used, 2) the effect of increasing drag force depends on the strength of the shockwave which is generated from the suborbital plane, 3) the shape of the parachute shows a periodic pattern, and 4) the parachute may have more than two mechanically stable shapes. These results suggest that the hypersonic parachute system is useful as a high-drag device for suborbital planes from the viewpoint of increasing drag coefficient.

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Temperature Gradient Growth experiments of In_xGa_1-xSb will be performed on the International Space Station (ISS) in 2012. In the GaSb/InSb/GaSb-sandwich system used, before growth, the dissolution process of GaSb into InSb takes place during the formation of the growth solution (melt). Solutal mass transport occurring during dissolution plays a significant role. In crystal growth on Earth, the large difference between the densities of InSb and GaSb leads to gravitational segregation in the melt and axial non-uniformity in crystal composition. In addition, the large separation between the liquidus and solidus curves in the phase diagram of this system further contributes to compositional non-uniformity. In order to have a better understanding for the effect of gravity on solutal transport during in this system, the dissolution process was numerically simulated under both Earth’s gravity and a microgravity level on the ISS. Numerical simulations showed that under Earth’s gravity, dissolution of the GaSb seed was enhanced due to the contribution of solutal natural convection. However, under microgravity diffusion mass transport was dominant in the melt, and the contribution of natural convection was not significant.

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The downward flame spread rate and limiting oxygen index (LOI) were investigated with varying oxygen level in normal gravity. Four kinds of diluents, nitrogen, argon, helium and carbon dioxide were used to change the thermal properties of the ambient gas. In order to clarify the impact of the properties on spread rate, simple scale analysis was developed for two-dimensional flame spread. The scale analysis predicted strong dependence of LOI on the flame temperature and the predicted LOIs agreed with the measured LOIs; the Damkohler number is the important factor for LOI of the downward spread. In the microgravity experiment, the flame spread rate was measured with varying opposed flow velocity at 18% oxygen level. The faster flame spread than that in normal gravity condition was observed when the opposed flow was about 15cm/s, which was consistent with the regime map predicted by the scale analysis.

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A numerical simulation study was carried out to investigate the effects of thermal and solutal Marangoni convections (thermosolutal Marangoni convection) on transport structures in a liquid bridge under zero gravity. The liquid bridge in the model represents a three dimensional half-zone configuration of the Floating Zone (FZ) growth system. Three dimensional field equations of the liquid zone, i.e. continuity, momentum, energy, and diffusion equations, were solved by the PISO algorithm. Computations were performed using the open source software OpenFOAM. The numerical simulation results show that the flow field becomes three-dimensional and time-depended when the solutal Marangoni number is larger than the critical value. It was also shown that not only flow patterns but also the azimuthal wave number (m) changes due to the competing contributions of thermal and solutal Marangoni convective flows.

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As humanity's activities expand to the Moon, Mars, and other extra-terrestrial bodies, it will be necessary to use local resources rather than bringing everything from the Earth. This concept is called In-Situ Resource Utilization (ISRU), which starts with excavation and earthmoving. The present study focuses on loosening and moving of the lunar regolith by a ripper (or rake) and a wide blade. After characterizing the lunar regolith and two of its simulants (JSC-1A), the relationship between the excavation energy and different conditions, namely, relative density— a ratio describing the density of a material with respect to its maximum and minimum density —, and tine spacing on a rake, is investigated with scaled experiments. Prior ripping decreases total excavation energy by up to 20% if the relative density is > 60%. This study has proven that the characteristics of the lunar regolith, i.e. angular grains and high relative density, led to the effectiveness of ripping.

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This paper proposes a concept of a center-actuated walking robot whose design is suitable for use of the parts fabricated from in-situ resources on the Moon, Mars, and other planetary bodies. The core idea of this concept is concentrating any actuator or movable components at one section of the robot body so that the rest of the body can be manufactured and assembled with limited resources and capabilities that are expected in the early stage of manufacturing away from the Earth. Justification of this concept is made from the economicalpoint of view as well as other aspects. As a walking mechanism, the proposed one consists of two rigid structures, which serve as legs, interconnected by one 6 degree-of-freedom actuator module (e.g. Gough-Stewart mechanism). It can be classified as a type of the walking mechanism with reduced degrees-of-freedom, which significantly reduces mechanical complexity, and thus makes a walking planetary rover a realistic option. Simulation results proved that a tetrapod based on the proposed concept can employ both static and dynamic gaits, such as static walk, trot, and pace. Potential of the concept, for example, reusability as man-made in-situ resources, is explored and future research topics are identified. Designers of space robots near future would be required to take advantage of in-situ resources. The work presented here is aimed at pioneering the robot design in the era of extraterrestrial resources.

