In the future, carbon fiber reinforced plastics (CFRPs) with high hydrogen gas barrier performance will find wide applications in all industrial hydrogen tanks that aim at weight reduction; the use of such materials will be preferred to the use of conventional metallic materials such as stainless steel or aluminum. The hydrogen gas barrier performance of CFRP will become an important issue with the introduction of hydrogen-fuel aircraft. It will also play an important role in realizing fully reusable space transportation system that will have high specific tensile CFRP structures. Such materials are also required for the manufacture of high-pressure hydrogen gas vessels for use in the fuel cell systems of automobiles. This paper introduces a new composite concept that can be used to realize CFRPs with high hydrogen gas barrier performance for applications in the cryogenic tanks of fully reusable space transportation system by the incorporation of a nonmetallic crystal layer, which is actually a dense and highly oriented clay crystal laminate. The preliminary test results show that the hydrogen gas barrier characteristics of this material after cryogenic heat shocks and cyclic loads are still better than those of other polymer materials by approximately two orders of magnitude.

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This paper describes the experimental study of impact testing and permeability testing of CFRP laminates for application of propellant tanks considering the reliability of the tank structures associated with foreign object impact damage. Three types of quasi-isotropic laminates with different laminate thickness were prepared and subjected to transverse impact tests, nondestructive inspections, and helium gas permeation tests in series. Transverse impacts induced delaminations, matrix cracking, and fiber breakage, which caused severe through-thickness gas leakage. Effects of impact energy and laminate thickness on the permeation-after-impact properties are discussed, and the critical damage mode controlling the severe gas leakage is identified.

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A new ceramic thruster for an interplanetary probe is currently under development. Monolithic silicon nitride (Si₃N₄) , which has good heat resistance and high fracture toughness among conventional structural ceramics, is a promising material for a high performance thruster. However ceramics are brittle compared to metallic materials. In order to evaluate reliability of the ceramic thruster as a space-use component, fracture behavior against micrometeoroid impacts was investigated. First the risk probability of the meteoroid impacts which may occur during a mission was estimated based on impact energy which may cause failure of the material. Second, damage of the silicon nitride ceramics by a possible micrometeoroid impact was investigated experimentally. Hypervelocity impact tests were carried out on the silicon nitride ceramic samples with a two-stage light-gas gun. Impacts at various velocities ranging from 1.0 km/s up to 4.5 km/s brought about three types of failure. However no shattering occurred by the hypervelocity impact with a possible energy. The experimental results together with the risk evaluation considering the flight mission conditions show that the Si₃N₄ ceramic thruster for the interplanetary probe would have no serious problems caused by a meteoroid impact during the flight mission even with local damage.

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This study evaluated effects of a low earth orbit (LEO) space environment on properties of a solid lubricant used for space applications. The tested lubricant was a bonded MoS₂ film with organic binder. The film was exposed to a real LEO space environment aboard the International Space Station (ISS) by the Space Environment Exposure Device (SEED) experiment for about 1400 days. Additionally, it was irradiated individually with atomic oxygen (AO), ultraviolet rays (UV), and electron beam (EB) on the ground. Fluences of these factors corresponded to exposure to the LEO environment by the SEED experiment. Friction tests in vacuum and surface analyses were carried out for the samples. Tribological behavior of the different samples was measured using classical reciprocating pin-on-flat friction tests. Furthermore, XPS analysis was performed for the film surface and rubbing tracks of the samples. Results show that the friction coefficient decreased by AO irradiation at an early stage of the tests. It resembled the result of the film exposed to the real LEO environment.

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The effect of hyperthermal atomic oxygen (AO) exposure on a surface property of Si-doped DLC was investigated. Two types of DLC were tested which contain Si atoms approximately 10 at% and 20 at%. Surface analytical results of high-resolution x-ray photoelectron spectroscopy using synchrotron radiation (synchrotron radiation photoemission spectroscopy; SR-PES) as well as Rutherford backscattering spectroscopy (RBS) have been used for characterization of the AO-exposed Si-doped DLC. It was identified by SR-PES that the SiO₂ layer was formed by the hyperthermal AO exposure at the Si-doped DLC surface. RBS data indicates that AO exposure leads to severe thickness loss on the non-dope DLC, in contrast, SiO₂ layer formed by the hyperthermal atomic oxygen reaction at the Si-doped DLC protects the DLC underneath the SiO₂ layer.

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Inflatable structure is an attractive candidate as ultra-light space structures with high packaging efficiency. They consist of flexible membrane and some components for the deployment. For such pneumatic structure with membrane, the damage of the membrane caused by space debris is one of serious problems for the utilization in space. To solve the problem, multi-cellular inflatable structures are proposed. Inflatable structures composed of multi-cells are robust against the damage of membrane. In this paper, a new type of inflatable cylinder which is composed of sealed multi-cells was investigated. Because deployment of the structure can be driven by the gas pressure of the sealed multi-cells, no complex components such as gas tanks and electric valves are required for the deployment. The fundamental characteristics of the deployment are investigated experimentally. As a result, it is shown that the inflatable cylinder composed of sealed multi-cells can achieve the deployment with simple hold-release mechanism even though some cells are damaged. Furthermore, by applying with inner bellows and outside constraint wires, the deployment sequence can be controlled and almost straight deployment can be realized.

