A space gravitational wave antenna, DECIGO (DECI-hertz interferometer Gravitational wave Observatory) will provide fruitful insights into the universe, particularly on dark energy, the formation mechanism of supermassive black holes, and the inflation of the universe. In the current pre-conceptual design, DECIGO will be comprised of 4 interferometer units; each interferometer unit will be realized by formation flight of three drag-free spacecraft with 1000 km separation. Since DECIGO will be an extremely challenging mission with high-precision formation flight, it is important to increase the technical feasibility before its planned launch in 2024. Thus, we are planning two milestone missions. DECIGO pathfinder (DPF) is the first milestone mission for DECIGO, and key components for DPF are being tested on ground and in orbit. In this article, we review the conceptual design and current status of DECIGO and DPF.

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A new concept of thrust modulation of solid propellant rocket motor is proposed. Some propellants cannot burn at intermediate pressure, while they can burn at lower and higher pressures. When one applies such a propellant to a motor, two combustion modes or two thrust levels are attainable without any change of the nozzle configuration. In the experiments different ignition conditions brought independent two combustion modes (low mode and high mode) in the same motor geometry. Some motors showed a natural transition from low mode to high mode. As an example, the alternative thrust levels were 50 N and 180 N. The natural transition was restricted with use of the partitioned grain. An active transition method was explored by exerting pressure perturbation through a vent hole with a ball valve. The valve system worked for the transition from high mode to low mode, but the reverse transition was not achieved well.

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To clarify the fuel gasification characteristics in a primary combustion chamber of a staged combustion hybrid rocket, the effect of fuel grain size on the regression rate of a grain was investigated experimentally. The grain size distribution in the combustion region achieved a steady state in 30 seconds burning duration. Examining fuel size distributions and fuel consumption rate at steady states, we obtained a history of fuel size and the regression rate of a grain in the combustion region. Regression rate increases with decreasing grain size. With a constant oxidizer flow rate, the regression rate is a function of grain size and independent to the initial grain size. After an initial transient the grain size decreases following the classical d-square law in droplet combustion: The square of the grain size decreases linearly with time. Although why the regression history of a grain in the combustion region follows the d-square law is not clear, this result is useful to estimate the fuel gasification rate of a staged combustion hybrid rocket.

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One-dimensional evaluation of Ramjet-mode operation was carried out on a rocket-ramjet combined cycle engine model. For simplicity, instantaneous mixing between the airflow and rocket exhaust, instantaneous heat release, and pressure recovery by a normal-shock wave were assumed. Shock wave location was so decided that the heat release at the injection (heat addition) location was to thermally-choke the combustion gas flow. By changing the injection location, it was shown that a further downstream injection resulted in a further thrust production and a further fuel flow rate requirement for choking, and a lesser specific impulse. Balancing the thrust production and the specific impulse in terms of the launch vehicle acceleration performance should be pursued. The total pressure loss within the engine model was dominated by the shock wave location, not depended on injection location and fuel flow rate, so that having shock wave penetration to further upstream location was beneficial both for thrust production in the engine and at the external nozzle.

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The purpose of this paper is to identify, for given technology levels (TRL) and mission requirements, those parameters that are critical for preliminary sizing of a hypersonic airbreathing airliner. Mission requirements will dictate a solution space of possible vehicle architecture capable of meeting cruise conditions as well as of taking-off (TO) and landing. In practice, once defined a range of cruise vehicle architectures, constraints are imposed (as to all passenger airliners), such as: 1. take off (=TO) and landing distance (so-called field length, FL): FL no longer than for the B-747-400, or 10000 ft; 2. completing TO with one engine off; 3. max acceleration at TO and climb-out (CO) = 0.4 g; 4. Hydrogen fuel (Meeting NOx emission limits (EINOx) is a further constraint not discussed in this paper). These constraints enable focusing on a realistic design out of the broad range of vehicles capable of performing the given mission. Thus a realistic vehicle must not only integrate aerodynamics and propulsion system; in fact, it is the result of many iterations in the design space, until performance and constraints are successfully achieved and met. The Gross Weight at Take Off (TOGW) was deliberately discarded as a constraint, based on Previous studies by Czysz. Typically, limiting from the beginning the TOGW leads to a vicious spiral where weight and propulsion system requirements keep growing, eventually denying convergence. In designing passenger airliners, in fact, it is the payload that is assumed fixed from the start, not the total weight. A parametric analysis of the hypersonic vehicle architecture is presented: in particular, optimal size, weight and geometrical shape are defined for different mission requirements. This analysis has shown that, it is possible to define a range of possible successful solutions for the European LAPCAT II project.

