In order to realize long-lived electric propulsion systems, we have been investigating an electrodeless plasma thruster. In our concept, high-density plasma from a compact helicon plasma source is accelerated by a magnetic nozzle for the thrust production. In order to optimize the thruster design, investigation of the thrust characteristics is needed. In this study, development of the electromagnetic thrust measurement system for compact helicon plasma thrusters and preliminary experiment of thrust measurement have been conducted. We have successfully developed a torsion-pendulum type thrust stand with measurement resolution of 10 μN. From the thrust measurement, the electromagnetic thrust increases with the Ar gas mass flow rate and plasma production power. The thrust force and the electron density jump are observed due to the discharge mode transition from the inductively coupled plasma to the helicon wave excited plasma. The electromagnetic thrust force is up to 340 μN (Ar gas mass flow rate of 1.0 mg/s, plasma production power of 400 W).

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Cost evaluation of a solar power satellite (SPS) in-space transportation using Hall thruster propulsion systems is conducted to obtain specific characteristics of Hall thrusters that yield significant transportation cost reduction. Therefore, the transportation scenario is optimized in the first place: the choice of reusing or disposing orbit transfer vehicles (OTVs), and the power of the propulsion system. The result shows that the case of reusing OTVs is superior to the case of disposing them, because the OTVs' manufacturing costs are predominant in the total cost. In addition, the power has only limited influence on the cost. Further, specific characteristics of Hall thrusters are discussed, which are necessary to achieve a target cost of $3.75 billion (300 billion yen) for the sum of two costs: the in-space transportation cost from a low Earth orbit (LEO) to a geostationary Earth orbit (GEO) and the propellant launch cost from Earth to LEO.

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Ultraviolet and mid-infrared radiations behind shock wave in Mars simulant gases are detected in order to expect the radiative heating for future Mars missions. According to the CFD calculation coupled with the radiation analysis code, SPRADIAN2, for a small-sized Martian aerocapture demonstrator, the radiation intensity from VUV wavelength region is dominant in the forebody region, while that from mid-IR wavelength region is dominant in the aftbody region. In VUV region, the absolute spectral radiances from CO 4^th positive band are evaluated by the absolute intensity calibration with a Deuterium lamp from 6.0 to 7.5 km/s. By fitting these profiles with Planck curves, equilibrium temperatures are estimated. The experimental values are less than 5% lower than the simulated ones. In mid-IR wavelength region, time histories of the radiation intensity from CO₂ ro-vibrational modes are measured by an InSb photovoltaic detector from 2.5 to 7.5 km/s. From the comparison of the relative values in equilibrium region, the IR radiation is highest at around 3.5 km/s and lowest at around 6 km/s.

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A resistance circuit ablation sensor is developed by using commercially available materials. The resistance circuit has a multi-layered configuration made of a CFRP rod, polyimide tube, and nichrome wire. The resistance of the spirally wound nichrome wire is measured via a four terminal method. The ablation sensor is embedded into a graphite test specimen and is heated in an arcjet wind tunnel flow in order to examine the operational characteristics of the developed sensor. The results show that the sensor can be used to measure the advancement of a high temperature location within the test specimen used in the present study.

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Preliminary studies are made to determine the aerodynamic and accommodation characteristics of the wind tunnel model under hypersonic rarefied flow. In the method, the wind tunnel models are exposed to the test flow in the hypersonic rarefied wind tunnel in JAXA. The displacements of the models due to the aerodynamic forces are measured during the testing. The flowfield around the wind tunnel model in the test section is also analyzed by using the numerical calculation. In the calculation, the accommodation coefficient of wind tunnel model surface is parametrically changed. The accommodation parameter of the wind tunnel model surface is then deduced from the comparison between the calculation and the measurement. For the sphere wind tunnel model, agreement between measurement and calculation could be obtained by present method with the accommodation coefficient from 0.8 to 0.9. It is also found that the determination of accommodation parameter becomes more achievable by using the displacement along the flow direction rather than that along the horizontal direction. On the other hand, further detailed modeling is required in the numerical analysis for the flat plate model. The uncertainties in the force measurement in the HRWT are discussed in this paper.

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Two-photon absorption laser induced fluorescence (TALIF) is applied to measure the recombination coefficient of atomic oxygen under several kPa pressure.ァThe recombination coefficient of oxygen on the oxide SiC is estimated as 6.4×10^-3 at the total pressure of 2,240 Pa, the surface temperature of 1,150K. From this result, we confirm that the recombination coefficient has pressure dependency.

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The effects of density, thickness and heat load upon the heat shield performance of the lightweight phenolic carbon ablators named LATS (Lightweight Ablator series for Transfer vehicle Systems) were examined quantitatively for both arc-heated test and re-entry heating conditions using a one-dimensional ablation analysis code. Thermal conductivity values of the ablator were tuned based on the arc-heated test results by matching the calculated temperatures to the measured data. Main findings are: (1) For both heating conditions, the heat shielding performance of the ablator has the same tendencies with respect to parameters of the ablator density, thickness and heat load. (2) The dependency of the back surface temperature upon the ablator density is small especially for a large ablator thickness. (3) The surface recession decreases with the increase of the density. However, the mass loss increases almost linearly with the increase of the density. (4) The ablator necessary thickness, with which the maximum back surface temperature equals to the pre-determined allowable temperature value, is nearly constant as the density changes. The ablator necessary mass increases almost linearly, with the increase of the density. (5) In considering the mechanism of nearly equal necessary thickness of the ablator, it is very important that the thermal diffusivity does not vary much with different densities of the LATS ablator. (6) From the point of the reduction of the ablator weight, the selection of a lower density ablator is more advantageous than that of a higher density ablator.