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As an effective countermeasure for suppressing space debris growth, the Aerospace Research and Development Directorate, Japan Aerospace Exploration Agency (JAXA) has been investigating an active space debris removal system that employs highly-efficient electrodynamic tether (EDT) technology as its orbital transfer system. A test flight experiment using a small satellite is planned for establishing and demonstrating the EDT technology. Precise tether dynamics during deployment phase is investigated in this study by numerical simulations in which the coiling of the tether is taken into consideration. The results of the simulation showed that the tether length becomes shorter than the full tether length when the coiling of the tether is severe. The in-plane libration angle of the tether grows larger when the spring constant of the coiling of the tether is specific value. The attitude motion of a mother satellite is less stable when the spring constant of the coiling of the tether is small because of the large tether tension.

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The particle swarm optimization algorithm is applied to the non-linear optimization of the variable geometric parameters of the adaptive truss structure. The design variables to be optimized are the length of the truss member of the variable geometry truss, and the purpose of the optimization is to enhance the identification accuracy of the structural stiffness parameters under the concept of self-identification. The objective function in the optimization consists of multiple cost functions; a function of the condition number of coefficient matrix of linear equation for the identification and the criterion for evaluating the linear independence of the mode shapes. When the control resolution of the actuator is considered in the optimization, the discrete value of the geometric parameters is optimized by the particle swarm optimization algorithm with a penalty function. The global optimum for the continuous geometric parameters is also searched to compare the optimization results with the discrete ones. Numerical experiments on the optimization of the two-dimensional variable geometry truss are presented to show the effect of the penalty function on the value of the objective function for each mode. The deviation of the optimized value of objective function is also discussed.

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In this study, the cooling performance of a C/C composite material structure with metallic cooling tubes fixed by elastic force without chemical bonding was evaluated experimentally using combustion gas in a rocket combustor. The C/C composite chamber was covered by a stainless steel outer shell to maintain its airtightness. Gaseous hydrogen as a fuel and gaseous oxygen as an oxidizer were used for the heating test. The surface of these C/C composites was maintained below 1500 K when the combustion gas temperature was about 2800 K and the heat flux to the combustion chamber wall was about 9 MW/m². No thermal damage was observed on the stainless steel tubes that were in contact with the C/C composite materials. The results of the heating test showed that such a metallic tube-cooled C/C composite structure is able to control the surface temperature as a cooling structure (also as a heat exchanger) as well as indicated the possibility of reducing the amount of coolant even if the thermal load to the engine is high. Thus, application of this metallic tube-cooled C/C composite structure to reusable engines such as a rocket-ramjet combined-cycle engine is expected.

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Fiber-matrix interface mechanical properties are the key design point for the composites, however, methodology for evaluating interface properties of C/Cs are not still well established. In this study, capability of the fiber bundle push-out method was demonstrated as a measurement technique of fiber-matrix interfacial mechanical properties of various C/Cs. First, to carefully determine the experimental conditions to give proper interface mechanical properties, the influences of a radius of an indenter and the thickness of specimens on the measured results were examined. From the experimental result, interface debonding stress, τ_d, and sliding stress, τ_s, could be obtained as a constant value independent from specimen thickness as long as tested using the same indenter size and the proper specimen thicknesses of < 400 µm. However, absolute value of especially τ_d showed dependence of on indenter size. This was shown to be the limitation of this method that τ_d obtained from this method cannot be treated as a quantitative value. To demonstrate the capability of the method, the change of fiber-matrix interface mechanical properties of C/Cs on the type of fibers and final heat treatment temperature. τ_d decreased with the heat treatment temperature in both C/Cs with pitch based carbon fibers, K633 and K321 and C/Cs reinforced with a carbon fiber with higher heat treatment temperature gave lower τ_d at the same heat treatment temperature. Main reason for the degradation of τ_d up to 2273K was attributed to the enlargement of debonding area at fiber-matrix interface. Degradation of τ_d was enhanced at the temperature range of > 2273 K by the structure change of the interface from the graphitization of matrices. Finally, The fiber bundle push-out method was proved to be a strong tool to investigate fiber-matrix interface mechanical properties of C/Cs.

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The objective of this paper is to describe the current state of the art of the research on Ultra High Temperature Ceramic materials with particular reference to their space applications, and also to report on the activities performed on UHTC in the past decade by the Italian Aerospace Research Centre in the specific technological field of structural thermal protection systems. Within several internal research project, various UHTC composition, mainly based upon Zirconium Diboride and Hafnium Diboride with added secondary phases and sintering aid were examined characterized in their mechanical properties and oxidation resistance. Two main composition were selected as the most promising for hot structure manufacturing: these materials were extensively characterized in order to obtain a comprehensive database of properties to feed the thermomechanical design of prototype hot structures. Technological demonstrators were manufactured by hot pressing followed by further fine machining with Electrical Discharge methods, and then tested at high temperature for long times in a plasma torch facility. The main outstanding results obtained are discussed in this paper. Future outlooks related to the UHTC technology and its further development are also provided.