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In this paper, described is the development of a numerical simulation system, what we call “Advanced Computer Science on SRM Internal Ballistics (ACSSIB)”, for the purpose of improvement of performance and reliability of solid rocket motors (SRM). The ACSSIB system is consisting of a casting simulation code of solid propellant slurry, correlation database of local burning-rate of cured propellant in terms of local slurry flow characteristics, and a numerical code for the internal ballistics of SRM, as well as relevant hardware. This paper describes mainly the objectives, the contents of this R&D, and the output of the fiscal year of 2008.

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The concept of spiking hydrocarbon fuels such as kerosenes with liquid silicon hydrides in order to render the fuel combination hypergolic and to improve the combustion efficiency is presented and preliminarily analyzed. In view of scarcity of available data, various approaches are used, among them quantum-mechanical ab initio calculations for the thermodynamics and shock-tube measurements for the kinetics of higher, liquid silanes. Based on these results and other data, performance predictions indicate that miscible hydrocarbon/silicon hydride fuels (HC/SH) have the potential to be stored in a single tank, to be hypergolic with many oxidizers, and to yield similar, partly better specific impulses (and volume-specific impulses) than hydrocarbon fuels without silane additives. A variety of hybrid HC/SH fuel combinations seems to be accessible, which might offer the possibility to design a fuel combination with characteristics adjustable in a wide range. The current and future availability of larger amounts of liquid silanes is discussed.

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Lithium Hydride is a hydrogen-rich compound with potential application as fuel, thanks to its high density and low molecular weight. It reacts exothermically with many substances and contains H2, suggesting its use where a much higher density (compared to that of LH2) would be beneficial. In this work LiH (solid at STP) has thus investigated as potential candidate for solid fuelled scramjets (SFSCRJ). Its thermochemical properties and issues associated to its combustion in a hot supersonic stream have been investigated; results show clearly that Li, released by thermal decomposition, plays a key role in the LiH performance. In fact, above the auto-ignition point liquid Li combustion with air increases local temperature and promotes LiH decomposition. To understand quantitatively these effects, a simplified physical model describing LiH “vaporization” and combustion was built and used in simulations of a notional SCRJ chamber by means of a CFD code. Results are intriguing: an intense and stable flame zone is predicted to be present over and downstream of the grain and high temperatures (of order 2900 K) are obtainable. Moreover, specific impulse and thrust density predicted at a flight Mach = 7 are also interesting, being 10,000 m/s and 200-300 m/s, respectively.

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Laser-supported detonation (LSD) waves are important because they can generate the high pressures and high temperatures necessary for laser propulsion systems. Although CO₂ laser beams, which have a wavelength of 10.6 μm, have been considered to be one of the most powerful sources of LSD waves, a glass laser beam (1.053μm), for example, also have high power. In this study, we numerically simulated LSD waves propagating through a diatomic gas, in order to evaluate the effects of incident laser wavelength on the construction of the LSD wave. We used the physical-fluid dynamics scheme, which has been developed for simulating unsteady and nonequilibrium LSD waves propagating through hydrogen gas.

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A laser ignition microthruster has been developed for microsatellites. Thruster performances such as impulse and ignition probability were measured, using boron potassium nitrate (B/KNO₃) solid propellant ignited by a 1 W CW laser diode. The measured impulses were 60 mNs ± 15 mNs with almost 100 % ignition probability. The effect of the mixture ratios of B/KNO₃ on thruster performance was also investigated, and it was shown that mixture ratios between B/KNO₃/binder = 28/70/2 and 38/60/2 exhibited both high ignition probability and high impulse. Laser ignition thrusters designed and fabricated based on these data became the first non-conventional microthrusters on the Kouku Kousen Satellite No. 1 (KKS-1) microsatellite that was launched by a H2A rocket as one of six piggyback satellites in January 2009.

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A two-dimensional applied-field magnetoplasmadynamic thruster (2D AF-MPDT) has been developed. A strong crossed magnetic field of more than 1 T can be applied to the 2D AF-MPDT. A thrust measurement system was newly developed. Successful operation with several applied magnetic field strengths from 0 T to 1 T was achieved and acceptable thrust efficiency was obtained. The discharge voltage increased when the magnetic field was applied. The thrust efficiency improves with increasing magnetic field. An operation limit such as a voltage hash was not observed. In this paper, the results indicate that the thrust efficiency of the 2D AF-MPDT can be improved with a strong applied crossed magnetic field.

