Although the surface is covered with rocks and is apparently lack of fine regolith, Itokawa’s surface show darkening and reddening by space weathering. Experimental results suggest rocky meteorite fragments can be weathered. The presence of opposition effect in rocky terrain of Itokawa suggested that the surface would be covered by particulate materials or porous enough to scatter light. Itokawa-derived fine grains (some of which have nanophase iron) in Hayabusa sample capsule clearly indicate the presence of fine particles on the surface of Itokawa and their effect on the optical property of the asteroid.

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Remote sensing observations with high resolution and high signal-to-noise ratio require a large-aperture optical system. To achieve these requirements, we propose an adaptive optics system that compensates the wavefront aberration due to misalignment or thermal deformation of optical elements. We use an image-based wavefront sensing technique to realize simple hardware architecture and high signal-to-noise ratio. For image-based sensing, a priori information is required in addition to the acquired images, which was achieved in our case using a phase diversity (PD) wavefront sensing method. We apply PD using a liquid crystal on silicon spatial light modulator (LCOS-SLM). For remote sensing in which the observed scene is moving relative to the observer, we modify the conventional PD method to estimate the parallel shift between the obtained images. We construct an adaptive optics system testbed using LCOS-SLM and a CMOS camera. The laboratory test results show that the proposed system improves the optical performance of remote-sensing sensors.

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Living organisms are always exposed and respond to various stresses from the external world - the environment, such as temperature, gravity, radiation or circadian. Understanding and addressing the effects of the environment on human health is a major issue. To clarify this issue, we are studying both human and medaka. Medaka is a good model organism to verify the effects of the space environment for their features including their short life cycle, productivity, mapped whole genome sequences, and transparency even in adults. We are trying to evaluate the heartbeat using transparent medaka with live imaging. Despite the importance of functions of internal organs, its observation has been difficult. Imaging technologies has developed remarkably in a decade. With the retirement of the space shuttle, experimental procedures without sample collection will be required. We established an approach to acquire the images and the methodology of the analysis from high-speed movies. We are subsequently going to introduce a method of verifying the movement of internal organs in medaka experiments using high-speed movies.

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In order to walk upright on the Moon or Mars without falling, a specific walking strategy to account for altered gravitational conditions must be verified. We have therefore been studying changes in the kinematics of walking at different gravitational loads using a body weight suspension system. Our simulation consisted of three gravitational conditions: 1 g (Earth); 1/3 g (Mars); and 1/6 g (the Moon). Surface EMG recordings were taken from the leg muscles of subjects walking on a treadmill. Cadence, stance phase duration, and step length were calculated from the walking velocity and steps. Subsequent experiments revealed that muscle activity and the duration of the double support phase decreased as simulated gravity was reduced. These changes are apparently caused not only by the direct effects of unloading but also by kinematic adaptations to the same. It can be said that humans walk slowly with a shortened stride and elongated stance phase in order to adjust to low gravitational conditions. One major limitation of our study that may have affected walking stability was the fact that the suspension system was fixed to an immovable frame. We have begun further studies using a newer movable body weight suspension system to achieve more realistic simulations.

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Mission-critical components of unmanned spacecraft, such as electronic equipment, are often mounted on the interior surfaces of structure panels. This study investigated debris impact damage to structure panels to assess the degree to which they can protect components. If debris perforates a structure panel but is stopped by the equipment chassis, the impact will not affect the probability of mission success. The ballistic limit of the chassis therefore equals to the damage limit of the structure panel. To estimate this damage limit, hypervelocity impact experiments were conducted on sets of a honeycomb sandwich panel, simulating a structure panel, fixed directly to an aluminum alloy plate, simulating an equipment chassis, with no gap between the two. Debris environment models show that alumina debris particles smaller than 1 mm in diameter are dominant in low earth orbit, and the average impact velocity is over 10 km/sec. However, advanced techniques are required to accelerate small solid projectiles to such speeds, so steel projectiles at 6 km/sec were used to simulate the impact pressure caused by alumina impacting at 9 km/sec. The depths of the resulting impact craters on the chassis plates were measured with an optical microscope, and the damage limit equation of the structure panels was derived from the crater depths. The calculated damage limit equation was compared with the SRL ballistic limit equations. As a result, it was found that the equation obtained in this study showed safety results but was too robust. The stand-off distance between honeycomb sandwich panel and aluminum alloy plate was effective to decrease depths of craters in the plate.

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As an effective means of suppressing space debris growth, the Aerospace Research and Development Directorate of the Japan Aerospace Exploration Agency (JAXA) has been investigating an active space debris removal system that employs highly efficient electrodynamic tether (EDT) technology for orbital transfer. This study investigates tether deployment dynamics by means of on-ground experiments and numerical simulations of an electrodynamic tether system. Some key parameters used in the numerical simulations, such as the elastic modulus and damping ratio of the tether, the spring constant of the coiling of the tether, and deployment friction, must be estimated, and various experiments are conducted to determine these values. As a result, the following values were obtained: The elastic modulus of the tether was 40 GPa, and the damping ratio of the tether was 0.02. The spring constant and the damping ratio of the tether coiling were 10^-4 N/m and 0.025 respectively. The deployment friction was 0.038ν + 0.005 N. In numerical simulations using a multiple mass tether model, tethers with lengths of several kilometers are deployed and the attitude dynamics of satellites attached to the end of the tether and tether libration are calculated. As a result, the simulations confirmed successful deployment of the tether with a length of 500 m using the electrodynamic tether system.