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Investigations of real gas effects of supercritical fluids on shock tube problems were numerically performed. Understandings of fluid behaviors under supercritical conditions are essential for the design of high pressure combustors. We developed a numerical code for simulations of flow fields under the supercritical conditions. Cubic types of equation of state are applied to evaluate properties of nearcritical and supercritical fluids. The present code was tested on shock tube problems, and the present result agreed well with a reference result. Comparisons of results between the real gas equations of state and the ideal gas equation of state showed differences in position of shock waves, contact discontinuities, and expansion fans due to the specific thermodynamic features of the supercritical fluids. In a condition that crosses the critical point, the results showed a large density jump with a small temperature jump at the contact discontinuity because of drastic changes of thermodynamic properties. This indicated that the real gas effects obviously appear when the initial condition is close to the critical point of the working fluids.

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Ablation experiments of silicon carbide (SiC) have been performed systematically in an air plasma freejet in our laboratory. In the experiments, ablation properties of weight loss and weight loss rate were measured, and the behavior of their ablation was observed by a video camera with neutral/density filters. The relation of the ablation properties and the behavior of ablation were discussed. The surfaces of SiC test pieces after the ablation experiments were also analyzed by electron probe micro-analyzer (EPMA). The same experiments of C were performed to compare with SiC. It was found that the behaviors of the SiC and C ablations were much different, and those materials had different ablation properties.

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The dynamic characteristics of a model of the earth-returning vehicle (HTV-R) were investigated in a supersonic wind tunnel. The free rotational technique was employed in the test. Two kinds of the test devices were prepared with different degrees of freedom. The 1-DOF device allowed the model to rotate only around a pitch axis, and the 3-DOF device allowed the model to rotate around the pitch, yaw and roll axes. The dynamic tests were carried out in a flow of Mach 1.5. Both the 1-DOF and 3-DOF tests gave the similar results on the damping and static aerodynamic properties in pitch. The rolling behavior was observed in 3-DOF test which was considered related to the uneven inertial properties.

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An inflatable re-entry vehicle is a candidate for future re-entry systems. Owing to the large area and configuration of the vehicle, it can afford a few advantages during the re-entry, descent, and landing approach, such as a decrease of aerodynamic heating and soft landing without requiring a parachute system. To investigate aerodynamic characteristics of inflatable reentry vehicle at low-Mach-number flight, wind tunnel tests were performed in JAXA Low-Speed-Wind tunnel. In this research, we investigated aerodynamic characteristics of 2 types of inflatable reentry vehicle, SMAAC and TITANS, at a low-Mach-number by using numerical simulation. Through the flow field simulation, it was indicated that the computed result of drag coefficient shows reasonable agreement with the experimental one. In the case of TITANS, the computed results showed good agreements compared with experimental results though it was confirmed that a blockage effect was observed.

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Innovation in technologies for high-speed atmospheric flights is essential for establishment of both supersonic/hypersonic and reusable space transportations. It is quite effective to verify such technologies through small-scale flight tests in practical high-speed environments, prior to installation to large-scale vehicles. Thus we are developing a small-scale supersonic flight experiment vehicle as a flying test bed. Several aerodynamic configurations are proposed and analyzed by wind tunnel tests. A twin-engine configuration with a cranked-arrow main wing is selected as the baseline of the first generation vehicle. Its flight capability is predicted by point mass analysis on the basis of aerodynamic characterization and propulsion performance estimation. In addition, a prototype vehicle with an almost equivalent configuration and dimension is designed and fabricated for verification of the subsonic flight characteristics of the experiment vehicle. Its first flight test is carried out and good flight capability is demonstrated. Furthermore a revised aerodynamic configuration with an air-turbo ramjet gas-generator cycle (ATR-GG) engine is being designed for the second generation design with improvement in flight capability at higher Mach numbers. Development of the engine, airframe structure, and autonomous guidance/control system is underway. This prospective flight experiment vehicle will be applied to flight verification of innovative fundamental technologies for high-speed atmospheric flights such as turbo-ramjet propulsion with endothermic or biomass fuels, MEMS and morphing techniques for aerodynamic control, aero-servo-elastic technologies, etc.

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The thermal analysis of a micro cubic satellite pointing to the Earth on a sun-synchronous and circular orbit has been carried out using one-nodal analysis. The altitude of the orbit is 500 km. The local time of descending node of the orbit is 11 AM. The combination of the solar absorptivity and the infrared emissivity on the surface of the satellite under which the temperature of the satellite is kept within the allowable temperature range, from 0 to 40 degree Celsius, has been clarified. As the heat capacity is larger, the number of the combinations of the solar absorptivity and the infrared emissivity increases. In order to increase the heat capacity of nano and micro satellites, the development of a heat storage material has been performed. It is desirable that the heat storage materials for micro and nano satellites have the characteristic of not phase- change but crystal transformation at heat storage because a container for heat storage material is not required. Trans-1,4- polybutadiene transforms crystal structure at the temperature of heat storage. Trans-1,4-polybutadiene is produced and the heat storage performance is measured. The produced trans-1,4-polybutadiene has the amount of heat storage of about 80 J/g at the heat storage temperature of 74 deg. C. This amount corresponds to about 70% amount of heat storage of a literature data (112 kJ/kg).The density of the produced trans-1,4-polybutadiene is 706 kg/m³.

