Hybrid rockets that use glycidyl azide polymer (GAP) as a solid fuel were studied. The linear burning rate spectrum of GAP was broadened when it was diluted it with polyethylene glycol (PEG). Self-combustible mixtures were used for the gas hybrid rocket motor; non-self-combustible mixtures were used for the traditional hybrid motor. Both were tested as functions of various experimental parameters, and the results of the traditional hybrid rocket are summarized in this paper. In particular, the fuel regression rates of GAP50PEG50 and GAP60PEG40 fuels were investigated, and the results indicate that the fuel regression rate of the fuel with high GAP content showed pressure dependability. Simultaneous measurements of the surface regression rate, which has attracted great interest, were conducted using ultrasonics in a special chamber, the results for the instantaneous surface regression rate during combustion were obtained. The obtained results are presented, and on the basis of experimental observations, a flame model of the traditional hybrid motor is discussed.

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We had numerically analyzed the charged particle profiles and the potential structure around a solar sail at 1.0 AU. The quantitative estimation of these issues can be of importance for the payload design of a solar sail such as the location of onboard scientific instruments and solar arrays. In addition, estimation of the electrostatic force on the IKAROS spacecraft was made to determine whether the force was significant for the deformation of the membrane. A 3-D electrostatic, full-Particle-In-Cell code was used to study precise charged particle behaviours around a solar sail, and MUSCAT, a spacecraft charging analysis tool, was additionally used to obtain differential voltage of the sail. The numerical results showed that a wake potential was formed due to a large ion wake in the downstream of a sail obstructed the diffusion of the photoelectrons to the downstream surface. The size of the photoelectron cloud around a sail was estimated to be 17.5 m in the upstream hemisphere at 1.0 AU. The floating potential of the sail was +8.3 V, where double Maxwellian photoelectron spectrum model was adopted in the computation. The reduction of the electron sheath due to the photoelectron cloud was recognized. The differential voltage on the insulator surface of the sail of -15.8 V was obtained by the MUSCAT computation, that showed the charging of a solar sail itself was not serious in this environment but would affect the photoelectron diffusion and the wake potential. The electrostatic force at the numerical grid on the membrane assuming the IKAROS spacecraft was estimated under the observed real plasma environment. The magnitude of the force was of 10^-8 N at the edge of the insulator that could not be negligible to a small-scale deformation on the membrane during the flight.

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This paper describes a modeling of attitude dynamics of spinning solar sail spacecraft under influence of solar radiation pressure (SRP). This method is verified and actually exploited in the attitude and trajectory guidance operation of Japanese interplanetary solar sail demonstration spacecraft IKAROS. IKAROS shows a unique attitude behavior due to strong SRP effect. This paper introduces a new generalized dynamics model called Generalized Spinning Sail Model (GSSM), which clearly explains physics behind the observed phenomena with only three parameters. Precise understanding of attitude dynamics through the GSSM led to 6 months of world first interplanetary trajectory guidance of solar sail-craft toward Venus. The GSSM also contributed to realize a zero-fuel spin axis maintenance for most of the flight path before the Venus flyby. In this paper, an overview of IKAROS attitude and trajectory control operation in the cruising phase, derivation of the proposed model and its implementation to the actual IKAROS operation are shown.

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The Japan Aerospace Exploration Agency (JAXA) makes the world’s first solar power sail demonstration of photon propulsion and thin film solar power generation during its interplanetary cruise by IKAROS (Interplanetary Kite-craft Accelerated by Radiation Of the Sun). It deployed and spans a membrane of 20 meters in diameter taking the advantage of the spin centrifugal force. It accelerates and controls the orbit using solar radiation pressure successfully. This is the first actual solar sail flying an interplanetary voyage. This paper presents the summary of development and operation of IKAROS and introduces the outline of the extended solar power sail mission toward Jupiter and Trojan asteroids.

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Japan Exploration Agency (JAXA) launched a powered solar sail “Interplanetary Kite-craft Accelerated by Radiation Of the Sun (IKAROS)” on May 21, 2010. One of the primal technologies demonstrated at IKAROS is the spin-deployment of the sail whose diameter is 20 m class. After the launch, two-step deployment operation was performed and successful expansion of the sail was confirmed. This paper shows the flight data and observed dynamic motion during the deployment. At the quasi-static first stage deployment, the spin rate of main body shows little oscillation after each step and then damped quickly. The damping ratio of the spin rate after a later step of deployment is estimated by curve-fitting to be 0.0127. At the dynamic second stage deployment, the nutation motion is maintained within ±1.5 deg/s and the spin rate of main body is converged quickly about 60s after the deployment start. These flight result and observed dynamic motion during the deployment are compared with the results of numerical simulations using multi-particle model. These results show that the multi-particle model can simulate the global behavior of membrane sufficiently except the dumping motion of in-plane oscillation between the main body and the expanded membrane.