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This paper analyzes the behavior of an interplanetary dust particle on a horseshoe orbit by solving equations of the circular restricted three-body problem (CR3BP) analytically to predict the distribution of dust particles around the Earth’s orbit, because in previous research numerical simulation showed that dust particles trapped on a horseshoe orbit contributes a number density of dust particles around the Earth’s orbit. An analytical solution is obtained by coordinate transformation and linearization of external forces on the assumption that a horseshoe orbit is an oscillation centered at Lagrange point L3. This paper describes properties of the horseshoe solution and also shows that the analytical solution of the CR3BP is consistent with that of numerical integration.

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A concept for forced reentry of small orbital debris with characteristic dimension ~ 10 cm or less from the highly populated sun synchronous orbit by injecting micron scale dust grains to artificially enhance drag is discussed. The drag enhancement is most effective when dust grains counter rotate with respect to the debris resulting in hypervelocity dust/debris impacts. While the natural drag on small debris with ballistic coefficient ~ 5 kg/m² in orbits with perigee above 900 km is negligible, it is sufficient to decay the orbit of the injected dust grains at a significant rate. This offers a unique opportunity to synchronize the rates of descent of the dust and debris to create a sweeping (snow-plow-like) effect on the debris by a descending narrow dust layer. The dust density necessary to de-orbit small debris is sufficiently low such that the orbits of active satellites which have larger ballistic coefficients are minimally affected. If deemed necessary, contact of the injected dust with active satellites may be avoided by maneuvering the satellite around the narrow dust layer.

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JAXA is developing analysis methods which can detect faint geostationary (GEO) objects that are not in the catalog provided by U.S. The stacking method, which uses numerous CCD frames to detect objects below the background noise level, has been developed and has been shown to work well, but has the drawback that detecting objects whose movements are unknown is extremely time-consuming. To overcome this, a new algorithm is developed which uses binarization of CCD images and calculates sum values instead of medians. When implemented in a field programmable gate array (FPGA), the algorithm reduces the analysis time by a factor of a thousand over the original stacking method, which enables us to analyze one night’s data before the following night’s observation. Another fast analysis method is a line-identifying technique which finds straight-line candidates in time-sequential CCD frames and can also detect faint GEO objects. Although the stacking method can detect fainter objects, the line-identifying technique has much shorter analysis time and can detect 40 cm-size objects in GEO using a 35 cm telescope. These two analysis methods may contribute to detecting many un-cataloged objects in the near future.

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Spacecraft is one of the largest and most complex systems of artificial products so that it necessitates application of systems engineering methods in its development process. Spacecraft development has a special difficulty because of its operations in low gravity and high vacuum environment, which makes system level tests impossible on the earth before launch. For this reason, it is common that design and manufacture are complementarily verified only by analysis. However, insufficient verification often results in on-orbit nonconformance, leading to unexpected anomalies in mission operations. This study is to propose to use Dual Vee Model developed in the recent advancement of strategic systems engineering methodology for identifying the root cause of such anomalies. Usefulness of this approach has been demonstrated by a real case of the communication interference, called multi-path, experienced on the Advanced Earth Observing Satellite-II (ADEOS-II) spacecraft. The Dual Vee Model consisting of the Architecture Vee and the Entity Vee turns out helpful to express the configuration items and the development sequence and to identify the process which could have caused the anomaly.

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Satellites are very complex systems and due to their knotty constructions, the servicer may not access the malfunctioning component without a long sequence of disassembly processes. One preferred method is to design and replace the components in modules for easy repairs. Having a universal design layout on satellite design would leverage an efficient maintenance service. However, since most satellites have unique designs which best suit their specific missions, the modularization of satellite systems haven't been widely adopted due to the lack of compatibility. This research aims to propose a design rule for the 'Easily serviceable satellite' in order to construct an assembly design guideline. Our goal is to optimize the configuration of satellite subsystems for easier service while maintaining their unique characteristics and requirements for their missions. These optimally rearranged subsystems can be either grouped or integrated to several service modules.

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The authors have been developing CAMUI (Cascaded Multistage Impinging-jet) type hybrid rockets, explosive-flee small rocket motors. This is to downsize the scale of suborbital flight experiments on space related technology development. A key idea is a new fuel grain design to increase gasification rates of a solid fuels. By the new fuel grain design, the combustion gas repeatedly impinges on fuel surfaces to hasten the heat transfer to the fuel. Suborbital flight experiments by sounding rockets provide variety of test beds to accumulate basic technologies common to the next step of space development in Japan. By using hybrid rockets one can take the cost advantage of small-scale rocket experiments. This cost advantage improves robustness of space technology development projects by dispersion of risk.