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Magnetic Sail is a propellantless propulsion system proposed for an interplanetary space flight. The propulsive force is produced by the interaction between the magnetic field artificially generated by a hoop coil equipped with the magnetic sail and the solar wind. Three-dimensional hybrid particle-in-cell simulations are performed to reproduce the plasma flow structure around the magnetic sail and to measure the propulsive force of the magnetic sail with two different coil currents. We report the characteristics of the magnetosphere, such as the profile of the magnetic field and the predicted thrust values obtained by simulations, which agree well with laboratory experiments, when simulations are carried out by considering the ion-neutral collision effect. The hybrid particle-in-cell simulation carried out without considering the collisional effect gave a thrust value of 2.3∼3.5 N, which can be applied to the thrust evaluation of the magnetic sail in a magnetosphere with size of 250∼300 km in a collisionless interplanetary space.

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In recent years there have been demands for small, simple, low-power electron sources for applications such as miniature electric propulsion systems, electrodynamic tethers, scientific applications, and spacecraft charge control. To satisfy these demands, we have started the research and development of carbon nanotube (CNT) cathodes. A laboratory model of a 0.5 mA-class CNT cathode was designed, fabricated, and tested. The emitter-to-gate separation distance and slit aperture width were varied to find effective configurations. In the selected geometrical conditions, the desired emission current of 0.5 mA was obtained at an applied voltage of approximately 600 V and with a drain current less than 10% of the emission current.

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An experimental study on coaxial laser-electromagnetic hybrid thrusters was conducted. The laser-electromagnetic hybrid thruster, consisting of a coaxial electrode configuration with an annular copper anode and carbon fiber rod cathode was used to produce laser-induced plasmas, which were further accelerated by electromagnetic force to improve thrust performance. Experimental measurement of impulse bit and mass shot was conducted. From the measurement, thrust performance showed impulse-bit of 2 ∼ 45 μNsec, momentum coupling coefficient of 5 ∼ 14 μNsec/J, specific impulse of 1000 ∼ 1400 sec and thrust efficiency of 3 ∼ 5 % for charge energies 0 ∼ 8.6 J and a laser pulse energy of 120 mJ. In addition, a significant improvement of thrust performance, could be obtained with the use of alumina propellant, which were an impulse-bit (Ibit) of 60 μNsec, a specific impulse (Isp) of 6,000 sec, and a thrust efficiency of 20% at charge energy of 8.6 J.

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Sputter monitoring system using continuous-wave cavity ring-down spectroscopy (cw-CRDS) was built for both lifetime assessment and contamination effects in Hall thrusters. We have performed measurements of sputtered manganese atoms from acceleration channel wall (stainless steel 316) in an anode layer type Hall thruster. The measurement strategy is based upon detection of manganese atoms via an absorption line from ground state at a wavelength of 403.076 nm. The path-integrated number density is 1.4 ± 0.3 ×10¹³ m^-2 at a discharge voltage of 200 V and an argon mass flow rate of 70 sccm. The number density is proportional to the discharge voltage, as expected. The number density and mass-loss have a relatively linear dependence. These results show the validity of the erosion sensor for Hall thruster lifetime estimation.

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Recently, a pulsed plasma thruster (PPT), which has advantages such as smallness, lightweight and low power consumption, has attracted attention again for the application to small satellites. Liquid propellant like water or alcohol provides higher specific impulse in comparison with conventional Teflon propellant. Nevertheless, water and alcohol propellants have problems: higher freezing point or contamination to satellites by the exhaust gas. In this study, we propose to apply DME (Di-methyl ether) to the PPT's propellant. DME can be stored as liquid, and requires no pressurant because DME has a vapor pressure of 0.6 MPa at 298 K. DME would also be potentially usable as heat transfer media, coolant, working fluid of heat pipe etc. We have designed and tested a prototyped pulse injector supplying liquid DME to discharge channels of the PPTs. Experimental results showed that the pulse injector successfully injected liquefied DME of 1.16 mg/shot, which is equivalent to the mass shot of a high power class PPT. The shape of injected liquid was dependent on the duration of the gating-pulse applied to the injector; a group of droplets were provided with a 20 ms gating-pulse duration, and a single stream with 25 ms or longer gating-pulse duration.