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Interplanetary Kite-craft Accelerated by Radiation Of the Sun (IKAROS) was launched in May, 2010. IKAROS deployed its membrane in space and generated electricity using its thin-film solar cells on the sail membrane. This paper focused on the fabrication process of the flight model of the thin-film solar cell array. IKAROS had 144 modules of thin-film solar cells with 360 W in total output power. One module was fabricated by stacking 4 layers; protective film, thin-film solar cell, anti-warpage layer and sail membrane. The size of one module is 220 mm x 300 mm x 152.5 μm. They were bonded each other with silicon glue. The connection between module and module or between module and sail membrane were also bonded with silicon glue. The fabricated thin-film solar cell array was examined from the point of possible damages during the fabrication process, resulting in no critical damage to the thin-film solar cell array. Moreover, it successfully generated in the orbit, in comparison to the assumed power generation from the results of the generation test before the launch.

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This paper introduces the application of low melting point thermoplastics (LT) to hybrid rocket fuel. LT made by Katazen Corporation has an excellent mechanical property comparing with other thermoplastics and prospect of high surface regression rate because it has a similar physical property with low melting point of paraffin fuel which has high regression rate probably due to the entrainment mass transfer mechanism that droplets continuously depart out of the surface melt layer. Several different types of LT developed by Katazen Corporation for this use have been evaluated in the measurements of regression rate, mechanical properties These results show the LTs have the higher regression rate and better mechanical properties comparing with conventional hybrid rocket fuels. Observation was also made using a small 2D combustor, and the entrainment mass transfer mechanism is confirmed with the LT fuels.

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Turbulent mixing field of hydrogen jets injected in supersonic streamwise vortices was numerically investigated using large-eddy simulation (LES). The jet and the streamwise vortices were introduced in a similar and simplified manner to those introduced by the alternating-wedge (AW) strut injector. The LES reproduced the large-scale wavy jet structure containing small-scale vortices as those observed in the past experiment of the AW strut. The streamwise vortices strongly promoted the jet spreading and the turbulent transition. Downstream of the turbulent transition, the fully developed turbulent state was achieved in the wavy jet structure. The instantaneous and time-averaged mixing efficiencies were quantitatively evaluated and compared each other to investigate the intermittency of mixing state. The mixing efficiency rapidly increased by the turbulent transition. The jet mixing in the streamwise vortices had lower intermittency compared to the transverse jet from the wall. The effects of the incident shock wave across the hydrogen jet were also investigated in detail. The turbulent properties such as the fluctuation intensity and the power spectrum hardly changed by the crossing of oblique shock wave. The mixing efficiency slightly increased after the oblique shock wave passed through the hydrogen jet.

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This paper discusses a 1-N class space propulsion thruster using SHP163 monopropellant and arc discharge assisted combustion. SHP163 is a hydroxyl-ammonium nitrate (HAN)-based monopropellant, which has relatively low toxicity and reactivity to materials such as metal and polymers, and can be stored in a liquid form due to its relatively low freezing point. Arc discharge, supporting a SHP163 exothermic chemical reaction, is used as an alternative to conventional particulate catalysts. Part of the pressurant is used as working fluid for the arcjet. Tests using prototyped thrusters show that combustion started and was successfully sustained with the assistance of a 1-kW class arc discharge. Thrust chamber pressure measurements yielded a characteristic exhaust velocity efficiency of 90 % at a specific power of 2.8 MJ/kg.

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A lab-scale combustion wind tunnel was developed for investigation of low-pressure ignition and flame holding in a sub-scale pre-cooled turbojet engine with hydrogen fuel in order to make engine start at high altitudes sure. The combustion wind tunnel is a blow-down type. A fuel injector of the sub-scale pre-cooled turbojet engine was installed into the combustion wind tunnel. Conditions in which a flame can be stabilized at the fuel injector were examined. The combustor pressure and equivalence ratio were varied from 10 to 40 kPa and from 0.4 to 0.8, respectively. The mean inlet air velocity was varied from 2 to 48 m/s. Flames stabilized at 20 kPa in pressure and 0.6 in equivalence ratio were observed. It was found that the decrease in the combustor pressure narrows the mean inlet air velocity range for successful flame holdings. Flame holding at lower combustor pressures is realized at the equivalence ratio of 0.4 in the low mean inlet air velocity range, and at the equivalence ratio of 0.6 in the high mean inlet air velocity range. Flame luminosity is the largest near the fuel injector. The flame luminosity distribution becomes flatter as the increase in the mean inlet air velocity.

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Based on a combustion characteristic that some propellants cannot burn in an intermediate pressure range while they can burn at lower and higher pressure, an active thrust modulation system was developed. The motor changed its thrust in dual combustion mode. It chose alternately high and low thrusts during combustion. The transitions from the low mode to the high mode were attained by a secondary ignition system, and those from the high mode to the low mode were done by a brief gas release system. Up to three times mode transitions were successfully demonstrated. Required time for the mode transitions and the effects of the thrust modulations on the thrust performance were evaluated. On the present motor scale, the transition time from the low mode to the high mode ranged from 0.12 s to 0.23 s, and that from the high mode to the low mode was from 0.19 s to 0.38 s. The thrust variable range was adjustable by the throat area. The ratio of the low-mode thrust to the high mode thrust was variable from 3 to 5. Specific impulse was decreased from 195 s to 175 s, when the number of thrust modulations was increased.