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A heat switch provides a thermal switching (high or low thermal conductance) as appropriate according to the thermal environment of a spacecraft. It has an advantage for a thermal control for lunar rover and planetary exploration spacecraft for which it is difficult to secure sufficient electrical power. A mechanical heat switch is being developed in our group for Japanese spacecraft including the lunar rover. Required on / off thermal conductances are 1 W/K and 10 mW/K respectively with lower than 5 cm of height as a goa1. The volume expansion of paraffin was introduced for the on / off mechanism according to the result of a trade-off study including feasibility, power and expected mass. The heat switch test model we fabricated has a mass of 660 g, height of 53 mm, and outer diameter of 62 mm. In our thermal performance test, the thermal conductance was 0.86 W/K when the switch was turned on and an on / off ratio exceeding 110 was measured. In this paper, a description of the heat switch, mission targets, conceptional design study and the thermal performance result of the test model are presented.

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An experimental study on the instability of a heated meniscus in a capillary tube is conducted to explain some of the temperature oscillations observed in the operation of loop heat pipes. In the past, several mechanisms have been proposed to explain the oscillations. Oscillations resulting from the location of the liquid-vapor interface inside the condenser or insufficient fluid charging are now well understood. Marangoni-driven interface instabilities were also proposed as a potential mechanism for oscillatory behavior. The present work will focus on the Marangoni-driven instabilities. Experimental results from testing conducted in ambient and in a saturated pentane environment will be presented. The capillary structure found in heat pipes is approximated experimentally using a glass capillary tube. The behavior of the interface is observed by using a video camera. The experimental results consist of visual observation of the interface and temperature measurements. The influence of several parameters leading to the destabilization of the evaporating meniscus is discussed. Due to varying heat loads and sink conditions encountered in space applications, the LHP may experience undesirable temperature oscillations; however, these oscillations can be avoided by appropriate countermeasures. Based on the experimental results of this work, it is concluded that a wick with a smaller pore size will be more resistant to the meniscus instability; thus, the extent of low amplitude/high frequency oscillations can be limited, or in some cases, eliminated.

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To achieve reliable transmission of detonation wave to a pulse detonation engine (PDE) combustor, authors examined a combination method of "predetonator", "reflector" and "overfilling of the driver gas" experimentally. A detonation wave propagates around our reflector changing its shape through three transition processes; from planer to cylindrical, toroidal, and planar again. Here, successful transmission to self-sustainable expanding cylindrical detonation wave is key issue. The authors used high sensitivity driver gas mixture (stoichiometric H₂-O₂ mixture) for the center of the cylindrical part to make the cylindrical detonation wave transmit in target gas mixture easily. To generalize the influence of the target gas composition on the necessary overfilling radius of the driver gas mixture, we employ stoichiometric H₂-O₂ mixture diluted by nitrogen or argon as target gas mixture. In this study, we showed that the ration of width of the cylindrical path on cell size of propagation limits of both dilution cases are about 1 when the driver gas is supplied enough to create a stable cylindrical detonation wave over 50 mm. Accordingly, when the cell size of the target gas mixture becomes over comparable size to the width of the cylindrical path, the stable expanding cylindrical detonation wave does not sustain.

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This paper introduces the development of laboratory scale hybrid rocket using low melting point thermoplastics (LT) as a hybrid rocket fuel. The LT developed by Katazen Corporation have the higher regression rate comparing with Hydroxyl-Terminated-PolyButadiene (HTPB) and excellent mechanical properties, which makes them as promising candidates for hybrid rocket fuel. Therefore, the small hybrid rocket motor of the LT was prepared for the rocket launch program of the space-education using N₂O as oxidizer. The motor length is variable and combustion efficiency was evaluated at several L/Ds. From the result was confirmed that characteristic exhaust velocity (C*) was increased significantly with increasing of the characteristic length (L*). From these results were suggested that the unburned fuel droplets did not contribute to combustion as a fuel. However, large L* will become large motor and will cause low fuel volume density in the chamber. In this study, C* function was investigated by static firing test using several types of baffle plate and several shape of aft combustion chamber as to small L* and confirmed large increasing of C* under the small L*. The details are presented and discussed in this paper.

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For the first step of investigating TNT equivalency, blasting tests without igniting to fuel samples in the air atmosphere were carried out in order to investigate the fragmentation characteristic of the fuel blocks. PP, PMMA and FT-0070 blocks were blasted by No.6 electric detonator or No.6 electric detonator + P-4. As a result, it is revealed that the blocks are fragmented into smaller particles if the original block size is smaller and if the power of the high explosives is higher. However, the effect of power of the high explosives is small in the case of this study. It is revealed that PP is not fragmented so small, FT-0070 is fragmented to powder and the fragmented particles size of PMMA is between PP and FT-0070.