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The most significant problem of a conventional hybrid rocket is its very low fuel regression rate; however, paraffin-based fuels are known to have a much higher regression rate. It was determined by Karabeyoglu et. al. that the materials with low surface tension and melt layer viscosity, such as paraffin-based fuels, generate more entrainment of liquid droplets, resulting in increased regression rates. It was very interesting model for us. Therefore we confirmed it experimentally by observing the burning fuels with a visualization chamber and revealed the high regression rate mechanism of paraffin-based fuels. And then we tried to increase the regression rate moreover with swirled oxidizer flow. It was quite effective, approximately twice of the regression rate of no swirl, but a large amount of liquid fuel was observed in the chamber. It is necessary to vaporize the liquid fuel rapidly to obtain high combustion efficiency. We tried to resolve this problem with a baffle plate set in the combustion chamber and obtained 3% combustion efficiency improvement.

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The Venus explorer, Akatsuki, was launch on 20 May 2010. After 200-day journey through the interplanetary transfer orbit, it reached the Venus at the altitude of 550 km on 7 Dec 2010. However, it experienced a trouble of the explorer's propulsion system and was not able to be the Venus orbiter. It now orbits the Sun with the period of 203 days. In this paper, we discuss the trajectory design strategies for Akatsuki mission by introducing the constraints which come from the observation orbit and the spacecraft system. The details of planning and the results of orbital maneuvers are also shown in this paper.

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Akatsuki is the Japanese first Venus exploration program. The spacecraft was successfully launched in May 21st 2010 from Tanegashima Space Center via H II-A vehicle. One of her missions is a flight demonstration of her telecommunication system developed for supporting Japanese future deep space missions. During her half-year cruising phase heading for Venus, we had conducted checkouts of the system. The key technologies are a deep space transponder, a set of onboard antennas, and power amplifiers. They all were newly introduced into Akatsuki. Their performances had been tested through the half-year operations and been investigated by using collective data obtained in the experiments. Our regenerative ranging was also demonstrated in these activities. We proved that the newly introduced telecommunication system worked exactly as we had designed and that the system performances were the same as evaluated on the ground. In this paper, we will summarize these in-flight validations for the Akatsuki telecommunication system.

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This paper shows the attitude operation results of Japanese interplanetary solar sail demonstration spacecraft IKAROS. IKAROS was launched on 21 May 2010(JST) aboard an H-IIA rocket, together with the AKATSUKI Venus climate orbiter. As IKAROS is the secondary payload, the development cost and period were restricted and the onboard attitude system is very simple. This paper introduces the attitude determination and control system. And as IKAROS is spin type spacecraft and it has the large membrane, the attitude control is not easy and it is very important to determine the long-term attitude plan in advance. This paper also shows the outline of the IKAROS attitude operation plan and its operation results.

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In this paper, the world's first application of hybrid rockets to rocket sled system is described. By using N₂O oxidizer which has self-pressurizing capability, simple handling and low-cost operation was enabled. On the other hand, accurate prediction of the time-velocity profile of the rocket sled becomes difficult because the thrust curve of the N₂O hybrid rocket varies with the outside temperature. Based on the preliminary thrust measurement, we predicted the velocity profile of the rocket sled and compared to the actual sled run. Both were well agreed with an error of 10 % in velocity.

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The long-term development of ceramic rocket engine thrust chambers at the German Aerospace Center(DLR) currently leads to designs of self-sustaining, transpiration-cooled, fiber-reinforced ceramic rocket engine chamber structures.This paper discusses characteristic issues and potential benefits introduced by this technology. Achievable benefits are the reduction of weight and manufacturing cost, as well as an increased reliability and higher lifetime due to thermal cycle stability.Experiments with porous Ceramic Matrix Composite(CMC) materials for rocket engine chamber walls have been conducted at the DLR since the end of the 1990s.This paper discusses the current status of DLR's ceramic thrust chamber technology and potential applications for high thrust engines.The manufacturing process and the design concept are explained.The impact of variations of engine parameters(chamber pressure and diam-eter)on the required coolant mass flow are discussed.Due to favorable scaling effects a high thrust application utilizes all benefits of the discussed technology, while avoiding the most significant performance drawbacks.

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The μ10 cathode-less electron cyclotron resonance ion engines, have propelled the Hayabusa asteroid explorer for seven years since its launch in May 2003. The spacecraft was focused on demonstrating the technology needed for a sample return from an asteroid, using electric propulsion, optical navigation, material sampling in a zero gravity field, and direct re-entry from a heliocentric orbit. The final stage of the return cruise and the subsequent trajectory correction maneuvers have been accomplished by using a special combined operation of neutralizer A and ion source B after the exhaustion of the other neutralizers' lives by the autumn of 2009. The total duration of the powered spaceflight was 25,590 h, which provided a delta-V of approximately 2.2 km/s and a total impulse of 1 MN·s. The degradation trends of the thruster performances have been investigated. It seems that the main cause of the degradation was the decrease in effective microwave power input to the discharge plasma induced by the increase in the transmission loss of the microwave feed system, and not due to the increase in the gas leakage through the accelerator grid apertures enlarged by erosion. Unintentional engine stop events have been summarized and analyzed. Most of them occurred due to the limit check errors of the backward microwave powers. Such errors can be decreased by carefully monitoring the trend change in microwave backward power as a function of xenon flow rate in future missions.