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To achieve reliable transmission of detonation wave to a pulse detonation engine (PDE) combustor, authors have proposed a PDE initiator, which consists of a predetonator and a reflector. A detonation wave propagates around the reflector changing its shape through three transition processes; from planer to cylindrical, toroidal, and planar again. Our previous study revealed that the transition to the cylindrical detonation wave upstream of the board plays a significant role in detonating hydrogen-air mixture in a 100-mm-diam-combustor. A self-sustainable condition of the cylindrical detonation wave is severe when the radius of the wave front is small. In cases using hydrogen-oxygen mixture as driver gas for the 100-mm-diam-combustor, we had to fulfill with driver gas entire upstream of the board at the critical condition for the transition to the cylindrical wave. On the other hand, curvature of the cylindrical detonation wave front becomes smaller with increasing radius of the front, so the self-sustainable condition of the cylindrical wave must be mitigated for a large bore combustor. In this study, we investigated the necessary filling diameter of the driver gas to detonate hydrogen-air cylindrical detonation by using a 500-mm-diam-cylindrical-combustor.

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This paper deals with a theoretical analysis of the low-frequency combustion instability induced by the combustion time lag of liquid oxidizer in small-scale hybrid rocket motors. We obtained the determined linear stability limit using the following parameters: the combustion time delay of liquid oxidizer, the residence time of a combustion chamber, injector pressure, chamber pressure, mass flux exponent, O/F, and the polytropic exponent of mixture gas in a combustion chamber. Kitagawa and Yuasa sometimes observed low-frequency oscillations, such as chugging, in their swirling-oxidizer-flow-type hybrid rocket engine. The obtained theoretical stability limit was compared with these experimental data.

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A pressure-fed engine with a regeneratively-cooled combustion chamber is studied in JAXA. Operation chamber pressure is approximately 1 MPa. A proposed propellant combination is liquid oxygen and ethanol. However, it is necessary to understand the critical heat flux when ethanol is used as a coolant for regeneratively-cooled combustion chamber because the saturation pressure of it is 6.3 MPa. In general, it is known that the cooling wall with fins improves the cooling performance. In this study, the effect of triangular fins on critical heat flux of ethanol in ethanol-cooled combustion chamber was investigated. As the result, it was found that the critical heat flux of cooling wall with triangular fins was 23 % higher than that of that without fin in the same velocity condition of the coolant. The critical heat flux increases by the triangular fins on the cooling surface due to the effect of the combination cooling with film boiling and nucleate boiling.

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The dual bell nozzle offers a simple and efficient altitude adaption through its contour inflection, which insures symmetrical and controlled separation at sea level and a large area ratio at high altitude. The understanding of the transition from one operating mode to the other is the key to its prediction. Intensive cold flow investigation has led to an empirical criterion for a precise transition prediction. However, the testing conditions may have a great influence on the actual transition. The temperature of the flow and the nozzle wall, especially when driving test series, can shift the nozzle pressure ratio leading to the transition. A test campaign in a high altitude chamber has also shown the dependence of the transition condition with the chamber pressure, i.e. with the density of the separated backflow. For a reliable prediction of the dual-bell nozzle flow behavior, both on test facility and in real flight conditions, all these influences have to be identified and taken into account.

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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 7–11 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. Direct-connect combustor tests were performed to evaluate effectiveness of a combustion enhancement technique termed auxiliary injection, i.e., a portion of fuel to be directly injected into airflow to provide ignition source for the residual fuel. Results of both the engine model tests at HIEST and the direct-connect tests are summarized and presented, and modification to the engine model for combustion enhancement was proposed.

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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 space launch vehicles. Either SSTO (Single Stage To Orbit) system or TSTO (Two Stage To Orbit) system with separation at high altitude needs final stage acceleration in space, so that the RBCC (Rocket Based Combined Cycle) engine should be operated as rocket engines. Performance of the scramjet combustor as the extension to the rocket nozzle, was experimentally evaluated by injecting inert gas at various pressure through the embedded rocket chamber while the whole sub-scaled model was placed in a low pressure chamber connected to an air-driven ejector system. The results showed that the thrust coefficient was about 1.2, the low value being found to mainly due to the friction force on the scramjet combustor wall, while blocking the scramjet flow pass’s opening to increase nozzle extension thrust surface, was found to have little effects on the thrust performance. The combustor was shortened to reduce the friction loss, however, degree of reduction was limited as friction decreased rapidly with distance from the onset of the scramjet combustor.

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The final purpose of this study is to develop a design tool for hybrid rocket engines. This tool is a computer code which will be used in order to investigate rocket performance characteristics and unsteady phenomena lasting through the burning time, such as fuel regression or combustion oscillation. When phenomena inside a combustion chamber, namely boundary layer combustion, are described, it is difficult to use rigorous models for this target. It is because calculation cost may be too expensive. Therefore simple models are required for this calculation. In this study, quasi-one-dimensional compressible Euler equations for flowfields inside a chamber and the equation for thermal conduction inside a solid fuel are numerically solved. The energy balance equation at the solid fuel surface is solved to estimate fuel regression rate. Heat feedback model is Karabeyoglu's model dependent on total mass flux. Combustion model is global single step reaction model for 4 chemical species or chemical equilibrium model for 9 chemical species. As a first step, steady-state solutions are reported.