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In order to design hybrid rocket engines, we developed a numerical prediction method to the internal ballistics, such as the characteristic of fuel regression rate. Our model includes quasi-one-dimensional flowfield and one-dimensional thermal conduction into the solid fuel. Besides, the energy-flux balance equation at the solid fuel surface is solved to determine the regression rate. In addition to convective heat transfer, radiative heat transfer is introduced into the energy-flux balance equation in order to improve the estimation accuracy of the regression rate characteristics. The emissivity of combustion gas is estimated by Leckner's empirical correlation. Parameters related to this correlation, such as flame temperature and chemical composition at the flame, are evaluated by the chemical equilibrium calculation in which Gibbs energy minimization is implemented. The calculation results are compared with the experimental data in an open literature. As a result, it is confirmed that the calculations with the radiation model predict the regression rate dependency on oxidizer mass flux more precisely than the calculations without the radiation model. However the estimated regression rates are overestimated compared to the experimental data. It is confirmed that one of the reasons for the overestimation is the use of the wrong value of the velocity ratio in the thermochemical blowing parameter definition.

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The long-term development of ceramic rocket engine thrust chambers at the German Aerospace Center (DLR) culminates in compact designs of transpiration-cooled fibre-reinforced ceramic rocket engine chamber structures. Achievable benefits of the transpiration cooled ceramic thrust chamber are the reduction of engine mass and manufacturing cost, as well as an increased reliability and higher lifetime due to thermal cycle stability. The transpiration cooling principle however reduces the engine performance. Due to the transpiration cooling the characteristic velocity decreases with increasing coolant ratio. The goal of the chamber development is therefore to minimize the required coolant mass flow. The wall temperature can be calculated using known heat transfer correlations, for example given by Bartz, and employing a model given in literature for the reduction of the heat transfer coefficient based on coolant mass flow. By this method the required coolant mass flow ratio for different chamber diameters and pressure levels can be calculated. This paper discusses the application potential of DLR’s ceramic thrust chamber technology for high performance engines. Parametric variations of engine sizing (such as chamber pressure and diameter) are performed. For large diameters and high chamber pressures the required coolant ratio is below 1%.

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Recently, bioethanol (BE) has attracted much attention as a clean rocket engine fuel because of its reusability and environmental suitability. On the other hand, aluminum which is expected to apply on a tank or structure materials because of its importance in weight decreasing shows corrosion property toward alcohol. In the present research, we confirmed the corrosion of aluminum alloys (A6061) by means of material compatibility experiments using BE. However, we succeeded the anti-corrosive effect for A6061 by alumite treatment and Ni coatings. Furthermore, we also examined the compatibility of non-metal materials such as FRPs and O-ring toward BE. We found that these non-metal materials have no compatibility toward BE but Ni-coated CFRP showed compatibility.

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The flame stabilization method using thermo-acoustic oscillation was demonstrated and the effectiveness of the sound for the flame stabilization was validated by investigation of the lift-off height change in a standing acoustic field which is a simple acoustic system. The insight of stabilization mechanism, focused on the variation of temperature, flow fields and reaction rate, was validated by CFD tool. To form the stabilized lifted diffusion flame and deal with as two dimensionally, 2D flame which formed in the mixing layer of the air and the fuel was employed for the analyzed flame. Thermo-acoustic oscillation was generated by forming flames in two stainless resonance tubes which are placed face to face. This oscillation was applied to the flame which is held by a cylindrical flame holder at the center of stainless resonance tube’s open ends. Standing acoustic field was generated by two loud speakers which are driven at the resonance frequency in a stainless resonance tube. As a result, the flame was stabilized by thermo-acoustic oscillation without the flame holder. It was found that the flame lift-off height decreased with sound both from the experimental result and the CFD result. Flow speed normal to the sound propagation direction and temperature oscillated with sound.

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In this paper, we extend the exiting method to conduct combustion stability analysis in axial injection hybrid rocket engines to apply for vortex injection systems by introducing the swirl strength as a new input of the linear model, extending the quasi one dimensional flows to axisymmetric three dimensional flows and estimating axial skin-friction coefficient in swirl flows and skin-friction shear stress with blowing by contrasting one with no blowing. Next, to validate the accuracy of the quasi-steady boundary layer combustion model we compare regression rates derived from this model with ones derived from experiments. Then, linearizing the quasi-steady model, we construct unsteady models. Finally, we couple the unsteady boundary combustion model with the thermal time lags model in solid fuels and conduct stability analysis. It is estimated that swirl systems also have a set of unstable poles of natural oscillations and swirl and can be more unstable than axial ones.

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Cavitation may cause a severe damage to rocket engine turbo-pump, and appropriate handling of cavitation is one of the key technologies for rocket engine development. Numerical simulation is quite useful technique, however, simulation result strongly relies on cavitation model. Cavitation flow is essentially strong unsteady and multi scale phenomena, scales of bubbles vary in size from small to large. Therefore, prediction of cavitation by numerical simulation is difficult. Many cavitation models have been proposed and researched. However, direct interface tracking approach has not been applied to cavitating flow and the model characteristic is not fully understood. This research shows direct interface tracking method can properly capture both cloud cavitation and super cavitation and grid resolution affects lift coefficient.

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Combination of a scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines (the combined system termed as Rocket Based Combined Cycle engine) are expected to be the most effective propulsion system for Booster stage of space launch vehicles. At hypersonic regime, it will be operated at rather high rocket engine output for final acceleration with some Isp gains due to air-breathing effects. In this regime, attaining thrust at this high-speed regime becomes very difficult, so that parallel injection of the fuel for scramjet combustion is favorable as the momentum of the injection can contribute to the thrust production. Thus, embedded rocket chamber was supposed to the operated as fuel rich gas generator at very high output. This configuration was tested at simulated flight Mach number of 11 at High Enthalpy Shock Tunnel (HIEST) with detonation tube as the source of the simulated rocket exhaust. However, combustion of the residual fuel in the rocket exhaust with airflow could not be attained in the past study, so that modification was carried out. Combustion was attained within the combustor, however, insufficient mixing was found to responsible for the rather low combustion efficiency, finding being from complementary mixing test at a blow-down type tunnel.