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In order to improve Hall thruster's lifetime performance, A new lifetime evaluation method, lifetime estimation using multilayered coated chips was developed for channel wall erosion rate distribution measurement. This method uses small chips coated with alternate channel wall material and marker. By detecting the marker emission as a signal of erosion progress, this method enables direct and rapid erosion rate distribution measurements. In this study, the capability of this new method for erosion rate distribution measurement was tested and validated. Multilayer coated chips of two kinds using different marker metals were embedded into the acceleration channel wall. The channel-wall erosion rate distribution was then measured. The measured erosion rates were, respectively, 0.90 nm/s and 0.75 nm/s at 2 mm and 4 mm upstream from the channel exit.

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In order to improve the thrust performance of a 1-mN-class miniature microwave discharge ion thruster, we investigate the dependence of inner plasma properties inside the thruster on operational conditions, krypton mass flow rate, incident microwave power and magnet field strength by laser Thomson scattering (LTS) technique. With an increase in mass flow rate, the electron temperature decreases and the electron number density increases at an incident microwave power of 16 W and the number of magnets of twelve. These results indicate that there is an optimum mass flow rate, which is 0.6 sccm in this condition. The electron number density and temperature increase with incident microwave power and is saturated at 8 W for a mass flow rate of 0.4 sccm and the number of magnets of twelve. With an increase in magnetic field strength, the electron temperature and the electron density suddenly jump from 9.8×10¹⁷ m^-3 and 5.2 eV to 1.7×10¹⁸ m^-3 and 7.3 eV at the number of magnets of thirteen at a mass flow rate of 0.4 sccm and incident microwave power of 16 W.

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With the increase of long-term space missions, the evaluation of lifetime of ion engines by numerical analyses becomes important. In order to develop a numerical code for the evaluation of ion engine lifetime, JIEDI (JAXA Ion Engine Development Initiative) tool development has been started. To evaluate the validities of boundary conditions such as upstream discharge region and downstream region conditions, a 3-dimensional full-particle code was developed. In the present study, the effects of the electron mass model introduced for shortening the calculation time were investigated. We found that there are differences in the distributions of charged particles and electric potential profiles in the downstream region among different electron masses. Consequently, the effects of electron mass on the energy peak of the ions impacting on the grid and the erosion distribution on the downstream surface of the accel grid were observed.

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Based on the success of the Japanese asteroid explorer Hayabusa, the ECR ion thruster μ10 will be installed in Hayabusa's successor, Hayabusa-2, and is scheduled to be commercialized for use in geostationary satellites within the next three years. To increase the thrust force of the μ10 as much as possible without major design changes, luminescence measurements were conducted using an optical fiber probe. The probe gave an internal view of the μ10, and it was discovered that there was plasma in the waveguide. As the plasma, the density of which is higher than the cut-off density, interferes with the transmittance of microwaves, the propellant injection location was changed. In addition to the change in propellant injection location, the grid system was also refined. These improvements increased the thrust force from 8.0 mN to 10.1 mN with a decrease in specific impulse by 40 sec from 3200 sec to 3160 sec.

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Emission data from a new geometry of Ionic Liquid Ion Sources using porous materials are presented. Ionic liquids are molten salts at room temperature which have near zero vapour pressure. When used in electrospray thrusters, these sources provide an efficient thrust mechanism and, due to their low vapor pressure, eliminate the need for complex propellant feed systems. By emitting pure ion beams rather than mixed ion/droplet emission, small high efficiency thrusters capable of providing specific impulse levels of several thousand seconds can be realized. The latest developments of these thrusters, using conical emitters micro-fabricated from porous substrates, are presented. Porous substrates provide liquid flow over a wide range of emitted currents, accommodating steady, stable emission of ion beams. Conical type emitters, fabricated on a planar porous substrate are well suited for high density 2-D arrays as they allow for both passive propellant feed through the bulk and simplified grid alignment compared with arrays formed from multiple emitter array substrates. This paper confirms that such conical emitters can provide beam currents of 100's of nA up to several μA with beams composed of pure ions with no charged droplets detected, as has previously been observed using alternative emitter geometries.

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Microfabrication of a 3x3 micro-plasmajet array with ultra-violet lasers and its 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 3x3 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 micro-plasmajet array thruster operated at 6.2 W with 1.46 mg/sec of propellant mass flow per nozzle element were 1.13 mN, 79 sec, and 24%, respectively.