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In an extensive series of studies on the initiation process of ammonium perchlorate (AP) combustion employing high-level quantum chemical techniques, we have investigated 85 reactions occurring in the gas phase, in solution and on AP crystal surface. Gas-phase reactions including the unimolecular decomposition of HOClO₃ and numerous bimolecular processes such as OH + ClO₃, NH_x + ClO_y (x = 2, 3; y = 0 - 4), HOCl + H/O/HO/HO₂ , HOCl + ClO_x (x = 1- 4) and HOCl + HNO_x (x = 1, 3), Cl/ClO/OH + HClO₃, ClO + NO_x/HO_x (x = 1 - 2), among others, have been studied by molecular orbital and quantum-statistical theory calculations. Several examples of these processes will be discussed. For the reactions occurring on the AP surface and in water solution, we have employed the periodic boundary and continuum solvent models (PCM) to study the effects of water on the sublimation and decomposition of AP. The results of these calculations agree quantitatively with available experimental kinetic data. The activation energies for redox processes in water solution were computed to be very high and no reactions leading to the production of NO_x, ClO₂ and HOCl species as proposed in the literature could be identified.

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The next generation solid propellant must fulfill the demand of low cost, and one of the prospective options is to use low melting temperature thermoplastic as a fuel binder. Continuous and small scale processing is possible for low melting temperature thermoplastic solid propellants (LTP) because of its easy manufacture process, which enables the elimination of large scale manufacturing facilities and the storage of propellants in small pieces. The latter enhances the flexibility of the manufacture schedule, and both of them will bring low cost of solid propellant, and LTP can be a part of the solution for the next generation solid launchers. Thermoplastics developed by KATAZEN have enough elastomeric properties and were evaluated as a fuel binder. Linear burning rate characteristics, mechanical properties of these LTPs are presented in this paper and discussions are made for the real applications.

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Ammonium perchrolate (AP) is the most useful oxidizer for the solid rocket motors (SRMs) under the present situation, however it is also the source of the environmental pollution close to the launch site. It is well known that HCl is exhausted through the firing of SRMs and its concentration reaches approximately 10 to 20 mole percent of the total exhaust. ‘Environmentally friendly’ and ‘High performance’ are required for the next-generation SRMs. In this study, ammonium dinitramide (ADN), which has recently attracted attentions as a halogen-free oxidizer was employed for a substitution of AP and the combustion properties of the composite propellants were investigated. Thermoplastic elastomer (TP) and hydroxyl-terminated butadiene polymer (HTPB) were used as a binder for this research. Pyrotechnic sensitivity analysis was conducted to estimate the safety of TP/ADN propellants. Strand burning tests were also carried out for all samples and the burning rate and temperature profile were obtained through these experiments. It was found that the burning rate characteristics of ADN-based propellants were influenced by the binders.

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This paper describes the influence of cathode configuration on performance of an arcjet thruster using dimethyl ether (DME) propellant. DME, an ether compound, has suitable characteristics for a space propulsion system; DME is storable in a liquid state without being kept under a high pressure, and requires no sophisticated temperature management such as a cryogenic device. DME can be gasified and liquefied simply by adjusting temperature whereas hydrazine, a conventional propellant, requires an iridium-based particulate catalyst for its gasification. In this study, thrust of a 1-kW class DME arcjet thruster is measured at a discharge current of 13 A, DME mass flow rates ranging 15 to 60 mg/s under three cathode configurations: flat-tip rods of 2 and 4 mm in diam. and 4-mm-diam. rod having a cavity of 2 mm in diameter. Thrust measurements show that thrust is increased with propellant mass flow rate. Among the tested cathodes, the flat-tip rod of 4 mm in diam. with 55 mg/s DME flow rate yielded the highest performance: specific impulse of 330 s, thrust of 0.18 N, discharge power of 1400 W and specific power of 25 MJ/kg.

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This paper reports operation characteristics of a steady-state, two-dimensional magnetoplasmadynamics (MPD) thruster using a hollow cathode. In order to improve cathode performance, a tungsten rod was replaced with a hollow cathode which had better electron emission performance and a longer lifetime. A rectangular acceleration section was employed between electrodes. The discharge voltage and thrust were measured with varying an applied magnetic field, discharge current and propellant mass flow rate.

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Lifetime evaluations by the numerical analysis of ion engines has become important for long-term missions. To use numerical analysis as a tool for evaluating the engine lifetime, a code that predicts grid erosion caused by the impact of ions quantitatively is required. In most codes, the electron number density is estimated from the Boltzmann relation because the calculation time of three-dimensional analyses becomes huge. In this paper, to evaluate the applicability of the Boltzmann relation in ion engine analysis, three-dimensional analyses of a Hayabusa type three-grid ion engine using a full-PIC (particle in cell) code and hybrid-PIC code using the Boltzmann relation were performed. The comparison of the results revealed that the electron number densities obtained by the two analyses agree well in almost the entire region. However, in the downstream neutralizing region, where the electric potential is positive, the electron number density for the hybrid-PIC code increases up to about seven times higher than that for the full-PIC code.