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A preliminary study on induction phenomena of laser-ignition was conducted. For observation of the initial phenomena, an experimental study on effects of a focused laser pulse (Nd:YAG, 335mJ/pulse, pulse width 5nsec) into a hydrogen-air mixture was conducted. Temporal evolutions of typical line spectra of a laser-induced plasma of the mixture were measured with the photodiode through specific band-pass filters for each spectrum for N⁺, H, and O, and also with a time-resolved ICCD spectroscopic analyzer. As the results, it was found that phenomena occurring after laser irradiation up to combustion completion were divided into three processes; i) absorption process of an incident laser pulse (2 x 10^-9~ 10^-8 sec), ii) plasma formation process (ionic emission by 400 nsec followed by atomic emission) (10^-8 ~ 10^-6 sec), and iii) ignition (10^-5 ~ 10^-4 sec), having various scales of the characteristic times. In addition, temporal evolution of electronic excitation temperature of N⁺ ions was estimated from the Boltzmann plots. It was shown that the temperature was about 25,000 K at initial 200 ~ 400 nsec, and then gradually decreased 20,000 K at 900 nsec. Moreover, electron densities were also estimated from the Stark broadening widths of H_α and H_β spectra. Temporal evolution of the electron density showed the electron density of about 4 x 10¹⁷ cm^-3 at initial 300 nsec, and then, down to 1.3 x 10¹⁷ cm^-3 at 900 nsec.

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A deep-throttable rocket engine using hydrocarbon fuel is needed an essential element of next generation space propulsion systems. However, deep-thlottling capability of a rocket engine is restricted within narrow range by some factors. Then, we proposed a gas/gas injection type rocket engine to solve the problems attending on deep-throttling. On the other hand, using this type injection system made some extra problems such as the balance between chamber cooling and gasification of the propellant, the power load of auxiliary machinery, and the range of turbopump operating point. In this paper, we adopt the propellant combination of Ethanol/Oxygen and analyzed the system balance and engine performance.

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Engine system analysis upon a Rocket-Based Combined Cycle (RBCC) engine for the booster stage of a TSTO launch vehicle was conducted, to evaluate the benefits of hypersonic air-breathing technology upon the performance of the launch vehicle. Hydrocarbon fuel (namely ethanol) and LOx was selected as the propellant of the vehicle to enable easy handling and good packaging of high-density fuel into the airplane-like vehicle shape. First, parametric study on the ramjet-duct portion of the RBCC engine was conducted for fixed embedded rocket engine parameters. Then the rocket engine parameters were varied for the fixed ramjet-duct parameters to attain best performance with a fixed ramjet-duct projected area. The analysis method and primary results are herein reported.

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A new capacitance type void fraction sensor was designed, produced, and tested. This sensor applies the difference between the relative permittivity ε of gaseous hydrogen (ε= 1.0) and that of liquid hydrogen (ε = 1.2). Following the sensor verification test using light diesel oil and air, a cryogenic experiment using liquid nitrogen was conducted. As a result, the void fraction measured by the sensor showed good agreement with the result obtained by an optical analysis using a high speed camera. One of the key problems on the sensor is an existence of the temperature drift caused by the change of the relative permittivity of the glass tube. In order to reduce the temperature drift, the length of electrodes and material of tubes were changed. Combination of short arc length electrodes and iupilon tube is ideal for reducing the unwanted temperature drift. The sensor which has shorter electrodes reduces the quantity of the temperature drift by 63% compared to the original sensor.

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Thrust components imparted by an electrodeless helicon plasma thruster including a magnetic nozzle are experimentally measured and analytically derived from momentum equations assuming a cold ion, negligible electron inertia and radial ion inertia. The magnetic nozzle of the maximum field strength of ~180 Gauss is provided by a solenoid coil located near the thruster exit. It is found that the total thrust can be given by the sum of an electron pressure force onto the source boundary and a Lorentz force generated by a radial magnetic field and an azimuthal electron diamagnetic drift current in the magnetic nozzle. Further, a permanent magnet helicon plasma thruster for more efficient and compact plasma thruster is also designed and tested.

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A modified magnetic multipolar plasma source is developed and tested for application as an ion thruster for space propulsion. The configuration consists of eight samarium-cobalt bar-magnets, symmetrically surrounding a cylindrical discharge volume of 21 mm diameter, with their magnetization in the radial direction. A donut magnet is placed at one axial end of the discharge volume with its magnetization along the cylindrical axis. A short monopole antenna, coaxial with the donut magnet is used for launching microwaves in the discharge volume. The resulting magnetic configuration due to the radial and axial magnetic fields, help in generation and sustenance of microwave induced electron cyclotron resonance plasma. Ions are extracted from the plasma by high voltage grids, attached to the end, opposite to the donut magnet. Two different magnetic configurations obtained by reversing the magnetization direction of the donut magnet are compared. Three dimensional magnetic field simulations and ion saturation current measurements at various microwave frequencies are used to determine the optimum magnetic configuration. Ion beam current extraction measurements, using the grids is carried out. Typical ion beam current of 9.9 mA is obtained, with calculated thrust of 0.63 mN at 8 W for 1600 MHz, with a Xenon mass flow rate of 20 μg/s.