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This paper presents the results of on-orbit system identification of the Engineering Test Satellite-VIII (ETS-VIII), which was launched by the H-IIA Launch Vehicle No. 11 in December 2006. On-orbit system identification experiments were performed using data acquired during normal operations in the initial check-out phase, such as station acquisition and wheel unloading. The satellite's unconstrained modes were identified using attitude rates and data from accelerometers located on the main body, solar array paddles and large deployable antenna reflectors at various solar array paddle angles. In addition, the constrained modal parameters of the solar array paddles and large deployable antenna reflectors were estimated for comparison with the analytical model.

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The influence of heat treatment conditions and severe plastic deformations (SPD) by ECAP on structural and mechanical properties development of EN AW 6082 aluminium alloy was investigated. The hardness and mechanical properties in dependence on two types of solution heat treatment (T_SA=400°C+slow cooling and T_SA = 550°C +quenching), natural or artificial ageing treatment (room temperature, 100°C, 170°C), several annealing temperatures (T_A=100 - 300°C) and ageing or annealing times after SPD were tested. The best properties for EN AW 6082 material were achieved after processing by (T_SA=550°C/1,5h + water quenching + 3xECAP/20°C + T_AA=100°C/30h) on level: 0,2%YS=408MPa, UTS=427MPa, El.=20,1%, Re.=29,3%, KCV_RT=16,5J.cm^-2.

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This paper presents a program for the multidisciplinary optimization and identification problem of the nonlinear model of large aerospace vehicle structures. The program constructs the global matrix of the dynamic system in the time direction by the p-version finite element method (pFEM), and the basic matrix for each pFEM node in the time direction is described by a sparse matrix similarly to the static finite element problem. The algorithm used by the program does not require the Hessian matrix of the objective function and so has low memory requirements. It also has a relatively low computational cost, and is suited to parallel computation. The program was integrated as a solver module of the multidisciplinary analysis system CUMuLOUS (Computational Utility for Multidisciplinary Large scale Optimization of Undense System) which is under development by the Aerospace Research and Development Directorate (ARD) of the Japan Aerospace Exploration Agency (JAXA).

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Radiation damage on dielectric thin-layer was studied. The dielectric multilayer coating is a promising technology for the use of solar light collection optics in Space Solar Power System (SSPS) in the geostationary orbit. We estimated spatial distributions of numbers of displacement atoms using a Kinchin-Pease model for the proton incidence, and Frenkel pairs created via self-trapped exciton process initiated by electron excitation in the target for the proton and electron incidences. Proton incidence gives damages near the surface of the target due to low energy protons dominating its energy spectrum. Proton irradiation experiment for a thin layer of several dielectric targets shows toughness for about 10-year-equivalent exposure time at geostationary orbit.

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This paper presents a thermal p-version multidimensional finite element for ablative materials which can be applied to the simple thermal structural analysis of an atmospheric re-entry vehicle's ablative thermal protection system. The element is integrated into the element library of a multidisciplinary optimization analysis system developed by the Japan Aerospace Exploration Agency's Aerospace Research and Development Directorate. As with a general thermal shell element, a user of this system can set the initial thickness of the virgin layer as a parameter instead of by coordinates, allowing parametric studies to be easily conducted. The element can be laminated with other general thermal element models; for example, an aluminum primary structure-strain isolation pad-ablative material shell configuration typical of a re-entry vehicle thermal protection system.

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White paints are used as thermal-control materials for spacecraft, often on complexly shaped structures on which it is difficult to apply multi-layer insulators. They present advantages over most polymeric films (single or multi-layer insulation) because they are more stable against atomic oxygen. Atomic oxygen in a low earth orbit environment strikes spacecraft and chemically reacts with materials, producing erosion or oxidation. JAXA has conducted space exposure experiments using the Space Environment Exposure Device on the Russian Service Module in the International Space Station to evaluate the tolerance of materials in a low earth orbit. Silicone-based white paints were exposed on the device for three different exposure periods, demonstrating high resistance. Silicon contaminants were observed on the retrieved paint surfaces.

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The objective of this study is to align a certain axis of a spacecraft in an arbitrary direction by using a variable-speed single-gimbal CMG. Because the total angular momentum of the spacecraft is conserved in inertial space, the final attitude of the spacecraft is restricted, and it is possible to achieve pointing control of the axis that is orthogonal to the gimbal axis. Using new parameters to express pointing errors, we have proposed a simple pointing control law that can be used to align the pointing axis in an arbitrary direction in inertial space. The validity of the proposed control law is verified by numerical simulations.