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For the application of Hall thrusters to small satellites, we have been developing a low power Hall thruster. As an initial test, we evaluated the thrust performance of a 50 W class miniature Hall thruster developed at Kyushu University. The outer diameter and the length of the acceleration channel are 18 mm and 7 mm, respectively. A torsional type thrust stand was developed, since the estimated thrust is 1-3 mN, which is too small to be measured by means of a conventional pendulum type thrust stand for Hall thrusts. The uncertainty of this thrust stand is less than 10% at 2 mN with calibration. The thrust, the discharge current, specific impulse and the thrust efficiency at xenon mass flow rate of 0.30 mg/s and discharge voltage of 125 V were 1.7 mN, 0.35 A, 600 sec and 11.4%, respectively. The thrust increased with an increase in discharge voltage and the thrust became 3.1 mN at discharge voltage of 250 V.

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Magnetic Sail, which generates thrust by utilizing the interaction between the artificial magnetic field and the solar wind plasma stream, is next generation propulsion for deep space exploration. In this study, thrust of Magnetic Sail scale model, which correspond to 10 N-class Magnetic Sail in space, for various tilt angles were measured with the parallelogram-pendulum thrust stand. The magnetospheric size L=0.1 m (60 Am² of magnetic moment) of Magnetic Sail was demonstrated and measured thrust for tilt angle 0 degrees (dipole-moment perpendicular to the plasma flow) was about 1.2 N which correspond to 10 N-class Magnetic Sail in space. Thrust increase with increasing of tilt angles and was about 1.5 times as large as that of 0 degrees case in the case of 90 degrees.

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In order to realize a long lifetime of an electric propulsion system, we have been investigating various electrodeless electric propulsion concepts utilizing a helicon plasma source. In one of our concepts, helicon plasma is electromagnetically accelerated using a rotating electric field in the presence of a diverging static magnetic field. This acceleration concept is called the Lissajous acceleration. Plasma acceleration experiments have been conducted and plasma acceleration was evaluated using a Mach probe. Although the experiments showed some features of the electromagnetic acceleration, most increment of the plasma velocity is caused by the increment of the electron temperature. The thrust (4.95 μN) did not reach feasible values for real applications, and therefore, it is important to find a better operational condition with the aid of a theoretical thrust model. We have developed a theoretical thrust model which consists of a trajectory analysis and an electric field penetration model in the electrostatic approximation. The model shows that experimental parameters are off from an optimum operational condition which provides the maximum thrust.

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Microfabrication of a 3 x 3 micro-plasmajet array with ultra-violet lasers and their thrust performance tests were conducted for nozzle elements with exit height of 0.5 mm and length of 0.5 mm. Stable discharge and operational conditions were confirmed for the 3 x 3 micro-plasmajet array thruster. To evaluate thrust characteristics of the arrayed plasmajet, thrust characteristics of the thruster were compared with those of the micro-single-nozzle plasmajet. It was shown that the thrust and specific impulse of the micro-plasmajet array were higher than those of the micro-single-nozzle plasmajet due to the multi-jet effect. The typical values of the thrust, specific impulse and thrust efficiency averaged per nozzle element of the microplasmajet array thruster operated at 6.2 W with 1.46 mg/sec of nitrogen gas propellant mass flow per nozzle element were 1.13 mN, 79 sec, and 24%, respectively. In addition, observation of the micro-nozzle discharges was conducted. To improve thrust performances and to achieve uniform discharges over all electrode elements, a new thruster with extended nozzle structure for which each electrode element was electrically isolated from nearby elements was developed. The typical values of the thrust and specific impulse of the multi-nozzle array thruster with 36 nozzle elements operated with 11 mg/sec of nitrogen gas propellant mass flow were 6 mN and 60 sec, respectively.

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Tohoku University is now finishing the development of its second 50-kg class Earth observation microsatellite RISING-2. RISING-2 is equipped with four camera systems with eight different imaging sensors, including a 5-m ground resolution cassegrain mirror telescope with three fixed wavelength filters in red, green, and blue, as well as a liquid crystal tunable filter. RISING-2 aims to achieve agile high resolution multi-spectral Earth observations. For this mission, Tohoku University has been developing a precise and agile attitude control system including multiple sensors and actuators. The pointing error is required to be less than 0.1 deg and the pointing knowledge shall be better than 0.06 deg. The designed star sensor is based on CCD image sensor which has a flight heritage by the university's first satellite RISING. This system utilizes a pyramid algorithm for secure constellation identification and works with a maximum update frequency of 2 Hz. Its availability is also very high in terms of the satellite's operating angular velocity. In this paper, the system integration of the star sensor as well as the results of its ground based performance evaluation is described in detail.