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Ion thrusters are promising propulsion devices not only for standard-sized spacecraft but also for small spacecraft. However, difficulty of the development of a miniature neutralizer has been a critical problem for miniature ion thrusters. We have proposed a novel ion thruster system to solve this problem. That is a switching operation, where a plasma source can select ion beam extraction or electron emission by electrical switching. The major challenge of this study is to find the method to effectively performs both functions from one plasma source. To match the system to small spacecraft, the both modes must be conducted using low microwave power and low mass flow as low as 1.0 W and 0.15 sccm. Firstly, fundamental characteristics of electron emission were investigated using an orifice plate. Based on this result, secondly, a special grid system for the switching operation was designed and examined. This grid system has two-different-size apertures for ions and electrons respectively. This grid showed high ion beam extraction performance, that is, 4.1 mA ion beam current at 1.0 W microwave power input. It also showed much higher electron current than the old grid system, although the electron current of 1.0 mA has not reached the targeted current yet. As a result, the experiment showed that the special grid system proposed here is an effective method to realize the switching operation.

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In order to quantify the accumulation effect of projectiles in target material for sputtering, sputtering yields of carbon for pure graphite and the projectile retaining carbon target material bombarded by xenon and helium ions are calculated with a Monte Carlo code ACAT. The ACAT results have indicated that the threshold energy for sputtering is reduced due to the accumulation of the xenon atoms in graphite. Meanwhile, the threshold energy for carbon sputtering is not largely influenced by the retained helium atoms in graphite. Mass ratio of the projectile to the target atom is the important factor for the low energy sputtering yield. An empirical formula for sputtering yields for graphite retaining xenon atoms has been proposed.

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A fundamental study of newly developed PPT type rectangular laser-electromagnetic hybrid thrusters was conducted, in which laser-induced plasma was induced through laser beam irradiation onto a solid target and accelerated by electrical means instead of direct acceleration only by using a laser beam. The performance of the thrusters was evaluated by measuring the mass shot and the impulse bit, and calculating the specific impulse and the propulsion efficiency. As a result, specific impulse varied from 5,300 to 7,100 seconds with increase of charge energy to the capacitor from 2.2 to 8.6 J with a laser pulse energy of 266 mJ. The maximum thrust efficiency of 22% was obtained with the smallest charge energy case. In addition, to optimize the acceleration channel geometry, thrust performance of seven various thrusters with various geometries were compared. The maximum impulse-bit and specific impulse obtained with the 5 mm (width) x 10 mm (length) thruster were from 28 μN and from 7,200 sec for the charge energy of 8.6 J.

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Pulsed plasma thrusters (PPTs) are electrothermal and electromagnetic thrusters that produce thrust in a discharge and have great potential as space engines owing to their simple structures. A coaxial PPT with solid propellant of cavity diameter of 3 mm and capacitor energy of 8 J has been designed and operated at Gifu University. To investigate unsteady phenomena of the PPT, optical measurements have been performed with photography using a high-speed camera. In past works, the luminescence from the PPT has been observed. The results show that the luminescence from the propellant continues after discharge. In the present study, to investigate the luminescence in detail, in addition to the photography of the luminescence without a filter, photographs were taken with a band-pass filter to obtain the luminescence from only the ionized gas, C⁺ for two different cavity lengths, 25 and 15 mm. By comparing the results between the unsteady behaviors of the luminescence with and without the filter, the luminescence phenomena were investigated. The luminescence from the propellant part without the filter in the last 34 and 15 μs for the cavity lengths of 25 and 15 mm was free of luminescence from C⁺. The time when the luminescence from C⁺ ends was close to the time of the end of discharge for each cavity length. These results suggests that the measurement of the luminescence from C⁺ is effective for understanding phenomena in the cavity during discharge.

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Magnetic plasma sails have possibilities to reduce drastically mission durations of deep space explorations by using the momentum of the solar wind to produce its thrust. In order to produce a practical magnitude of the thrust, the magnetic plasma sails are required to inflate applied magnetic fields by injecting plasma from the spacecrafts. In this study, the magnetic field inflation by the plasma injection is numerically simulated by one component plasma model, which can be performed more robustly than those by hybrid simulations. The fundamental features of the injected plasma flow field and the electromagnetic field around the magnetic plasma sail are investigated, especially on the effects of the direction and the density of the plasma injection. The magnetic field can be inflated by the plasma injection in the one component plasma model. The direction and density of the plasma injection have strong impacts on the flow field and magnetic field inflation.