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Two-dimensional axisymmetric particle-in-cell simulations with a Monte Carlo collision algorithm, external circuit analysis, and measurements of plasma densities and ion beam currents have been conducted for a micro RF ion thruster. The plasma source of the thruster is 5.0 mm in radius and 10 mm in length with a five-turn helical coil around a cylindrical quartz chamber, where the plasma is generated at low RF powers (< 10 W). The numerical results have indicated that high RF frequencies (> 100 MHz) lead to a high coupling efficiency between the RF power and the plasma, although too high RF frequencies (> 300 MHz) degrade the efficiency. The optimum gas flow rate for ion beam extraction is determined to be 10 μg/s. These results are qualitatively in good agreement with the experimental results, which validates our numerical model, and thus the model would be useful to estimate the performance of the micro RF ion thruster.

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For a 500-to-900-kW-class steady-state self-field magnetoplasmadynamic (MPD) thruster, the thermal design is conducted with a combination of magnetohydrodynamic (MHD) and thermal analyses, where a heat flux evaluated from the MHD analysis is imposed on the electrode as a boundary condition in the thermal analysis. The increase in anode-to-cathode radius ratio improves the thrust performance, but can simultaneously raise the temperature locally at the anode exit edge and the cathode tip due to the concentration of the discharge current and/or the insufficient heat removal. It is suggested, however, that a thruster without electrode melting should be realizable even at such a high input power by setting an appropriate cathode radius and enhancing heat removal from the electrodes by means of heat pipe, although the achievable performance of the thruster depends on the assumption of the temperature of cathode root.

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Numerical studies were conducted on a small ion engine system called μ1, which was developed for micro-spacecraft propulsion. For its miniaturized application, the μ1 ion engine is very small in size and its beam current and propellant utilization efficiency are lower than those of standard-sized ion engines. As it operates in a low perveance range with a molybdenum two-grid system, the acceleration grid experiences very severe erosion, and the extraction ion beam performance degrades substantially because of the scattering of highly diverted ions and neutrals and the impact of backstreaming charge exchange ions. In this study, the decrease in the propellant utilization efficiency and the lifetime of the μ1 ion engine were firstly assessed using the JIEDI tool, considering the effect of the change in extraction ion beam performance. Lifetime analysis showed that the μ1 grid system can operate over its required lifetime of 10000 h, however, the propellant utilization efficiency decreased significantly from 0.42 at the beginning of life to 0.30 at the grid structural failure of 27000 h. The agreement of grid erosion patterns after the first 150-h of operation between the simulation and the experiment supports the validity of the analysis. In addition to the lifetime assessment, the grid parameters were parametrically searched to achieve both high propulsion performance and long life. One of the strategies is to double the diameters of the screen and accelerator grid holes as well as the accelerator grid thickness, keeping the separation distance between the screen and accelerator grids unchanged.

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Solar sail is a spacecraft that has a large-scale membrane to utilize the solar radiation pressure for its thrust. Hence, maintaining the membrane structure during space flight is a critical issue to keep thrust performance of the spacecraft. In this paper, we focused on the electrostatic force due to spacecraft charging on the membrane as one of the possible factor to cause the deformation of the membrane structure. We had estimated the electrostatic force via charging simulation for the IKAROS spacecraft in solar wind plasma at 1.0 AU. We had also made a structural analysis for the deployed membrane of IKAROS with the electrostatic force. The structural analysis showed that the electrostatic force could hardly affect the membrane structure in this case.

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Magneto-Plasma Sail (MPS) is a sail space propulsion system to accelerate the spacecraft in the solar wind direction by capturing the solar wind energy. MPS’s significant feature is it’s a utilization of magnetic field inflation process, in which a small magnetosphere by on-board coils is inflated by low-speed plasma released from the spacecraft. In this study, a parametric survey is conducted using the axis-symmetric two-dimensional magnetohydrodynamic (MHD) simulation to investigate the relation between the thrust characteristics and the injected plasma parameters (the mass flow rate and the thermal beta value at the injection point). As a result, the thrust gain, which is defined as the ratio of the MPS spacecraft thrust force to the Magnetic Sail thrust force, and the specific impulse have the maximum value against the thermal beta value, and in addition, increases with lower mass flow rate. It is indicated that the maximum value of the thrust gain is 3.77 when the thermal beta value is 25 and the mass flow rate is 100 in the non-dimensional parameter (when a magnetic moment of the spacecraft is assumed 6.28x10⁷ Wb・m, the thrust force is 60 N and the mass flow rate is 3.6 mg/s).

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Magnetoplasma Sail (MPS) is one of the next generation in-space propulsion systems that utilize the interaction between the solar wind and the magnetosphere around a spacecraft inflated by the plasma injection. In order to evaluate the thrust characteristics of MPS with plasma injection, the thrust measurement of a scale-model MPS (10-100 N class) was conducted in the laboratory using a pendulum-type thrust stand. The thrust was observed to increase by plasma injection, from 0.04 N to 0.1 N. The thrust gain, the ratio between the thrust with plasma injection and the thrust without plasma injection, was 2.5 (maximum thrust gain). The thrust gain of MPS was observed to increase with increasing the β_k value at the plasma injection point, and to our knowledge, the thrust increase by plasma injection of MPS was observed for the first time in this study.