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This paper presents a new method for generating motion profiles of rotating bodies, for example, the joint angle motion of mechanical systems such as space manipulators and antennas. In the planning of a motion from the given initial conditions to the final conditions, there are many constraints that should be considered, for example, actuator torque limits, rate limits, jerk limits, etc. Furthermore, flexibility of a system cannot be ignored in general. Vibration suppression is of great importance in some applications. For stationary boundary conditions (rest-to-rest cases), if the only constraint is the torque limit, a simple bang-bang type profile is the optimal one. However, if we consider more general boundary conditions, complex constraints, and vibration suppression conditions, the problem becomes considerably complicated. In this study, motion planning for a fixed motion time is considered. The control input or the acceleration of the motion is expressed as a linear superposition of triangle waves whose weight coefficients are determined in order to optimize the motion profile. In this formulation, all the boundary conditions and several constraints, including vibration suppression conditions, are expressed as linear equality and inequality constraints. Therefore, by appropriately setting the performance index, the problem becomes a linear programming problem and can be solved efficiently.

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This paper presents a new attitude determination algorithm for a spacecraft in low earth orbits using magnetic sensors and rate sensors (e.g. quartz rate sensor gyros) for generic missions including earth observations. The algorithm is applicable to mainly a three-axis spacecraft, including when it is tumbling just after the separation from a launch vehicle or due to some anomaly. The advantage of the proposed approach is that it does not require additional attitude sensors such as sun sensors. First, an attitude determination algorithm using only magnetic sensors is introduced presuming that the attitude change during the measurement period is very small. Then the algorithm is extended to combine magnetic sensors with rate sensors, where both the three-axis attitude and rate sensor biases are estimated. The estimation error is statistically analyzed. Numerical examples are given to validate the proposed algorithm and its accuracy evaluation.

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Tohoku University and Rikkyo University are carrying out the project of Venus observation with high precision using a balloon-borne telescope. In this paper, the outline of Balloon-Borne Telescope for the optical remote sensing of Venus is introduced, and the simulation model of three-stage control method is constructed. For this observation, the pointing technology with high precision to restrain the slight moving of image is necessary. The target precision is only 0.1 arc seconds. The dynamics and control model is defined firstly, and the model parameters are determined by the experimental verification. By developing the numerical simulation tool, the motion can be estimated in the simulator, and the control strategy can be more easily optimized compared to the gain adjustment only based on experiments.

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Future scientific missions, such as virtual telescopes or interferometers, will require precise formation flying, such that the relative positions of spacecraft are controlled very precisely. The present paper discusses how to suppress relative position variation during one orbit. In the present paper, we propose a new approach to choose an adequate initial position and velocity, such that the relative position is maintained when averaged over one orbital motion, and design a predictive, discrete feedback controller. An along-track formation control in low earth orbit is taken as an example, where the J2 term and air drag are the two representative natural disturbances. Numerical simulations are performed in order to validate the proposed controller and investigate how the size of minimum impulse affects control accuracy.

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This paper deals with in-plane motion of a solar sail spacecraft to rendezvous with a planet. In the interplanetary rendezvous problem, the spacecraft's velocity must coincide with the orbital velocity of the planet when it reaches the planet's orbit. Thus, the spacecraft's radial and tangential velocities as well as its orbital radius are controlled by one control input, i.e. the spacecraft's pitch angle. In this paper, we propose a trajectory design method which can reduce the amount of computational iterations considerably. This is applied to a rendezvous mission to a planet in a circular orbit and is achieved by utilizing locally optimal control techniques. A hidden problem in the method is pointed out, and a countermeasure is proposed. Then, numerical results of the proposed method are shown and compared with the results obtained by a fully numerical iteration method. Finally, some mathematical properties of a sailcraft's governing equations are discussed in the framework of nonlinear control theory. We show analytically that a solar sail spacecraft can rendezvous with any planet in any elliptical orbit by using only pitch angle control.

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This paper explains the attitude control method for a 50 kg class satellite QSAT. It uses three-axis magnetorquers for control and a gravity-gradient extension boom for enhancement of the attitude stabilization. We divide the mission period into three main attitude-control phases. The first phase refers to the de-tumbling which ends when one particular satellite axis is roughly along the local Earth-pointing direction. In the second phase, we extend the boom. The third phase refers to the normal mission mode, when accurate attitude control will be achieved by means of the magnetorquers. We construct a number of normal-mode control algorithms and evaluate their performances by simulations. We calculate the required control torque by the PD (Proportional-Derivative) control method. The differences of the methods appear in the calculations of the required magnetic moments. In this paper, we simulate the following methods; 1) simplified linearized dynamics equation method, 2) the LQR (Linear Quadratic Regulator) method, and 3) the cross product method which utilizes only data from the magnetometer and the rate gyros. From the simulations, we conclude that “simplified linearized dynamics equation method” is the most efficient method for QSAT.

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Solar sail is one of the promising propulsion systems for future deep space exploration missions as it does not require any fuel to acquire propulsive force. Although folding method and deployment mechanism of the sail have been intensively developed, attitude control of the solar sail, which is necessary for the orbital control by the solar sail, has not been much studied. This paper discusses the attitude dynamics and the control method of a spinning type solar sail spacecraft. The spinning type solar sail, where the sail equipped around the spacecraft hub is to be deployed and extended by centrifugal force, has no rigid structure supporting its membrane. This type of mechanism has the advantage in its simple and lightweight structure, however, the attitude control is difficult due to the flexibility of the membrane. In this paper, we introduced a mathematical dynamics model including first oscillation mode of the membrane which can handle coupled motion of a rigid spacecraft and a flexible membrane, and analytically developed a controller that can avoid unnecessary oscillatory motion. The performance of the controller was verified by numerical simulations using more precise multi-particle numerical model.