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Thermocapillary-driven flow is realized in a liquid bridge by a non-uniform surface tension distribution caused by an axial temperature difference. It is known that a unique behavior of the particles suspended in the liquid bridge emerges after the onset of the oscillatory flow states, which is named the particle accumulation structure (PAS) after Schwabe et al (Microgravity Sci. Technol. 1996). In the present study, we focus on one of the two types of the PAS, SL-2 PAS (after Tanaka et al. (Phys. Fluids 2006)), appeared under larger thermocapillary effect. Especially we pay our special attention to the effect of ambient gas flow around the liquid bridge on the SL-2 PAS. We reconstruct the spatio-temporal variations of the trajectory and velocity of the particles on the SL-2 PAS by applying three-dimensional particle tracking velocimetry (3D PTV).

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In this paper, a new algorithm that measures the centroid position of a spot in the image obtained from an image sensor with sub-pixel accuracy was presented. The lens forms a spot on the image sensor, which spreads over a few pixels on the image. The algorithm is used to carry out centroid calculation using the brightness values of the spread pixels, and it increases the accuracy of the center detection of the spot image. In the calculation, a weight table is used, which has a circular form with a blurred border. The weight value is multiplied by the corresponding brightness value, which further increases the accuracy. An image containing a spot and random noise is produced in the simulation. Centroid calculation is carried out using the produced image, and the accuracy of the centroid position is evaluated. It is confirmed that the presented algorithm can detect the centroid position with an accuracy of 0.02 = 1/50 pixel when the spot is defocused. It is 1.5 times more accurate comparing with one of the traditional algorithms and the presented algorithm proves to be quite effective.

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Space communications by lightwave have various advantages such as smaller and lighter equipment, higher data rate, larger communication capacity, and limited risk of interference with other optical communication systems. Many organizations in the world have been researching and developing the optical space communications technologies for space infrastructures to support various space activities. At the National Institute of Information and Communications Technology (NICT) in Koganei, Japan, a laser communication demonstration between the NICT optical ground station and a low earth orbit satellite was successfully conducted in 2006 to 2009. The polarization characteristics of an artificial laser source in space were measured through space-to-ground atmospheric transmission paths. Stokes parameters and the degree of polarization of the laser beam transmitted from the satellite were measured. These results contribute to the link estimation for quantum key distribution (QKD) via space and provide the potential for enhancements in QKD worldwide in the future. After this experiment, NICT has begun to develop a Small Optical TrAnsponder (SOTA) onboard a small satellite, which project is called the Space Optical Communications Research Advanced TEchnology Satellite (SOCRATES). In this mission, basic QKD experiments are planned. In this paper, the result of the field experiment between two buildings and the verification plan of basic QKD experiments is presented.

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The First Quasi-Zenith Satellite "MICHIBIKI" was successfully launched on September 11, 2010. After the initial functional verification, we have started the domestic experiments to evaluate performance of the Quasi-Zenith Satellite System(QZSS), especially the effect of QZSS on the improvement of availability for GPS Positioning, Navigation and Timing (PNT) service. “MICHIBIKI Field Observation on Multi-Locations in Various Conditions” is one of technical demonstration activities, started after the initial verification of “MICHIBIKI”. The purpose of the experiments is to show the “MICHIBIKI” capability to enhance GPS availability. The experiment has been conducted in several different conditions with statistical manner. There are two approaches for the experiment. One is “Multi-locations” aiming to collect as much as data to evaluate statistically and the other is “Various conditions” aiming to evaluate the effect of surrounding environments as well as different use cases. By using multi-frequencies high-end GNSS receiver and handheld GPS logger, Fixed Point Observation and Moving Point Observation have been conducted all over Japan in collaboration with lots of partners. Analysis results of post-processing show the improvement positioning availability by “MICHIBIKI”. In this paper, we will report the overview of the experiment and current analysis result of observed data.

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This paper describes an evaluation of a 16 Amplitude-Phase-Shift-Keying (16APSK) direct Radio Frequency (RF) signal processing transceiver designed to make reconfigurable communication equipment smaller and more efficient. An experiment board made for the evaluation mainly comprised Field-Programmable Gate Array (FPGAs) for modulation and demodulation, an A/D converter, and a D/A converter. Evaluation of 750 Mbps and 16APSK confirmed that making the transceiver’s Intermediate Frequency (IF) signal processing part less complex enables device scale to be substantially reduced while keeping performance equal to that obtained with conventional digital communication technology. To reduce the number of conversion steps and obtain our destination frequency in the L band, the modulation portion uses a bandpass filter (BPF) to select an aliasing component with a 1.5 GHz sampling frequency. In the same way, the demodulation portion uses direct under-sampling technology to sample the L-band using a signal with a 1.5GHz sampling frequency. An amplitude equalizer inside the transceiver avoids the bit error rate (BER) degradation caused by linear amplitude distortion at the aliasing frequency range of the sampling frequency response in the L band. As a result we succeeded keeping good performance compared with the IF band loopback test using no conversion.