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In order to evaluate neutralization phenomena of ion thrusters, ion beams and plasmas were observed with a two-dimensional visualized ion thruster and a miniature microwave neutralizer. This neutralizer was appropriated for this study because of the low consumption power, the low carrier gas flow rate, the low luminescence and the wide controllable range of emission current. Through the observation and electrical measurements, it was confirmed that imperfect neutralization affected the ion beam optics, caused stagnation of emitted ions, refluxed the ions into the grid system, and increased space potential in the downstream region resulting in the formation of “virtual anode.” Since the shapes of observed virtual anodes were semicircular, it was also confirmed that two- or three-dimensional evaluation is necessary for the more detail understandings of neutralization phenomena.

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An Electrodynamic tether (EDT) is an attractive propulsion device for active debris removal systems. One of the key components of the EDT system is an electron emission device, and we have studied a field emission cathode (FEC) using carbon nanotubes (CNTs) because of its simplicity and potential capabilities. Since EDT systems are operated in low earth orbit (LEO), the CNTs in the FEC may be affected by atomic oxygen (AO), so the effect of AO irradiation on the FEC was studied. We conducted AO irradiation tests on FECs using laser detonation beam facilities and compared the pre- and post-irradiation electron emission characteristics. As a result, the FEC could not emit electrons when the total AO fluence of 3×10²⁰ /cm² was irradiated perpendicular to the emission surface. In this case, CNTs on the emitter surface disappeared by AO irradiation. When the irradiation direction was parallel to the emission surface, on the other hand, the FEC could emit electrons after the irradiation although the required voltage increased up to 1.5 times as high as that of pre-irradiation condition. This comparison indicated that the direction of AO irradiation has strong effect on the performance degradation of the FEC on orbit.

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The JIEDI (JAXA's Ion Engine Development Initiative) tool was developed to assess the ion acceleration grid erosion of an ion thruster. The validation and the sensitivity analysis of the input parameters of the JIEDI tool are conducted in this paper. We compare the simulation results of the JIEDI tool and show that it successfully reproduces the full lifetime test of the μ10 prototype model. Through sensitivity analysis, we found that the sticking factor (the ratio of sputtered materials sticking onto a grid surface to sputtered materials hitting the grid surface) is the most sensitive input parameter when estimating the accelerator grid erosion and electron backstreaming time. The accelerator grid mass loss was 189% larger for a sticking factor of 0 compared to that determined for a sticking factor of 0.78. For the worst case scenario of grid erosion (without re-deposition (sticking factor = 0)), the uncertainty in the neutral mass flow rate through the grid holes is important when estimating the accelerator grid erosion. A 25% change in the neutral mass flow rate caused by a 6% change in the propellant utilization efficiency, corresponds to about a 20% change in the accelerator grid mass loss as well as the increasing rate of minimum potential on the axis. In contrast to the accelerator grid erosion, the uncertainties in the discharge voltage and grid gap (that affect the trajectory of the mainstream ions), are important when estimating decelerator grid erosion. A 40% change in the discharge voltage corresponds to about a 90% change in the decelerator grid mass loss.

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Although investigations of propellant distribution within the ion thruster are necessary to improve its performance, it is difficult to measure the distribution because the propellant is neutral: a non-charged particle. For this study, a low-pressure measurement system with a differential pressure gauge was developed for the evaluation. Through the experiments, it has been confirmed that the measurement error was less than 10% in the pressure range of 0.01 Pa to 1 Pa, and that the neutral particles are distributed almost uniformly within the discharge chamber in the annular propellant-supply case. The pressure measurement system might be useful for the evaluation of the neutral density distribution not only in an ion thruster, but also in other electric propulsion systems.

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Two-dimensional simulations of Lissajous acceleration were conducted by a code based on Particle-In-Cell (PIC) method and Monte Carlo Collision (MCC) method for understanding plasma motion inaccelerationarea of an electrodeless plasma thruster and for finding the optimal condition. Obtained results showthat azimuthal current depends on a ratio of electron drift radius to plasma region length, and peak value is in proportion to AC frequency. To eliminate a disturbance on rotating mode due to electron cyclotron motion, the cyclotron frequency should be much higher than the AC frequency. When the high density plasma reduces electric field penetration, the azimuthal current is suppressed in a low level. In order to improve the electric field penetration, we should apply the high magnetic field and the high AC frequency. When a ratio of the collision frequency between electron and neutral argon to the cyclotron frequency is lower than 0.01, collisional loss of the azimuthal current is little or nothing.

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This paper proposes direct focusing of repetitive high-power laser pulses on an arbitrary surface of a vehicle in the atmosphere to generate a blast wave at each pulse and push the vehicle along an impulse or thrust vector. Fundamental research was conducted on the interaction between a focused high-power laser pulse or blast wave and a surface with arbitrary curvatures. The characteristics of the impulse or thrust vector generation on the surface were numerically simulated. For simplicity, some fundamental shapes of the surface or vehicle were assumed for the simulation (i.e., planar and semicircular bodies) to examine shock–surface interactions. The impulse vector characteristics were also investigated experimentally. The results showed that the composition of the impulse vector (or each component) reached about 90% of the total impulse in the initial 10 μs. In this duration, a significant high-pressure region induced by the shockwave was localized near the laser irradiation spot; it acted on the surface vectors to induce a primary thrust. Based on these results, the directions of impulse vectors are primarily determined by the local surface vectors of the laser irradiation spots.