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Cation (Nb, Y, Al or Si) doped TiO₂ thin films with anatase phase were fabricated on SiO₂ glass by conventional Pulsed Laser Deposition method. Ionic radius of these cations is larger or smaller than that of Ti. As deposition state, amorphous was formed. Annealing for crystallization was necessary at least 500°C for 10 hours. However, the glass is difficult to be crystallized deposited at 500 °C. Suppress a formation of oxygen deficiency is possible by doping of another cations.

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This paper describes the experimental assessment of gas permeability of carbon fiber/toughened epoxy laminates under cyclic loadings as a fundamental research for the development of composite cryogenic propellant tank for reusable space transportation system. A thin-ply technique, which may contribute to the enhancement of damage resistance of CFRP laminates, is introduced herein, and the damage accumulation behaviors and gas permeability are compared between thin-ply laminates and standard laminates subjected to cyclic tension and impact.

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Thermal control materials such as white paints and germanium-coated polyimide film were evaluated with respect to their space environmental tolerance for materials selection of the Dual-frequency Precipitation Radar of the Global Precipitation Measurement satellite (GPM/DPR). Though peeling off and cracking occurred in one paint material during the thermal shock test, other paints showed good tolerance against thermal shock, atomic oxygen, and ultraviolet ray irradiation. Germanium coating on polyimide film was also verified as high atomic oxygen tolerant barrier. Comparing different thickness germanium coatings, it seems that a 1000 angstrom Germanium film has fewer defects and risk of AO undercutting than a 525 angstrom Germanium film.

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A novel deployable rhombic dodecahedral module is proposed and its mechanical characteristics are investigated. This module has a rhombic dodecahedral shape and consists of struts, cables, and vertex joints. Two types of stowed configurations, planar and linear, are applicable to the module. Cables are provided to ensure the stability of the module. The module is deployed to its objective shape from the stowed configuration by changing the lengths of the cables. In order to investigate the structural characteristics of the deployable module, mechanism analyses are carried out, and a prototype is built and tested. The results clearly indicate that the proposed rhombic dodecahedral module is deployable and that it is stable in the deployed configuration.

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Folding FEM analyses are presented to examine folding properties of a two-dimensional deployable membrane for a precise deployment simulation. A fold model of the membrane is proposed by dividing the wrapping fold process into two regions which are the folded state and the transient process. The cross-section of the folded state is assumed to be a repeating structure, and analytical procedures of the repeating structure are constructed. To investigate the mechanical properties of the crease in detail, the bending stiffness is considered in the FEM analyses. As the results of the FEM analyses, the configuration of the membrane and the contact force by the adjacent membrane are obtained quantitatively for an arbitrary layer pitch. Possible occurrence of the plastic deformation is estimated using the Mises stress in the crease. The FEM results are compared with one-dimensional approximation analyses to evaluate these results.

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Health Monitoring in aerospace applications becomes an emerging technology leading to the development of systems capable of continuously monitoring structures for damage with minimal human intervention. A promising sensing method to be applied on hot structures and thermal protection systems is the electrical resistance measurement technique, which is barely investigated up to now. This method benefits from the advantageous characteristics of self-monitoring materials, such as carbon fiber-reinforced materials. By measuring the variation of the electrical resistance of these materials information on possibly present mechanical damage can be derived. In order to set up a database on electric properties of relevant materials under relevant conditions and to perform a proof-of-concept for this health monitoring method a facility has been laid out, which allows for the measurement of the electrical resistance of thermal protection system relevant materials at temperatures up to 2000°C. First preliminary measurements of the surface resistance of a graphite sample have been performed and are presented. It has been proven necessary to make some modifications to the setup. Therefore, the remaining measurements with graphite and C/C-SiC samples are subject of further investigation which will be performed in the future.

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The Spinning solar sail is expected to be a future space exploration system. Considering the dynamic deformation of the sail membrane is an indispensable factor in designing the spacecraft. But the mathematical model of the sail membrane is complex. Computational analysis is difficult and time consuming. Therefore this has a negative impact on simulation, design and control problems. The model reduction technique is required to shorten the design period. It is a necessary step in order to put the gossamer structure to practical use. In this presentation, the dynamic property of the square type spinning solar sail is revealed, and the requirements for constructing a reduction model are revealed. Then, empirical model reduction techniques are applied to gossamer structures, and the issue of constructing a low-order model is summarized.

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Tension-stabilized structures are very useful for building large and light space/ground structures at a low cost. However it is difficult to apply a transverse damping force to such a structure. This paper describes new way to damp these structures with a tuned mass damper (TMD). It proposes a tuning method to find the optimal placementes, masses, stiffnesses, and damping coefficients of one or several TMDs attached to a cable.