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Microgravity experiments were performed to evaluate liquid propellant retention force of a porous metal. In our gas-liquid equilibrium propulsion system, porous metals are equipped in a storage tank and surface tension in pores of the porous metals holds liquid propellant in them, which ensures expelling of only gaseous propellant from a storage tank. The performance of a porous metal for liquid retention was evaluated by two different microgravity experiments. In the first one, the acrylic tank filled with semilunar porous metals and liquid propellant was rotated by a motor. The performance was evaluated by considering a force balance between liquid retention force and centrifugal force acting on the liquid propellant absorbed in porous metals. In the other one, the internal pressure in the tank was reduced by propellant ejection from a nozzle. The performance was evaluated by exhaustion time. As a result, it was found that liquid retention force was equal to, or higher than analytical value (bubble point pressure), and extractable gas volume of each ejection strongly depends on the existence of bubbles in liquid propellant.

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Channel wall erosion of a 500 W class SPT-type Hall thruster was computed by using a 2D3V Fully Kinetic Particle-In-Cell Direct Simulation Monte Carlo model. Electrons are treated as particles and their trajectories are directly calculated in order to achieve self-consistent simulation. Compared with previous studies, the calculation acceleration technique of artificial electric permittivity was not implemented to avoid change of physics. In addition, the BN channel wall was included into the calculation domain to improve the wall sheath calculation for accurate wall erosion prediction. The calculation results of both case “without BN wall” and case “with BN wall” are presented and compared with the experimental data. Better agreement was achieved by the new case “with BN wall” rather than the conventional case “without BN wall”, which indicates the importance of calculating the electric field within the insulator wall.

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Ionization oscillation of a Hall thruster is one of the most serious problems. In order to suppress the ionization oscillation of an anode-layer-type Hall thruster, we tried to reduce the magnetic field in the anode hollow using a magnetic shield. Firstly, we validated the effectiveness of a magnetic shield with numerical simulation which was developed for the UT anode-layer-type Hall thruster (power: 1.7 kW, thrust: 93 mN, I_sp: 1900 s). Secondly, we developed a new anode-layer-type Hall thruster in which anode hollow magnetic flux density is reduced using a magnetic shield and measured its oscillation amplitude and thrust efficiency. As a result, the oscillation amplitude was drastically reduced by the magnetic shield; the oscillation amplitude was less than 0.4 in the whole operation range. Moreover, the thrust efficiency is also improved by more than 10 percent by the magnetic shield.

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The inductively coupled plasma cathode (ICP/C) has been developed as electron source for ion thrusters to liberate the thrusters from the limitations of hollow cathode, such as the lifetime limit. Therefore, in order to improve the power consumption efficiency of the ICP/C, its ignition and electron emission characteristics were investigated experimentally as functions of orifice dimensions and ion collector shape in this study. It is obtained that the factors affecting its plasma ignition are not only the applied RF voltage and the vessel inner pressure but also the conditions in the vessel. Its ignition capability is enhanced with increasing the orifice length and decreasing the orifice diameter because the inner pressure increases at constant xenon mass flow rate. On the other hand, there exists an optimum diameter of orifice for electron emission performance of the ICP/C. Additionally the pole collector is preferable to the cylindrical collector for electron emission. When the pole collector was inserted in the cathode, the typical performance was 0.48 A of the anode current at 24 W of RF power, 0.10 mg/s of xenon mass flow rate and 40 V of anode voltage.

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Solar wind plasma flow with interplanetary magnetic field (IMF) and the thrust of the magnetic sail are examined by time-dependent, two-dimensional, X-Y Cartesian, hybrid particle-in-cell (PIC) simulations. The hybrid-PIC simulation model is that the ions are treated kinetically as particles and the electrons are modeled as an inertia-less (mass-less) fluid. In this simulation, the real solar wind parameters around a near-earth orbit are used. The direction and strength of IMF are set to +Y direction which is perpendicular to the solar wind flow (+X direction) and 3 nT. Expressed in r_L/L (the ratio of an ion Larmor radius r_L of the solar wind at the magnetopause to a representative length of magnetic field L), when r_L/L = 0.1 (in the case of MHD scale), magnetopause is formed accompanied by a fast magnetosonic bow shock. When r_L/L = 2.0 (in the case of ion inertial scale), the electromagnetic interaction results in the formation of a magnetosphere with standing whistler waves. The drag coefficients, which is the thrust normalized by the solar wind inertial force, of both scales with IMF tend to increase compared with the cases without IMF because the incoming IMF accompanied by the solar wind piles up at upstream of the spacecraft. Also, on the ion inertial scale, the generation mechanism of Whistler wave and the influence of that on the thrust performance are revealed.

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Magnetic sail is a propellantless propulsion system used in space, which is capable of generating a propulsive force by the interaction between the magnetic field generated by a hoop coil and the solar wind plasma flow. Three dimensional hybrid (ion particles and electron fluid) particle-in-cell (PIC) simulation and scale-model experiment are performed to investigate the characteristics of the plasma flow around a magnetosphere on the ion inertial scale where an ion gyro radius rLi is comparable to the representative size of magnetosphere L. It is found that the dark region around magnetospheric boundary appearing in the experimental photograph corresponds to the region where the plasma density increases due to the plasma trapped by the magnetic field. The induced current which is both perpendicular to the plasma flow and the magnetic field also increases in the magnetospheric boundary region hence this region coincides with the magnetopause current layer. The width of the magnetopause current layer has a good agreement between the numerical simulation result and experimental result. Also, the predicted thrust value of 0.34 ± 0.01 N obtained by the hybrid simulation agrees well with the experimental result when numerical simulation is carried out by considering the ion-neutral collision effect. The hybrid PIC simulation carried out without considering the collisional effect gave a thrust value of 0.4 ± 0.01 N (increasing by a factor of 1.3), which can be applied to the thrust evaluation of the magnetic sail in a collisionless interplanetary space.