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In recent years, small spacecraft missions are increasingly utilized in wide variety of applications. They promote technology demonstration and space development by the short-term and low-cost development. Further progress of small spacecraft needs small thrusters—propulsion system with light weight, low operating power, and high efficiency. A miniature and low-power microwave discharge ion thruster is one of the candidates. The miniature ion thruster can operate with 1 W microwave power and it has specific impulse higher than 1000 s. However, microwave behaviors in the transmission line of complex geometry and discharge chamber with plasma absorption make it difficult to modify or improve the thruster. It is important to understand overall behavior of the microwave from the input connector to the discharge chamber. In this study, 3D microwave analysis was conducted in the transmission line and discharge chamber and it was revealed that impedance mismatching, even if it is short-distance, has an important effect, making resonance. Secondly, that resonance was formulated using a signal flow graph, and reflectivity of the plasma was experimentally measured using that formulation. A set of the techniques makes it possible to estimate the net plasma absorption of the device with different transmission lines or to optimize the transmission line for a newly designed plasma source.

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To present a feasibility of high-altitude flight of a laser-propelled vehicle at supersonic speed, we have developed a flight simulator which has fluid-orbit coupling calculation module to reproduce impulsive flight reaction driven by blast waves. By high-power energy transmission through arrayed lasers together with the genetic algorithm (GA) controlled sub-laser, the supersonic flight is successfully achieved in the simulation for 32.5-g vehicle, while the angular offset should be suppressed as small as possible. Rather than translational position, controlling angular offsets by the GA operation is especially important to attain the km-order flight on the premise of the active control. Additionally, the vehicle weight, the vehicle size, and the input energy are scaled up to assess the stable flight of 10-kg vehicle. The active control technique has enough possibility to launch kg-order vehicle at supersonic regime with the optimized beaming strategy.

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A space inflatable extension mast designed on an idea of structural rigidization simulation system has been projected and provided to verify its engineering technology and to obtain the structural property for a long-term operation in space environment. The inflatable mast extended successfully in orbit, and has sent the telemetry data for more than seven months, so far. The experiment progress meets both its minimum and full success criteria for this mast. The simulation model of this mast is made up based upon the ground test, and predicts the natural frequency in orbit. Lightweight extendible masts are fundamental and essential structural elements to construct space structures; therefore a pivotal first step of conditions to utilize space inflatable structures has been actually achieved.

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This paper discusses a contact dynamics which is happened when a massive spacecraft named H-II Transfer Vehicle (HTV) is captured by a Space Station Remote Manipulator System (SSRMS). If the point of the first contact is different from the planned one, then HTV's motion will be completely change and may collide with ISS. Therefore, if we can predict motion of HTV depending on the telemetry data in real operation time, it will be useful to assure safety of the mission. To develop a dynamics simulator, at first, we studied telemetry data obtained in the 1st HTV (HTV-1) capture operation. Then, we developed a simple models with some parameters. In that phase, some preliminary experiment and deliberation in kinematics was conducted to model mechanical behavior and determine parameters. Then, optimization of parameters was conducted to identify some unknown parameters of contact dynamics based on telemetry data. And with following the course of optimization, we discuss key parameters which have large effect on dynamic HTV's behavior and can be used in developing a low calculation-cost model. Finally, we apply key parameters and compare the results with an actual telemetry data of HTV-2 to verify its efficiency.

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This paper addresses the design of a rib-stiffened shell antenna with the adaptive structure system by simultaneous optimization of the structure and the actuators to improve the surface accuracy of the space antenna. A rib-stiffened shell antenna is proposed for the antenna structure concept with the adaptive structure system. The simultaneous design optimization is performed for the 1/6 element model of rib-stiffened shell antenna. The results of the design optimization show that the higher precision surface can be obtained by the simultaneous optimum design than the individual optimum design. The feasibility of the simultaneous optimum design is examined experimentally. The surface error obtained by the experiment is in agreement with the surface error of the design optimization, and thus, the feasibility of the simultaneous optimum design is verified.

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A novel method for controlling the shape of an antenna reflector based on surface accuracy analyses using an optical shape measurement system such as a stereo vision or photogrammetry system is developed, and its feasibility is demonstrated. In the control procedure, intentional deformations are added to a reflector surface by using surface adjustment mechanisms, and the corresponding changes in the surface accuracy of the measured surface shape are analyzed using an optical shape measurement system. Then, the control inputs for the shape correction of the deformed reflector are directly determined from the information about the changes in the surface accuracy. Some numerical simulations and experiments are performed, and the feasibility of the developed method is demonstrated. The results of the investigation show that deformations in the reflector are properly corrected by the developed method when its shape is measured at appropriate points.

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To meet severe requirement of recent large and precise space structures, extremely high structural shape stability on orbit is needed. The low thermal deformation is one of the most important characteristics. This paper describes a novel technique for spacecraft's thermal deformation test on the ground. The required measurement accuracy in the test becomes higher and higher depending on the required structural shape stability. In the proposed test technique, the heat load is applied rapidly and locally to shorten test time. Due to the short test time and restricted applied heat load, the environmental changes are suppressed. As a result, good measurement accuracy can be achieved. Furthermore, the transitional thermal deformation is measured continuously so that we identify the cause of thermal deformation through the correlation between temperatures and measured deformations. The effectiveness of proposed technique is shown through the test results of ASTRO-H which is a next X-ray observation satellite.

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