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This paper discusses the in-planar and out-of-planar dynamics of a Tethered Satellite System (TSS), which aims to inject a satellite into a different orbit plane. The orbital transfer is achieved by only varying the tether length in a gravity gradient field. A previous work treating in-plane orbital transfer induces the system's pitch motion, and proposes a control procedure. For a transfer in a different orbit plane, the system's roll motion must be controlled as well as its pitch motion. First, this paper shows the system's governing equations of motion, and compares the features of the roll motion with its pitch motion. Then, a control method is proposed to increase the angular momentum of the roll motion. Finally, the effectiveness of the proposed method is demonstrated by numerical simulations.

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In Space Systems Dynamics Laboratory at Kyushu University, the mini-satellite QSAT is being developed. This satellite aims at investigating the plasma physics in the aurora zone in order to better understand spacecraft charging and at conducting a comparison of Field-Aligned-Currents observed in orbit with ground-based observations. In order to achieve the mission objectives, the spacecraft attitude must be determined. The attitude determination concept of QSAT is based on a combination of the Weighted-Least-Square and Linearized-Kalman filter estimation methods. The Weighted-Least-Square method produces the optimal attitude-angle observations at one point in time by using the Sun sensor and magnetometer measurements. The recursive Linearized-Kalman filter combines the angular observations with the attitude rate measured by the gyros to produce the optimal attitude solution.

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Reaction wheel assembly used as main actuator of an attitude control system can be a major source of disturbances. Although vibration isolator for a reaction wheel assembly is required to attenuate a disturbance, a control torque has to be transmitted for an attitude control of a satellite. This paper introduces a new design method of vibration isolator for reaction wheel assembly. Proposed method is based on H-infinity formulation and optimizes not only parameters of vibration isolator, but also parameters of satellite's controller simultaneously. At first, numerical model of a parallel link isolator is formulated. Next, the formulations of new design method and constraint conditions are described. After that, results of numerical analysis and availability of proposed method are shown.

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Engineering Test Satellite-VIII (ETS-VIII) is equipped with two large deployable reflectors (receiver (RX) and transmitter (TX) LDRs) that are deployed on orbit. The deployment sequence consists of two steps: one is the boom deployment step and the other is the reflector deployment step. The results of deployment at each step is monitored by a camera attached to the satellite. In order to deal with off-nominal scenarios due to camera malfunctions, the boom deployment angle is estimated from the telemetry data obtained from the attitude and orbit control system (AOCS) as well as the images captured by the camera. This paper deals with the estimation of the boom deployment angle from the AOCS telemetry data; further, the results of the on-orbit deployment are presented.

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In this paper, a control strategy for a spacecraft (“Pursuer”) to rendezvous with an uncooperative object (“Target”) is dealt with. This strategy consists of a feedforward optimal control phase (phase 1) and a feedback control phases (phase 2 and 3). In the feedforward optimal control phase, the target's future attitude motion is predicted from the result of motion estimation using image data in order to determine the final state for the maneuver. It is expected that there will be errors in the result of motion estimation for the target and error compensation is expected to be necessary. For this purpose, the pursuer estimates and predicts target's motion during maneuver, and recomputes the optimal control input using it. A quasi-optimal control input is proposed and named “Adjusted Control Input” in order to adapt this error compensation. And its performance is investigated through a numerical simulation of the pursuer's position control.

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The Japan Aerospace Exploration Agency (JAXA) is studying the feasibility of using the solar power sail as a new propulsion engine for deep space exploration missions. In this paper, the sail shape and equipment layout for missions utilizing small-sized solar power sails are proposed. The two-stage deployment method of the sail is also proposed. The sail need to be deployed statically at the first stage, and two types of deployment mechanisms are introduced. On the other hand the second stage of the deployment can be performed dynamically, and the oscillating motion of the membrane is converged by tethers connecting the membrane to the main body. The deployment motions are analyzed by numerical simulations using multi-particle models in order to verify the deployment. They are compared with the results calculated by finite element method models. The numerical simulation results are discussed from the technological viewpoint of the sail deployment dynamics and mechanisms.

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This paper presents a generic method for optimal motion planning for three-dimensional 3-DOF multi-link robotic manipulators. We consider the operation of the manipulator systems, which involve constrained payload transportation/ capture/release, which is a subject to the minimization of the user-defined objective function, enabling for example minimization of the time of the transfer and/or actuation efforts. It should be stressed out that the task is solved in the presence of arbitrary multiple additional constraints. The solutions of the associated nonlinear differential equations of motion are obtained numerically using the direct transcription method. The direct method seeks to transform the continuous optimal control problem into a discrete mathematical programming problem, which in turn is solved using a non-linear programming algorithm. By discretizing the state and control variables at a series of nodes, the integration of the dynamical equations of motion is not required. The Chebyshev pseudospectral method, due to its high accuracy and fast computation times, was chosen as the direct optimization method to be employed to solve the problem. To illustrate the capabilities of the methodology, maneuvering of RRR 3D robot manipulators were considered in detail. Their optimal operations were simulated for the manipulators, binded to move their effectors along the specified 2D plane and 3D spherical and cylindrical surfaces (imitating for example, welding, tooling or scanning robots).