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Laser-supported Detonation (LSD), which is one type of Laser-supported Plasma (LSP), is an important phenomenon because it can generate high pressures and temperatures for laser absorption. In this study, using thermal-non-equilibrium model, we numerically simulate LSPs, which are categorized as either LSDs or laser-supported combustion-waves (LSCs). For the analysis model, a two-temperature (heavy particle and electron-temperature) model has been used because the electronic mode excites first in laser absorption and a thermal non-equilibrium state easily arises. In the numerical analysis of the LSDs, laser absorption models are particularly important. Therefore, a multi-charged ionization model is considered to evaluate precisely the propagation and the structure transition of the LSD waves in the proximity of the LSC-LSD threshold. In the new model, the transition of the LSD construction near the threshold, which is indicated by the ionization delay length, becomes more practical.

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The pulsed plasma thruster (PPT), has attracted attention again as a micro-thruster because of its compactness, light weight, and comparatively low power consumption. On the other hand, the propellant utilization efficiency of a conventinal Teflon PPT is relatively low among electric propulsion devices because a propellant that originates from late-time ablation produces negligible thrust. The liquid propellant PPT (LP-PPT), in which water or ethanol is fed with an injector, was proposed to overcome these difficulties. Thrust measurements show that a LP-PPT provides higher specific impulses than a conventional PPT. However, water requires temperature management for propellant storage due to its relatively high freezing point. Moreover, even if ethanol, which has a sufficiently low freezing point, is used as propellant, a pressurant is necessary, as well as water, because the vapor pressures are insufficient for self-pressurization. In this study, we propose to use dimethyl ether (DME) as the propellant. DME, which has a freezing point of 131 K at 1 atm and a vapor pressure of 6 atm at 298 K, can be stored in tanks as a liquid, and requires no feeding pressurant. We designed a DME pulsed plasma thruster to evaluate performance. Thrust measurement yielded a specific impulse of 430 s for a coaxial type at a capacitor-stored energy of 13 J.

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The plasma behavior in a magnetic thrust chamber system for a laser fusion rocket is numerically simulated using a three-dimensional (3D) hybrid particle-in-cell (PIC) code and a one-dimensional (1D) radiation hydrodynamic code. The magnetic thrust chamber and an applied magnetic field with a suitable geometry generate an impulse from the interaction between the diamagnetic current in the laser-produced plasma and the magnetic field generated by a magnetic coil. A 1D radiation hydrodynamics code is used to compute the hydrodynamic evolution of a radiating plasma heated by laser beams or external radiation sources. By combining this code and a 3D hybrid PIC code, a series of numerical simulations are performed to investigate high-energy laser injection onto a fuel target and the ablated plasma behavior of the system. A thrust energy of 0.37 J and an impulse bit of 31.6 μNs are obtained for an incident laser energy of 4.0 J. This impulse bit could mostly be generated by interactions between a slowly expanding plasma (expansion velocity of ~20 km/s) and a magnetic field. To optimize this system, it is important to reduce the expansion velocity of the laser-produced plasma.

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In order to validate the grid erosion evaluation code for the lifetime validation of ion thrusters, the electron number density/temperature in the vicinity of a screen gird in a 30 W class microwave discharge ion thruster were measured by means of laser Thomson scattering (LTS) technique. A photon counting method and a triple grating spectrometer were used against a small Thomson scattering signal and a strong stray laser light. Observed Thomson scattering spectrum tells that the electron energy distribution function was Maxwellian. From this spectrum and the Rayleigh scattering calibration using nitrogen gas, electron number density and electron temperature were calculated to be (3.8±0.2)×10¹⁷m^-3 and 6.2±0.1 eV, respectively at incident microwave power of 8 W and krypton mass flow rate of 6.2 μg/s. The ion saturation currents estimated from the LTS measurement are in good agreement with ion beam currents through the screen grid for several conditions. These results show that LTS technique is a useful non-intrusive tool for measuring plasma property in the vicinity of the screen grid.

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A series of successful ion-propelled missions opened a new era of space exploration and development using high-Isp ion engines. Modifications and improvements of the current ion engines and the development of new ion engines are underway to enhance mission performance using ion engines. However, the life qualification tests for ion engines are difficult, time-consuming, and costly processes, which is a serious problem for the future use of ion engines. Therefore, the numerical evaluation of ion engine life has become very important. In 2006, JAXA started the JAXA Ion Engine Development Initiative (JIEDI) project to develop a numerical tool for the life qualification of ion engine optics. The developed code (JIEDI-1) can analyze the wear of the grid system caused by ion and neutral sputtering with a reasonable computational time. In this paper, the JIEDI-1 code is explained in detail and several examples of its use are presented.