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Molecular contamination is a concern for sensitive optics of astronomy satellites and earth observation satellites. The contaminants decrease the transmittance of lenses and optical filters, and the reflectance of mirrors. The absorptance of the diethylhexyl phthalate, as a model contaminant, depended on ultraviolet irradiation time and the contaminant thickness.

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To conduct remote inspection missions, the authors has proposed Morphable Beam Device (MBD) and developed an experimental device using a bendable beam without any articulated joints. In the device, a beam is deployed, enabling a wide range of shapes and lengths. In this paper, a prototype of an MBD is introduced and a beam shaping theory for two beam shaping mechanisms of slide and rotation types is discussed and verified with experiments.

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Three models of deployable membrane space structures consisting of a membrane, inflatable tubes, and connective cable networks are investigated with the aim of developing suitable modules for future hierarchical modular space structure systems on a scale of hundreds of meters. To a flat spirally folded membrane, inflatable tubes are attached in the circumferential direction, the radial direction, or both. Deployment experiments on laboratory-scale hand-made conceptual models are carried out, and their details are presented. The deployment of inflatable tubes in three different folding patterns is also studied, and smooth deployment of the tube in a modified zigzag folding pattern is demonstrated.

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This study deals with the in-plane motion of a free-floating satellite equipped with thrusters whose force directions are fixed to the satellite. The system's governing equations form nonintegrable second-order “nonholonomic” constraints due to the fixed force directions to the satellite. This paper treats a satellite with three thrusters, and at first assumes that the magnitudes of the thrusters can continuously be changed from zero to a specified positive value. Under the assumption, it is shown that the nonholonomic governing equations can be transformed into holonomic ones by a feedback of the rotational angular velocity. Then, this paper shows the procedures to control the satellite's position and attitude precisely. Finally, we restrict the magnitude of the thrusters to be constant, and discuss the control techniques for such systems.

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This paper presents a procedure to design multiple flyby missions. When the missions to observe space debris or near Earth asteroids (NEAs) are considered, it is more efficient to achieve many observations by a single spacecraft during the flight. Using the generating functions of canonical system, two-point boundary-value problem can be solved by function evaluations. Using this property, a strategy to optimize the multiple flyby sequence is proposed. This enables us to choose an optimal sequence among a long list of targets in which minimum fuel consumption is achieved. This paper focuses on the selection of sequence with pre-fixed interval time. The illustrated example using Hill two-body problem is shown.

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A ground-based spacecraft simulator is useful for experimental validation of advanced control laws. The minimization of the offset between the center of gravity and the center of rotation of the simulator is necessary for accurate simulation of a zero-gravity space environment. This study focuses on the automatic balancing of a ground-based spacecraft simulator on which moving masses required for the balancing are installed. We propose a novel automatic balancing method based on the recursive least squares approach. The proposed method is tested on a spacecraft simulator equipped with spherical air bearing. Experimental results show that accurate offset compensation and accurate space environment simulation are achieved using the proposed method.

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In this paper, we consider constellation of two orbiters around the Mars. We assume two orbiters, whose orbits are required to be orthogonal. After Mars orbit insertion, it will be needed to adjust their orbital planes appropriately for required configuration under perturbation. We discuss how to transfer S/C to the seperated orbits after simultaneous insertion, and how to maintain these orbits during the mission phase. We adopt the frozen orbit for one orbiter to fix the orbit axis, and utilize J2 perturbation on the argument of periapsis or the right ascension of ascending node for keeping the orthogonality between two orbital planes. Maneuver plan is also reported.

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The study on the post- HINODE Solar Observation Mission has been started by members in the Solar physics community. One candidate of the mission targets is the observation of the high latitude region of the Sun, which requires the injection of the space observatory (spacecraft) into the orbit largely inclined with the ecliptic plane. Reported in this paper is the trajectory design result for this orbit transfer by way of the Solar electric propulsion and the Venus/Earth gravity assists. The sequence is divided into two phases, the Venus Earth Gravity Assist (VEGA) phase and the sequential Electric Propulsion Delta-V Earth Gravity Assist (EDVEGA) phase. The designed trajectory through the sequence is provided, and it is compared with the trajectory solely using the sequential EDVEGA strategy.