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The lunar explorer SELENE (also called KAGUYA) carried thirteen scientific mission instruments to reveal the origin and evolution of Moon and to investigate the possible future utilization of Moon. In addition to the scientific instruments, a high-definition TV (HDTV) camera provided by the Japan Broadcasting Corporation (NHK) was carried on KAGUYA to promote public outreach. We usually use housekeeping telemetry data to derive the satellite attitude along with orbital determination and propagated information. However, it takes time to derive this information, since orbital determination and propagation calculation require the use of the orbital model. When a malfunction of the KAGUYA reaction wheel occurred, we could not have correct attitude information. This means that we don’t have a correct orbital determination in timely fashion. However, when we checked HDTV movies, we found that horizon information on the lunar surface derived from HDTV moving images as a horizon sensor was very useful for the detection of the attitude of KAGUYA. We then compared this information with the attitude information derived from orbital telemetry to validate the accuracy of the HDTV derived estimation. As a result of this comparison, there are good pitch attitude estimation using HDTV derived estimation and we could estimate the pitch angle change during the KAGUYA mission operation simplify and quickly. In this study, we show the usefulness of this HDTV camera as a horizon sensor.

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The Japanese lunar orbiter SELENE discovered three giant holes, Marius Hills Hole, Mare Tranquillitatis Hole, and Mare Ingenii Hole, that exceed tens of meters in diameter and depth, and that are possible lava tube skylights. These lunar-hole structures and possibly underlying lava tubes have great significance for lunar science and potential utilization, and thus should be priority targets of future lunar exploration. It is essential to acquire detailed information on terrain around the holes, multi-layered outcrops on the inner walls of the holes, debris and thermal conditions on the floor of the holes, and entrances of predicted lava tubes extending from the bottoms of the holes. The Smart Lander for Investigating the Moon (SLIM) mission, a future Japanese lunar lander mission being considered, is planned to land in an area within 100m around the Marius Hills Hole. Rovers will be dispatched from SLIM and approach the hole. SLIM and its rovers are small but the mission return could be huge.

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A Mars Aero-flyby Sample Collection (MASC) mission has been proposed in a Mars exploration project at Japan Aerospace Exploration Agency (JAXA). The MASC vehicle enters the Martian atmosphere, captures dust particles and atmospheric gases at sampling altitudes between 30 and 50 km, and returns back to Earth. In order to improve the feasibility of this project, the development of its sampling system during flying in the Martian dusty atmosphere is crucial. Since silica aerogel has been used as a capturing medium for micrometeoroids and space debris, it is also planned to be used for the MASC mission. However, the capture of hypervelocity micron-size dust particles during the Martian atmospheric flight using aerogel is challenging. The aerogel is exposed to significant aerodynamic heating during sampling, and thus, the effect of heated aerogel on the dust particles must be evaluated. This work attempts to evaluate the impact on aerogel and the survivability of the dust particles inside the capturing medium by carrying out light gas gun and Van de Graaff experimental tests. By comparing the cases between normal and heated aerogels, the survivability of the dust samples as well as the heating effect has been investigated.

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WISE-CAPS (Web-based Integrated Secure Environment for Collaborative Analysis of Planetary Science) is Web-GIS based platform for analysis and sharing of lunar and planetary exploration data. As this system is entirely web-based, users do not need to install any special software to use WISE-CAPS (only web browsers required). The WISE-CAPS is constructed fully using open source software and is also open standard compliant. This means this system can exchange data as layers, both importing and exporting. The system has free scale zooming (in and out) without reloading web pages, point designation and several advanced features. Using this platform, users can put their data and share with other people. Another unique feature is web-server based user control function which can limit data browsing to specific individuals (ID) and groups. This paper describes outline of WISE-CAPS and some examples of its current implementation, and future prospection.

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Variations of Mars' rotation provide us information concerning both the interior structure and the surface mass redistribution of Mars. Precession and nutation of Mars mainly reveal the core-mantle subsystem, besides length-of-day variation and polar motion of Mars are generally referred to the atmosphere-cryosphere sub-system. As one of the missions of MELOS (Mars Exploration with Lander-Orbiter Synergy), we are proposing areodetic observations using space geodetic techniques including four-way Doppler measurement (FWDP) and inverse VLBI (IVLBI). FWDP is ranging rate measurement of target spacecraft via a relay spacecraft. Utilizing the heritage of four-way Doppler measurements by SELENE, we plan to track the MELOS Landers relayed by the MELOS Orbiter. We also introduce the new technology of IVLBI. In the framework of this technology, one ground radio telescope observes the phase-shift between mutually coherent signals on the orbiter and the landers. The estimated accuracy for the rotation obtained by IVLBI is better than 1 mas (milli-arc second) including the systematic phase noise. The measurements will provide the core radius estimation with the errors of less than 200 km by 50 weeks observations with two landers.

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