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The trajectory design associated with the deflection of potentially hazardous asteroid (PHA) is considered. At the terminal mission phase, low-thrust guidance control is used to compensate state errors and to change the intercept angle from nominal one to meet mission requirements. In this paper, the robust trajectory design using low-thrust acceleration is considered. We employ the generating functions approach proposed by Guibout and Scheeres (2004) and investigate its application to robust terminal guidance control. Our approach is illustrated through the study of the asteroid deflection mission.

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The 50-kg class micro satellite RISING-2 is now under development by Tohoku University and Hokkaido University. The development is at Flight Model phase and some components are tested to evaluate its specifications. In this paper, the endurance tests using vacuum chamber and thermal test chamber in Tohoku University are described. There was no problem in those tests. In addition, the attitude control accuracy of the RISING-2 was analyzed in the simulation. This simulation based on component specifications and also included noise data. The results show the pointing error angle was less than required error angle.

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We developed a novel coaxial rotary valve for a multi-tube PDE. Since this single valve can supply three different gases (fuel, oxidizer and purge gas) into a combustor, the unification of the valve systems for three different gases is possible by using our newly designed valve. A PDRE system can be simple and lightweight by using this valve, and thus its thrust-weight ratio can be increased. We proposed the design of a multi-tube rotary-valved PDRE system by this rotary valve. Moreover, in preparation for a multi-tube rotary-valved PDRE, we carried out the multi-cycle operation experiment by the single-tube rotary-valved PDRE system. The combustion wave velocity was measured to confirm the operation of the PDRE system. Deflagration-to-detonation transition (DDT) was confirmed and DDT distance decreased under the condition of high operation frequency. In addition, a maximum operation frequency was 159 Hz.

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The presence of an atmosphere on Mars signifies that an aircraft could travel in its atmosphere using the aerodynamic forces of flight. A reconnaissance aircraft offers the possibility to obtain high resolution data on a regional scale of several hundreds to thousands of kilometers, which cannot be achieved with rovers or satellites. However, conventional aircraft design cannot be applied for a Mars Airplane because of the nature of the Martian atmosphere and the constraints from transportation from Earth to Mars. This paper presents the conceptual design of a fixed-wing airplane for Mars, and the design rationale undertaken following the constraints set by the Martian environment. As a result, the main wing area was calculated to be 1.2 m². It was folded using two hinges to be packed into an aeroshell. The proposed Mars Airplane has four notches on its main wing. They make packing easier to keep the wing area maximized, making the Mars Airplane more compact.

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Hayabusa sample return capsule (SRC), separated from the mother spacecraft, entered the earth atmosphere over the desert of the Australia on June 13, 2010, and landed safe on the ground after passing through the excessively high aerodynamic heat load with about 14 MW/m². All the component modules of the SRC have been recovered by June 15. The present paper overviews the reentry operation and flight of the SRC and post-flight analysis of the recovered heatshields: The post-flight analysis program is planned by 3 steps and is under progress : The first step is reconstruction of the reentry flight trajectory and the flight environment. The reentry trajectory is to be reproduced synthetically taking account of the accuracy of the reentry orbit determination, the SRC landing point, atmospheric density and the wind on the reentry day. And the second step is to overview the SRC through non-intrusive method such as X-ray CT-scanning. The 3-dimensional laser-scanning has been also carried out for measuring the surface recession. Regardless of difficulty in the surface roughness, emissivity dispersion etc., 3-D surface has been numerically reconstructed and compared with the CAD data at the pre-flight.

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The increased drag and convective heating associated with flow at hypersonic speeds have a significant impact on the design of vehicles. To alienate this problem, a thorough literature review proclaims the research on using aerospikes and aerodisks as passive control devices. However, these studies have been conducted only on simple axisymmetric geometries and not on representative lifting body shapes. As a preliminary design study, the efforts were focused on analysis of using forward-facing aerospikes and its effectiveness with varying angle of attack (-10° to +10°) and spike-nose configuration, on key aerothermodynamic parameters of drag and heat reduction for prospective application to lifting body configuration, forming a delta shape with non-axisymmetric forward stagnation surface. Flow visualization was carried out using the schlieren technique. Measurements made at freestream Mach number 7 with six-component force balance system revealed a 100% increase in the Lift/Drag Ratio and no significant change in pitching moment coefficient compared to No-spike case, with aerospikes proving worthy at high angles of attack; thereby indicating their practical feasibility for eventual future applications to spacecrafts.

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We organize Ig Satellite Design Contest to collect the NEEDS and to establishment of NEEDS database for microsatellite. It will be polestar for the satellite developers. This paper shows an example of NEEDS collection and derivation of the key technologies from the NEEDS, which should offer a strategic direction to the SEEDS, the satellite developers for example.