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Grid erosion analysis of a μ20 ion engine was carried out. The design principle of the grid optics for the μ20 ion engine is fairly unique, in that the acceleration grid is ion-beam machined to produce an optimum grid shape by the direct impingement of ion beams. By adopting a new grid surface model and strategy for updating the grid surfaces, the highly eroded grid surfaces of the μ20 ion engine were successfully analyzed. Parametric studies were also conducted to assess the grid lifetimes for different acceleration grid voltages.

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In order to estimate the erosion of accelerator grids of an ion engine, the low-energy sputtering of carbon material under Xe ion bombardment and the redeposition of low-energy sputtered particles on the carbon material are studied through the molecular dynamics (MD) simulation. For the normal incidence of Xe, the MD result of sputtering yield almost agrees with the experimental result by Williams et al. (AIAA-2004-3788). However, the experimental result shows less incident angle dependence than the MD result because it performed on a rough surface. For the redeposition on amorphous carbon (a-C) surface, the reflection and the deep trapping are suppressed. The mean deposition rate with deposition energy up to 100 eV is 0.82∼0.75 taking account of incident angle distributions.

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A 20-cm diameter electron cyclotron resonance xenon ion thruster for space propulsion is under development that generates 500 mA of ion beam current at a microwave discharge power of 100 W. It does not have any moving mechanical parts for microwave impedance matching. Extracted ion currents and reflected microwave powers were experimentally investigated around a nominal frequency of 4.25 GHz for different flow rates. Optimized frequency tuning within 0.6% of the nominal frequency minimized the microwave reflection and maximized the ion current at each flow rate between 0.39 and 1.27 mg/s. However, constant frequency operation at 4.266 GHz is recognized as the best strategy because it provided with fare performance in wide range of flow rate and almost minimum reflection during tentative stop of beam extraction after high voltage breakdowns.

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In order to validate the concept of a magnetic thrust chamber for laser fusion rockets, the interaction between laser-produced plasma and the magnetic field produced by a permanent magnet/coil is investigated. Time variation in magnetic flux density in the magnetic thrust chamber was observed by means of several loop probes. The magnetic flux density at the midpoint between the permanent magnet and the target was found to decrease, approaching zero at 50 ns after laser ignition; it then recovers to its initial value after 1 μs. This result demonstrates the concept of the magnetic thrust chamber; the diamagnetic cavity was observed. The impulse bit of the magnetic thrust chamber is estimated from the time variation of the magnetic flux density; it is 0.6 μNs at a laser energy of 0.6 J and laser pulse duration of 2 ns. The estimated impulse bit is in good agreement (order of magnitude) with that measured using a pendulum thrust stand.

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On-orbit identification of wrinkle states in membrane structures is a key technology for the operation of space structures with high performance and high precision. This study investigates the applicability of elastic waves for the identification of wrinkle states in membranes in both theoretical and experimental aspects. Cross-sectional changes generated by the wrinkles are focused on, and the induced dispersion characteristics of elastic wave are used for the identification of wrinkle states. The identification methodology is theoretically derived, and experimental demonstration is also presented.

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The purpose of this study is to investigate the effect of attachment errors of two beam-type flexible appendages mounted on a rigid body along the spin axis. For such a system, the sufficient conditions for achieving asymptotical dynamic stability are maximization of the moment of inertia about the spin axis and existence of a lower limit of the beam's natural frequency against the spin rate; these conditions can be acquired by Lyapunov's direct method. However, such conditions are limited to ideal configurations; for realistic satellites, attachment errors are unavoidable. In this study, each appendage that simulates a continuously flexible beam is modeled as a mass particle connected to the rigid body's space through springs. This paper illustrates mathematical formulations for acquiring the equilibrium state and determining the dynamic stability using Lyapunov's direct method. Numerical examples show that attachment errors either improve or worsen the stability in terms of the sufficient conditions.

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We propose a digital autonomous power scavenger with a microprocessor. The proposed system is a completely self-powered one that does not require any external power supply at all, and can thus be used portably at any site. Nevertheless, the digital approach enables the power scavenger to be programmable and thus, it affords some versatility with regard to control schemes. The proposed digitalautonomous system is much more advanced and progressive than clumsy analog-autonomous ones. It can be implemented in multiple-input multiple-output systems to scavenge electrical power from even complicated structural vibrations. We determined the value of the storage capacitance that gives the best balance between scavenging power and consumed power.

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A lobed-pumpkin balloon, currently being developed in ISAS/JAXA as well as in NASA, is a promising vehicle for long duration scientific observations in the stratosphere. Recent ground and flight experiments, however, have revealed that the balloon has deployment instabilities under certain conditions. In order to overcome the instability problems, a next generation SPB called 'tawara' type balloon has been proposed, in which an additional cylindrical part is appended to the standard lobed-pumpkin balloon. The present study investigates the deployment stability of tawara type SPB in comparison to that of standard lobed-pumpkin SPB through eigenvalue analysis on the basis of finite element methods. Our numerical results show that tawara type SPB enjoys excellent deployment performance over the standard lobed-pumpkin SPBs.