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This paper presents a novel attitude control device which is called three dimensional reaction wheel (3DRW). 3DRW consists of only one levitated spherical mass which can rotate around arbitrary axes. This leads to the reduction of the weight and volume of the device as compared to existing reaction wheel. Furthermore, this device has no mechanical contact between rotor and stator, so the failure caused by the mechanical contact would be reduced. In this paper, the results of the analysis and experiment on the dynamics and control of 3DRW are shown. In the experiments of the rotation control, the air bearing system is used. Using this device, the characteristics of rotation of the spherical mass are obtained. To verify the feasibility of the concept of 3DRW, the experiments of angular velocity feedback control are carried out. The results of experiments are applied to the numerical simulation of the attitude control for microsatellites, and the feasibility of 3DRW is verified.

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Reported in this paper are the results of the orbit maneuver compensation in KAGUYA's Lunar transfer. Because of the uncoupled allocation of the attitude control thrusters, extra velocity increment (δv ) is induced whenever KAGUYA performs an orbit maneuver. Since the observed level of δv was unacceptable range from the point of maneuver accuracy requirement, it was compensated by means of deducting estimated δv from the orbit maneuver command. The δv estimation model was updated step-by-step during the Lunar transfer, which leaded to significant improvement of the orbit maneuver accuracy and resulted in the omission of the last trajectory correction maneuver. The method of the compensation and its results are introduced in detail.

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In September 2005, Hayabusa (MUSES-C) spacecraft successfully had a rendezvous with asteroid 25143 Itokawa. After the arrival, Hayabusa made detailed observations of the asteroid during its rendezvous period (about three months). As the results of various kinds of scientific analysis, a variety of physical parameters of Itokawa (e.g. size, volume, mass, and density) were derived. As to the orbit determination of Hayabusa spacecraft, during the cruise phase, the radiometric (2-way X-band range and Doppler) data were used for analysis. On the other hand, during the approach phase or rendezvous phase, we could obtain the optical data by means of star tracker or optical navigation camera, thus both the radiometric and the optical data were used for orbit determination. The present paper will report on the results of the orbit determination of Hayabusa during the close proximity phase. We will also mention about the mass estimation of Itokawa in this period. The data used in this analysis are 2-way X-band Doppler data and the position data, which were calculated from optical navigation camera's data. As well as the large orbital maneuvers and the gravitational acceleration of Itokawa, the effect of solar radiation pressure, and the effect of attitude control are also taken into account for the calculation. As to the gravity model of Itokawa, a spherical-harmonics gravity model or a polyhedron gravity model are used depending on the situation.

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To conduct remote inspection missions, an experimental device using a flexible beam without any articulated joints was developed and it is called Morphable Beam Device (MBD). A beam is deployed, enabling a wide range of shapes and lengths in a microgravity environment for various space applications, such as an inspection device, a crew support device and a variable geometry beam member of space structures. In this paper, a beam shape modeling made by the MBD considering retro-bending and beam shape generation algorithms are described, and feasibility of the proposed methods are demonstrated.

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The application of the variable geometry (VG) wing to a lifting re-entry body is expected to enhance the control capability of its aerodynamic characteristics and, as a result, to widen the corridor for the flight trajectory. In the present study, the flow field around a plain delta wing having three chord-wise hinges, one is on the wing root and the others on both sides of the mid-span of the wing, at Mach number 3 is numerically investigated by solving the Euler equations. The effects of the angle of attack and the “tip-down” bending angles around these hinges are clarified. The results show that the lift-to-drag ratio is hardly affected by the tip-down angle and that the overall lift and drag forces vary almost proportional to the change in the projected wing area by taking the tip-down configuration. The center of pressure moves backward by the tip-down effect.

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In the present work, a plasma wind tunnel using magnetic nozzle was developed and the flow characteristics were investigated. With stronger magnetic filed applied, the plasma flow was more compressed and squeezed in the radial direction, and elongated downstream, which was obviously confirmed by the captured images. The axial distributions of translational, rotational, vibrational, electronic excitation, and electron temperatures were measured by emission spectroscopy. Translational and rotational modes were in equilibrium, and their temperatures decreased downstream. The vibrational mode and electron translational mode were almost frozen. It was experimentally confirmed that the equilibrium or nonequilibrium phenomena occurred in magnetic nozzles in the similar way as seen in conventional solid nozzles.