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This paper describes design and current developing status of a flexible standard bus system for a new series of small scientific satellites planned in the Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (ISAS/JAXA). The satellite series are called SPRINT series, which stands for “Small space science Platform for Rapid INvestigation and Test”. First, features realizing a flexible standard bus are mentioned. Design concepts such as alternative options and selectable options are introduced. Second, the current status of the standard bus and SPRINT-A, which is the first satellite of SPRINT series using the standard bus and planed to be launched in 2013, is explained.

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In order to shorten time span for spacecrafts' onboard software, we have developed a Model Driven Architecture (MDA) tool named SIB2 generator: integrated development environment for onboard software. To SIB2 generator implemented is not only automatic source code generation but miscellaneous functions which accelerate implementation of application software. We also paid deep attention to enhance adaptability of the automatically generated code into various onboard components. This paper reports development-completion and design rational and results for initial version SIB2 generator.

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As a first step in the development of a space elevator, we focused on the stratospheric platform that would be used as a relay station of communication and high-resolution earth observation base at 20~30 km altitude. We propose a new system for connecting the platform to the ground by a tether that would allow the payload to be transported by a climber moving along the tether. We call this system a stratospheric elevator. The stratospheric elevator offers not only the chance to verify the space elevator technology but also the chance to extend the life of the platform by making its maintenance easy. The result of our initial study of the system showed that we can build a stratospheric elevator using currently available materials and technology. The total tether system, including a 200 kg payload and a 200 kg climber, would weigh about 2 tons with considering the effect of the high-speed airstream when we built the system at 20 km altitude, and a platform of about 80 m diameter could sustain the total system.

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The first and second H-II Transfer Vehicle (HTV) successfully completed their missions in Nov. 2009 and March 2011. Now, JAXA has the capability to upload lots of logistics to support human activities on the International Space Station (ISS). Between these HTV missions, JAXA established "ISS Cargo Return Vehicle Research and Development Office" as a new section to study a new spaceship concept. It is called "HTV-R" and an features improved type of HTV, which enables JAXA to recover various samples of experiments conducted on the ISS. HTV-R equips a new re-entry vehicle called as "HRV(HTV Return Vehicle)" for the enhanced function. Although most of the HTV-R mission does not differ from original HTVs, once the HRV separates from the HTV-R after completion of the de-orbit maneuver, the HRV will fly autonomously and conduct lifting reentry to the atmosphere with controlling the attitude and trajectory to the predefined splashdown point. Because HRV flight profile resembles that for the human returning flight, JAXA considers that the study for HTV-R is not that of a cargo recovery system but also as the pathfinder of human transport system.

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The Epsilon rocket, formerly called Advanced Solid Rocket (ASR) launcher, proceeded to the full development phase in August 2010 and its launch site was officially declared to be the Uchinoura Space Center (USC), the home of Japanese solid propellant rocket. The primary purpose of Epsilon is to provide small satellites with a responsive launch that means a low cost, user-friendly and ultimately efficient launch system. The slogan is “Small, Cheap, Fast and Reliable”. This outcome is also a result of the excellent endeavors of those who devoted themselves to the next generation solid propellant rocket. However, this is not the final destination. Now that the development was approved, the most important is what the next step should be beyond Epsilon. This paper deals with the significance of the development of Epsilon launch vehicle and how it contributes to the possible evolution of future space transportation systems.

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The purpose of “Alloy Semiconductor” crystal growth project is to make clear the factors for crystal growth of a high-quality bulk alloy semiconductor by investigating (1) solute transport in liquid and (2) surface orientation dependence of growth kinetics under microgravity and terrestrial conditions. The temperature gradient furnace Gradient Heating Furnace (GHF) onboard “Kibo” is used for the growth of an In_xGa_1-xSb bulk crystal which is a potential substrate material of optoelectronic devices such as thermo-photo-voltaic cells and gas sensors, since the band gap and the lattice constant of the crystals are tuned by adjusting the composition. The current status of the space experiment project will be reported in the presentation.

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The space astrometry satellite, Nano-JASMINE (Japan Astrometry Satellite Mission for Infrared Exploration), is a 5 cm diameter telescope designed to be sensitive to wavelength of zw-band (600< λ <1000 nm) and to achieve astrometry measurements of nearby bright stars (zw <7.5 mag) with an accuracy of 3 milli-arcseconds. This satellite will be launched by Cyclone-4 rocket from Brazil in 2013. A mission system flight model has been fabricated, and the performance has been evaluated; wave front errors were measured for the telescope mirrors and a radiation damage test was performed for a fully depleted charge coupled device image sensor used as a detector. This paper reports results of such a performance evaluation.