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A loop heat pipe (LHP) is a two-phase heat transfer device that utilizes the evaporation and condensation of a working fluid to transfer heat, and the capillary forces developed in fine porous wicks to circulate the fluid. LHPs have been gaining increased acceptance for spacecraft missions, and recently, small LHPs on the order of a few hundred watts have been investigated for this purpose. In this study, a 100W class small LHP with a polytetrafluoroethylene wick as the primary wick was designed and fabricated for thermal vacuum testing. The LHP has a thermoelectric converter (TEC) to control the loop operating temperature. The thermal vacuum test was conducted to evaluate the LHP's thermal performance under a space-simulated environment such as ultra-high-vacuum, and black body radiation, except for a gravitational effect. The loop showed large thermal hysteresis before and after the large and small head loads. The TEC was able to control the loop operating temperature with a small amount of electrical power.

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For the worst-case analysis of a space structure's shape on orbit, various factors such as effects of thermal deformation, aged deterioration, and material hysteresis should be considered. Furthermore, the parameters of each factor have some uncertainty, such as in material properties. Therefore, in shape prediction, the consideration of these various factors and their parameters' uncertainties leads to a combinatorial explosion. To solve this problem, the factors are classified by the mode shape of deformation. If the mode shapes of some factors have high correlation, those factors are categorized in the same group. Within each group, maximum and minimum deformations are analyzed considering the uncertainty in the parameters. Among the groups with low correlation, deformations are evaluated using a combination of the maximum and minimum deformations from each group. As a result, the combinations of factors and parameters are drastically reduced. Such a shape prediction method was applied to a large deployable antenna structure of ASTRO-G. In this study, the performance of this antenna is evaluated using GRASP analysis for the predicted antenna shapes.

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In the last decade space agencies and industries have focused their attention on proposing and developing a new class of large space structures (LSS). LSS have potentially broad applications in space including low-stiffness precision shaped antennas for mobile communications satellites, narrow-band broadcast services and remote sensing, low-stiffness planar structures for large solar arrays. A lightweight design responding to the low gravity environment and live loads due to gravity gradients, extreme thermal variations, and solar wind create new and challenging problems in dynamic analysis of LSS. Since LSS will be designed for minimum weight and be quite flexible a significant problem will also arise on developing the capability to suppress and control flexible modes of vibration, some of which are closely packed with their frequencies that in turn could be very close to the ones of the attitude controller. On account of this it is necessary to have a very accurate model of the LSS. In the present work a hybrid FEM-continuous formulation, based on the classical modal reduction of a LSS under gravity and gravity gradient forces, will be presented. Analyses will be also performed to evaluate the effects of modal truncation on the robustness of attitude control laws.

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In this paper, the response of closed-cell aluminum foam sandwich structures under low velocity impact of an ogival-ended mild steel projectile was studied. An analytical model was proposed to predict the penetration depth as well as residual velocity based on the dynamic cavity expansion theory and the Poncelet resistance formula using in the closed-cell aluminum foam sandwich panel. A high-pressure air gun was employed to execute penetration depth tests. Two tests had been done and the results were measured, those are, the residual velocity of the first layer (aluminum plate) and the penetration depth of the sandwich panel. Simultaneously, the finite element analysis software LS-DYNA was utilized for describing the penetrating process of steel projectiles normal impact. It was found that both residual velocity and penetration depth showed nonlinear variation against impact velocity, also the analytical results showed acceptable deviation with those from tests and FEM simulations during specific impact velocities.

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In this study, the shape repeatability of cable-network structures is examined. Cables used for cable networks have nonlinearities between the tensile load and elongation. Therefore, the shapes of cable-network structures have some uncertainty, which leads to diminished shape repeatability. In order to clarify the influence of such nonlinearities, numerical simulations and experiments are carried out. The mechanical properties of the cables are experimentally investigated and their nonlinear characteristics are numerically modeled. The experimental model examined in our previous study is employed as the analysis model. The shape repeatability of the models caused by the nonlinear characteristics of cables and rigid body rotations of nodes are investigated through numerical simulations. The modeled nonlinear characteristics of the cables used in the numerical simulations are validated by comparing with experimental results.

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In this paper, the spring-mass system model developed for simple numerical simulations of thin membranes is enhanced by taking into account the properties of buckling and creases. The model is applied to the numerical simulations of centrifugal deployments of folded square membranes that are small-scale models for solar sail spacecraft “IKAROS”. First the folding and deployment methods are reviewed. Then the formulation of the enhanced spring-mass system model is explained. Numerical simulations of the centrifugal deployments of two kinds of folded square membranes with different crease intervals are performed and the numerical results are compared with the corresponding experimental results. The deployment behaviors are discussed and the validity of the spring-mass system model is examined.

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A magnetized plasma sphere generated by a rotating magnetic field (RMF) technique has been proposed as an effective shield for innovative space transportation systems, such as a space elevator, from energetic particle radiation found in the Van Allen radiation belts, for example. In the proposed system, the RMF field drives electron ring current inside and outside of the antenna. Then, the generated ring current forms a magnetic sphere which acts as a plasma shield. In this work, generation of electron ring current by RMF has been evaluated numerically. The shielding effect of the generated magnetized plasma sphere has been demonstrated numerically with a simple ring current model. The results showed a combination of azimuthal drift motion of and radial Lorenz force on energetic charged particles. Also, an experimental study has been initiated. The developed IGBT invertor system and a RMF antenna set are also briefly introduced. As an initial experimental work, plasma formation has been attempted in a vacuum tank by using scale models of RMF antenna sets.

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