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This study investigates the trajectory analysis of small solar sail demonstration spacecraft IKAROS considering the uncertainty of solar radiation pressure. Estimation of solar sail force model in space is the key factor for successful solar sail navigation because the solar sail have large uncertainty due to the flexible membrane. Since the sail wrinkles after the deployment and its surface will suffer from degradation, the solar sail force model is difficult to develop on the ground. In this paper, a practical analysis of estimating the solar sail force model from Doppler and range observable is investigated. This is demonstrated by orbit determination including parameter estimation of solar sail model. Some examples are described to investigate better parameters to estimate the solar sail force model.

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This paper presents a novel attitude control scheme for a spacecraft. It involves the use of a double-gimbal control moment gyro (CMG), wherein the rotational speed of the momentum wheel is kept constant. Tilting the spin axis of the momentum wheel around either an inner or an outer gimbal axis can generate attitude control torques in one rotational direction. In order to compensate for the absence of control torque in the other rotational direction, a combination of rotations of the inner and outer axes is designed to generate a cyclic motion about the spin axis of the momentum wheel resulting in three-axis attitude control. Analytical and numerical algorithms are developed to compute the control inputs, i.e., the tilt angles of the inner and outer gimbal axes, which are responsible for the desired change in attitude. The effectiveness of the proposed control scheme is verified by numerical simulations.

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This paper presents an approximation method of the state transition matrix for orbits around a primary body and subject to arbitrary perturbation forces. By assuming that the behavior of the perturbation sources is sufficiently slow compared with the orbital period, which covers most of practically useful cases for orbits around a primary attracting body, this method provides a functional form of the approximate state transition matrix composed of a sum of elementary analytic functions. The resulting state transition matrix is expressed in a series expansion form with a small number of constant parameter matrices and osculating orbit parameters at the initial epoch, and is valid for several tens of orbital revolutions without updating the parameters. Numerical simulations show that this method is valid for arbitrary eccentricity orbits with the semimajor axis ranging from LEO up to around 10 Earth radii when applied to Earth orbits. Due to the simplicity of the resulting approximate form, the formulation provided in this paper is suited for implementation onboard spacecraft, and a fast and iterative computation of linearized orbital dynamics with full perturbation forces.

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In recent years, medium-sized satellites that require a large torque are equipped with CMGs. In this study, we deal with limit cycle oscillation induced by CMG gimbal friction. Using a novel describing function that represents the strong nonlinearity of gimbal friction, we analytically predict the amplitude and frequency of the limit cycle. A novel controller, PID+Sign, based on PID control is proposed. The controller estimates the magnitude of the friction force by using an integrator for the gimbal angle error and reduces the limit cycle oscillation by friction compensation. Simulation and experimental results show that the proposed controller can be used to reduce the limit cycle oscillation and achieve accurate attitude control.

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This paper presents the development of large CMGs for on-ground attitude maneuver tests. The sizing of the CMGs, the overall design of the CMG system, and the results of development tests are shown. The developed CMGs are mounted on a large “dynamics simulator” test facility, which can provide an on-ground test environment of three-axis free rotational motion via air floating. The maneuver demonstration tests are performed using a cluster of four CMGs. The results show that the CMG cluster is capable of maneuvering the dynamics simulator with high agility, and the gimbals are steered with accuracy. Finally, two kinds of feedforward control laws for agile and smooth maneuvers are proposed.

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Precise relative GPS navigation is essential technology for rendezvous and formation flying of spacecraft in Low Earth Orbit. Key design issues of precise relative GPS navigation software are studied and a novel formulation is proposed for the mission which requires high accuracy when the separation distance is up to several km. The navigation filter estimates float navigation solutions by extended Kalman filter with elaborate dynamics models, and resolves ambiguities of integer carrier phase biases to achieve high accuracy fix navigation solutions. A relative navigation software with the proposed formulation is implemented and evaluated in two different ways. One is a test using spacebourne GPS receivers and a GPS signal simulator to evaluate the performance and sensitivity of the software against variation of parameters. The other is a test using actual telemetry data from Gravity Recovery and Climate Experiment (GRACE) to demonstrate the software performance. The design, implementation, and results of the evaluation are presented and discussed on this paper.

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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 a single orbit for along-track formation flying. We first introduce the control function distribution among the two spacecraft. We then focus on the translational control, design three different controllers and perform numerical studies to compare them in regard to the smallest possible variation (i.e. the achievable control accuracy) and the total delta-V. It has been shown that different controllers perform better in suppressing the effects of air drag (constant acceleration) or the J2 term (mixture of several sinusoidal waves).

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