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Ka-band communications is one of the most important technologies for increasing the amount of data acquired in deep space missions. As a first step toward developing this technology, a Ka-band extender was attached to an existing X-band transponder. The Ka-band extender can generate Ka-band signals from the transponder’s signals. The extender is designed so as to reuse design elements of the X-band transponder. This approach ensures reliability of the extender without additional qualification, lowers production costs, and allows for flexibility in the Ka-band extender configuration. For instance, the minimum configuration is a simple upconverter, which is realized by sharing circuits to the greatest extent possible with the transponder. The extender is compatible with the Ka-band specifications in Consultative Committee for Space Data Systems standards. Properties such as a flexible coherent ratio, high-speed analog and digital modulation, and ultralow phase noise for radio science missions are provided. Here, a breadboard model of the Ka-band extender was evaluated in experiments. The Allan variance of the Ka-band output signal was less than 1 × 10^-12 (at 1 s), 1 × 10^-13 (at 10 s), and 1 × 10^-14 (at >100 s) when an external reference signal was used. The Allan variance degradation and phase noise degradation, which were caused by the internal phase locked loop or frequency translation loop, were also measured. The measured phase noise degradation was about 25 dB from the theoretical value.

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The Kaguya mission completed by collision of the Kaguya spacecraft on targeted place of the Moon on June 11, 2009. The Kaguya science team makes endeavors to archive data for data opening to public, and to study using the Kaguya data. Although data analysis and science study are ongoing, scientific achievements obtained so far are summarized concerning with the origin and evolution of the Moon as a ultimate science target of Kaguya mission: Ubiquitous identification of pure anorthosite in outcrops of central peaks of large craters by Multi-band Imager (MI) and Spectral Profiler (SP), discovery of multi reflectors of radio waves under large mares and ocean in the nearside by Lunar Radar Sounder (LRS), confirmation of free-air gravity anomaly in the whole Moon and identification of farside anomaly different from nearside one of mass concentration by Relay Satellite Transponder (RSAT), confirmation of lunar global topography by Lunar Altimeter (LALT), re-estimation of crustal thickness by Kaguya data of gravity and topography, re-estimation of formation ages of farside mares by crater counting using high resolution images of Terrain Camera (TC), confirmation of magnetic anomalies and mini-magnetosphere by Lunar Magnetometer (LMAG) and Plasma Angle Composition Experiment (PACE), reconfirmation of global distribution of radio-active elements K, U and Th by Gamma-ray Spectrometer (GRS), and discovery of solar wind proton reflection and access into lunar wake by PACE.

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This paper proposes a new method for identification of locations on an irregular-shaped small body. A 3D shape model of polygons is a representation of a small body, and an ID number of polygon can be a unique identifier of a particular location on the body. We also propose a graphical interface to specify locations of interests by identifying particular polygons in a 3D graphics display. With these concepts, we develop database system with a web-based interface for archiving images obtained through asteroid explorations missions. Database for this system stores images and a list of polygon IDs that are included in the field of view (FOV) of each image. We prepare the polygon ID-image FOV databases for Hayabusa and NEAR Shoemaker asteroid missions, by simulating actual FOVs of each images from ancillary data of spacecraft and shape models of target asteroids. Accuracy of FOV estimation and consistency of polygon ID databases are also assessed.

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In this paper we propose a new data format for lunar and planetary exploration called "XPEF" (XML-based Planetary Exploration Data Format). The XPEF is based on XML, a framework widely used for data description and exchange, and data themselves are stored in individual format with their thumbnails and packed as a single archive. Therefore, XPEF is database-friendly, easy to display, and robust from data corruption. Our purpose in designing a new data format is to improve the difficulty which researchers meet in using the PDS (Planetary Data System). Currently, most of lunar and planetary exploration data are dealt with the PDS. The data format used in PDS, however, has several inconvenient aspects from the view point of current informatics such as poor interoperability of database systems and insufficient capability for data integrity assurance and data browsing. In this paper, the concept and design of XPEF and implementation status of XPEF processing system, including data translation mechanism from the data format in the PDS to XPEF over the network called Messaging Network, will be presented.

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Since 2008, a new plan for next Mars exploration program has been proposed and discussed by scientists and engineers in Japan. This exploration program is named MELOS, or Mars Exploration with Lander and Orbiter Synergy, a long-awaited program in the planetary science community in Japan after unsuccessful end of Nozomi Mars orbiter and ongoing challenge of Akatsuki Venus orbiter. The goal of the whole program is to understand Mars as a system, by elucidation of Martian climate, atmospheric escape, internal structure, surface environment and interaction by them. A series of missions has been planned, and several spacecraft including orbiters and landers are under discussion to be launched in early 2020's. In this paper, we investigate launch opportunities during early 2020's and estimate the payload mass in each case. Feasible interplanetary transfer trajectories from Earth to Mars are proposed. Preliminary design of insertion sequence into the Mars orbit and some orbit candidates derived from mission requirements are also shown together with numerical simulation results.

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Using 4-way Doppler tracking with relay satellite OKINA, KAGUYA obtained the first precise gravity field of the lunar farside. Multi-frequency differential Very Long Baseline Interferometry (VLBI) observation using OKINA and OUNA improved the accuracy of gravity, through precise determination of OKINA's orbit. The current gravity model is SGM100i involving VLBI data. Laser altimeter (LALT) on board KAGUYA obtained the first precise global topography of the Moon with range accuracy of 5m. The correlation of spherical harmonics coefficients between gravity and topography become higher than that of the previous model. Gravity signatures of far-side impact basins are mostly explained by topography except for the central high. Combined with topography data, we estimate Bouguer anomaly and the crustal thickness variation of the Moon.