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In order to predict the thermal and mechanical loads during re-entry, the URANUS (Upwind Relaxation Algorithm for Nonequilibrium Flows of the University of Stuttgart) has been being developed at the Institute of Space Systems (IRS) of the Universität Stuttgart. For the accurate determination of the thermochemical conditions, advanced thermochemical relaxation models for the gas-phase as well as sophisticated gas-surface interaction models have been developed. The Navier-Stokes equations for the 11-component air flow, which consists of N₂, O₂, NO, N, O, N₂⁺, O₂⁺, NO⁺, O⁺, N⁺ and e^-, have been derived by the Chapman-Enskog method from the Boltzmann-Equation. The linearized system of equations is solved fully coupled and fully implicitly, employing Newton's method. In the paper, the modelling is described in some detail and selected simulation results are presented.

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An aeroshell made from membrane material have an advantage of reduction in the aerodynamic heating, because its small mass and large area enable us to make the low-ballistic-coefficient flight, in which the vehicle decelerates at very high altitude with low atmospheric density. In this paper, we propose a new concept of mini re-entry system for small satellites. This vehicle is called "FEATHER" (Flexible Expanded Aeroshell with Tiny payload Harness for Entry and Recovery). "FEATHER" is a novel re-entry and recovery system, featuring the autonomous aeroshell deployment, the low-ballistic-coefficient re-entry with less severe aerodynamicc heating and so on. FEATHER is composed of the membrane aeroshell made from the high-temperature cloth called ZYLON®, an outer frame made of Shape Memory Alloy (SMA) and a payload. When the aeroshell receives the aerodynamic heating, the temperature of SMA frame rises and restores the circular shape as memorized beforehand. Then the membrane aeroshell is automatically deployed. Therefore the vehicle can achieve the low-ballistic-coefficient flight with a drastic reduction in the aerodynamic heating without any additional sensors, controllers and actuators. The preliminary studies made on FEATHER system so far including the hypersonic wind tunnel experiments are presented in this paper.

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A laser driven plasma wind tunnel using a 2 kW class continuous wave laser was developed as a high speed and high density atomic oxygen generator. Firstly, its operation conditions were examined. Using argon and oxygen as working gases, laser sustained plasma (LSP) was successfully produced in the plenum pressure range from 0.30 MPa to 0.95 MPa, and then the LSP was expanded to the vacuum chamber through the convergent-divergent nozzle. Next, plume characteristics were evaluated by Pitot probe and laser absorption spectroscopy using an absorption line of ArI 772.38 nm. As a result, the Mach number and the specific enthalpy around the center were 5.0 to 6.5 and 3.0 MJ/kg to 5.2 MJ/kg, respectively. The maximum flux density of atomic oxygen was estimated as 2.2×10²¹ cm^-2s^-1.

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To design the LOX/LNG ablative combustor, it is indispensable to build up the mathematical ablation model. In this paper, the mathematical model of SiFRP has been developed, which successfully predicts the penetration depth of both charred and decomposed zones and also the temperature profiles in the various kinds of experimental results which have been conducted to verify the model. Using this model, the peripheral-zone temperature profiles in the combustor, that is to say, combustor wall heat flux profiles, are estimated to reproduce the char penetration depth which has encountered in the ground firing tests, and the model has been successfully applied to design a LOX/LNG combustor.

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An existing computational fluid dynamics code which simulates a high enthalpy arc heater flow for air, named ARCFLO3, is updated to be able to calculate a high temperature carbonaceous flow. The flowfield is assumed to be in thermochemical equilibrium and the thermodynamic and transport properties for high temperature carbonaceous gas mixture are calculated though the Yos' mixture rule by compiling the most updated set of collision integrals. The radiative transport in arc heater is calculated in a fully coupled manner with flow motion by using a 3-band radiation model developed in the present study. The upgraded code is used to calculate the operational characteristic parameters such as arc voltage, chamber pressure, heater thermal efficiency and mass-averaged enthalpy for existing 60MW arc heater geometry by using CO₂ as a test gas. The result shows that by specifying the wind tunnel operating conditions routinely used for air, the similar level of the operational characteristic parameters is obtained, and that a high enthalpy environment during a manned Martian atmospheric entry flight can be produced for the arc heater geometry.

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A facility for the determination of total and spectral emission coefficients is described. Values of measured total emission coefficients of the different high temperature ceramics such as HfO₂, Al₂O₃, Yt₂O₃ and the metals/alloys Tungsten, TZM and PM1000 in the temperature regime of 750 - 1800 K are given. The drastic influence of the oxidation state of PM1000 on the emissivity is discussed. Additionally, results of an investigation of the influence of surface roughness and surface topology on emissivity are presented. Therefore, three SSiC samples with root-mean-squared roughness surface roughness from R_q = 0.05 to R_q = 0.66 have been prepared using common finishing operations. The tests showed that the emissivity increased about 10 % with an increase of the surface roughness even in the regime where R_q values are in the same magnitude or much smaller than the maximum emitting wavelength. Knowledge on emissivity is a basis for the analysis of thermo-chemical behavior and related properties (i.e. recombination coefficients) as the two properties influence each other. In a second step a high enthalpy inductively heated plasma wind tunnel (PWK3) is described and a methodology to derive recombination coefficients is explained. Recombination coefficients for the above mentioned materials have been determined in pure oxygen plasma. The methodology for the determination of recombination coefficients of ceramic and metallic Thermal Protection System (TPS) materials in single species gases used at IRS and its latest improvements are presented. Test results for the recombination coefficients in oxygen plasma are shown between 1469 and 2072 K.

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A method of motion reconstruction from a image streak taken by a single video camera, and a procedure to calculate aerodynamic force and moment from the motion are proposed. In this method the motion of an object is separately measured with translational and rotational motion. For the translational motion, the size and its location of a moving object is used to get the instantaneous position of the object. For the rotational motion, the attitude of the object is calculated by tracking featured points on the object. In addition to accomplish the aerodynamic measurement by this method, an evaluation was done for the measurement errors in the aerodynamic force, which are caused from image distortion by aberration of a lens. A series of the experiments for the flying ball showed the validity of the proposed method and the aerodynamic characteristics of the flying ball were analyzed using this method.

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This paper deals with an experimental investigation and visualization on the dynamics of laser induced single bubble in liquid nitrogen near an aluminum wall in cryostat. The cryogenic liquid nitrogen has the characteristic feature of the low latent heat, surface tension, and viscosity, as compared with the normal temperature water. Cavitation phenomena in cryogenic liquid may be different from the phenomena in water at normal temperature. However, there have been only a few reports on the single bubble dynamics combined with laser beam irradiation in cryogenic liquid. In this experiment the plasma is produced by focusing of Nd:YAG laser beam of second harmonics in liquid nitrogen. Laser induced bubble dynamics and interaction with the solid wall have been studied by flow visualization. This paper also shows the comparison between the dynamics of laser-induced single bubble in liquid nitrogen and in distilled water. We have clarified that the behavior of laser induced bubble in liquid nitrogen is different from that in distilled water at normal temperature.

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JAXA is currently performing studies on a Hypersonic Turbojet Experimental Vehicle, which involve a hypersonic flight test of a Small Pre-cooled Turbojet Engine. The aerodynamic performance of this airplane was examined at the JAXA hypersonic, supersonic, and transonic wind tunnel facilities. The 6-degrees-of-freedom forces and pressure distribution around the model were measured and evaluated. This airplane satisfies the lift-to-drag ratio requirement for a flight test at Mach 5. In addition, the results indicate that this airplane has longitudinal and directional static stability if the moment reference point is x/l smaller than 0.35. A separation occurs at the external expanding nozzle. Therefore, a redesign is necessary to solve these problems.

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In the re-entering phase of space vehicle into the atmosphere strong shock wave is generated in front of the hypersonic vehicle, and high enthalpy flow with radiating, non-equilibrium real gas effects is generated behind the shock wave. As well as usual convecting heat transfer, the radiative and non-equilibrium real gas heating from the shock-heated air ahead of the vehicle plays an important role on the total heat flux to the wall surface, which affects the heat shield design of the re-entering vehicle. As for the research of such non-equilibrium high enthalpy flows behind strong shock waves, several spectroscopic studies have been reported by using shock tubes or shock tunnels. The direct measurement, however, of molecular vibrational/rotational temperatures from ground state, has not been performed, and the experimental data are in quite insufficient conditions. In this paper nonlinear spectroscopic measurement is performed on the vibrational/rotational temperatures of nitrogen behind the hypervelocity shock waves from 4-5km/s by CARS method. The CARS spectroscopic signals have been successfully observed by free-piston, double-diaphragm shock tube experiments, and the measured results are compared with the simulated spectra to decide the temperatures by spectral fitting method.

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The goal of this paper is to describe the development of a hardware-in-the-loop simulation and verification environment for Field Programmable Gate Array (FPGA) based on-board computing systems. The underlying simulation environment is the Model-based Development and Verification Environment (MDVE). MDVE is an infrastructure for model-based engineering developed by EADS Astrium. A simulation environment based on MDVE was developed at the Universität Stuttgart. Recently, the demand on applying new high density FPGA technologies for innovative spacecraft on-board computing systems is rising. The small satellite “Flying Laptop” which is built by the Universität Stuttgart is the demonstrator of a FPGA-based on-board computer. In order to develop and verify the hardware and control algorithm of the computer, an extended simulation interface between MDVE and FPGA-based computing systems is established. This environment is capable of software verification and real-time simulation/verification configuration, and enables not only on-board software development but also functional real-time hardware evaluation of all the satellite components under precise space environment models. This paper describes the detailed implementation of this simulation interface and illustrates the obtained simulation results on attitude control algorithm verification and power budget calculation as well as communication timing analysis, which ensure the validity of the implementation.

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In space development, early anomaly detections of the space systems, especially detecting symptom of anomaly, and early measures against them are very important. In this study, we introduce a model acquisition method for spacecraft telemetry data using multivariate statistical tools and auto anomaly detection with the acquired model. We expand conventional limit checking to more adaptive one by considering relationship of telemetry data series using Kernel PCA. This study is aimed for reducing the cost of the modeling of spacecraft systems and improving the anomaly detecting accuracy. We applied our study to a real satellite telemetry data provided from Japan Aerospace Exploration Agency (JAXA), and detected not only anomalies but symptom of anomalies that conventional limit checking can not detect.

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