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Measurements of lunar gravity and rotation are important because they provide information of the physical state of the lunar interior. Previously only passive LLR (Lunar Laser Ranging) using CCR (corner cube reflectors) has been applied for the detailed study of lunar librations, rotation variability. As for candidate instruments for SELENE-2 (forthcoming lunar landing mission by JAXA) and future lunar missions, we propose VLBI (inverse-VLBI and differential VLBI) for gravity measurement to constrain tidal Love number, LLR and ILOM (In-situ Lunar Orientation Measurement) for libration measurements.

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EJSM (Europa Jupiter System Mission) is an international Jovian system mission with three spacecraft. Coordinated observation of Jovian magnetosphere is one of the important targets of the mission in addition to icy satellites, atmosphere, and the interior of Jupiter. JAXA will take a role on the magnetosphere spinner JMO (Jupiter Magnetospheric Orbiter), whereas ESA will launch JGO (Jupiter Ganymede Orbiter) and NASA will be responsible for JEO (Jupiter Europa Orbiter). One possibility is to combine JMO with a proposed solar sail mission of JAXA for Jupiter and one of Trojan asteroids. Since Trojan asteroids could be representing raw solid materials of Jupiter or at least outer solar system bodies,involvement of Trojan observation should enlarge the scope and enhance the quality of EJSM.

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The South Pole-Aitken basin (SPA) is the largest (2500km in diameter), deepest and presumably oldest impact basin in the solar system. SPA was characterized by an ellipse with axes 2400 by 2050 km, but more precise topography and gravity data are necessary to discuss the structure of large basin like SPA. We use localized representation of gravity potential where Slepian functions were used to estimate the gravity field over certain areas of the Moon and a spherical cap area with a radius of 40 degree from the SPA center. The direction of an ellipse denoting the depression is found to be similar to the previous result. The region with the thinnest crust is offset southward from the center of SPA. We analyzed interior structure of small basins in and around SPA. We interpret that a positive gravity anomaly corresponds to a Moho uplift. Just around the rim of SPA, a distinct Moho uplift beneath Schrödinger corresponds to the presence of olivine and obscure circular structure Amundsen-Ganswindt has a significant Moho uplift, suggesting a buried impact structure.

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The Greenhouse gases Observing SATellite (GOSAT) monitors atmospheric carbon dioxide (CO₂) and methane (CH₄) globally from space since February 7, 2009. The Thermal And Near infrared Sensor for carbon Observation Fourier-Transform Spectrometer (TANSO-FTS) and Cloud and Aerosol Imager (TANSO-CAI) acquire global data every three days. Onboard and vicarious calibrations have been performed, and differences from the pre-launch tests have been analyzed. The resulting radiometric, geometric, spectroscopic, and polarimetric characterizations are presented. In addition, we describe the recent status of on-orbit instrument parameters such as pointing accuracy, interferogram quality, and micro-vibration. Finally, the TANSO-FTS Level 1 data processing and correction algorithm and its updates on the ground are presented. The two-year on-orbit operations on orbit are summarized in chronological order.

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Gravity resistance is a principal graviresponse in plants. In resistance to hypergravity, the gravity signal may be perceived by the mechanoreceptors located on the plasma membrane, and then transformed and transduced via the structural continuum or physiological continuity of cortical microtubules-plasma membrane-cell wall, leading to an increase in the cell wall rigidity as the final response. The Resist Tubule experiment, which will be conducted in the Kibo Module on the International Space Station, aims to confirm that this hypothesis is applicable to resistance to 1 G gravity. There are two major objectives in the Resist Tubule experiment. One is to quantify the contributions of cortical microtubules to gravity resistance using Arabidopsis tubulin mutants with different degrees of defects. Another objective is to analyze the modifications to dynamics of cortical microtubules and membrane rafts under microgravity conditions on-site by observing green fluorescent protein (GFP)-expressing Arabidopsis lines with the fluorescence microscope in the Kibo. We have selected suitable mutants, developed necessary hardware, and fixed operation procedure for the experiment.

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A wide variety of organic compounds have been detected in such extraterrestrial bodies as carbonaceous chondrites and comets. Amino acids have been confirmed in extracts from carbonaceous chondrites and cometary dusts. It was suggested that these organics were formed in quite cold environments. We irradiated possible interstellar media, such as a frozen mixture of methanol, ammonia and water, with high-energy particles. Amino acid precursors with high molecular weights were detected in the irradiated products. Such complex amino acid precursors are much more stable than free amino acids against radiation, and heat. It is suggested that interplanetary dust particles (IDPs) brought much more organics than meteorites and comets. However, characteristics of organic compounds in IDPs are little known, since they have been collected only in terrestrial biosphere. We are planning the Tanpopo Mission, where IDPs would be collected in aerogel equipped on the Exposure Facility of the International Space Station. In addition, amino acids and their relating compounds would be exposed to space environments to see their possible alteration processes.

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