The main objective of this research is automatic generation of more accurate DEM/ortho-image and, by extension, to update existing DEMs and ortho-images database automatically by using high resolution satellite images. In the first part, we tested the possibility of automatic GCP and tie-point generation for DEMs and ortho-images generation. We extract Ground Control Points (GCPs) automatically from high resolution satellite image by using existing ortho-images and existing DEMs. In second part, we tested whether high resolution images can be used to update existing DEMs and ortho-images. For this, we performed automatic generation of DEMs/ortho-images by using true GCPs obtained by GPS surveying. For testing the method proposed, we used IKONOS, SPOT 5, QuickBird, Kompsat-2 high resolution satellite images. Assessment for results supports that automatic generation of DEMs/ortho-images was performed successfully. We generated a DEM from IKONOS images fully automatically by using GCPs generated from existing DEMs and ortho-images. Newly updated DEM was compared with existing data. Updated DEM shows more detailed surface, and the result of validation shows that the accuracy is similar with that of existing DEMs/ortho-images.

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The TOKYO model is a laboratory model and the TSUKUBA model is an air-borne model of GOSAT TANSO-FTS. The TOKYO model is devolved to demonstrate the application of the Fourier transform spectrometer to the satellite remote sensing of short wave spectral region and as well as the feasibility of the CO₂ and CH₄ column-density on the ground. The TSUKUBA model is manufactured for the air-borne experiment. This paper presents the objectives, design, and performance test results of these models.

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Tetra-ol glycidyl azide polymer (GAP) is one of the best candidates for the solid fuel of gas hybrid rocket system because of self-combustibility and high heat of formation. Combustion model of GAP was developed by Beckstead et al. and they applied it to tri-ol GAP successfully. We have applied this model to tetra-ol GAP as an initial attempt, and numerical simulation showed that maximum temperatures in the gas phase exceeded those of experimental results significantly, and calculated burning rates were much higher than strand burner data, thus, modification of the model taking account of combustion incompleteness was found to be necessary. Modifications of combustion model were made taking the residue analysis results into account as Blow Off Mechanism. Simulated final temperature in the gas phase and burning rate are lowered effectively and coincide well with experimental data adjusting kinetic parameters.

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We observed Al droplets near propellant surface with the CMOS camera, achieved the observation of Al droplets near propellant surface through 1MPa to 8MPa and measured the diameter of Al droplets near the burning propellant surface. We also observed the propellant section cut by the microtome and measure the size of Al droplets in the pockets. We investigated the relationship between the diameter of Al droplets near the burning propellant surface and the pocket sizes. When the direction of acceleration is the same to the burning direction, the acceleration affect the diameter of Al droplets, and the size of Al droplets become larger than those under static conditions.

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An innovative Panel ExTension SATellite (PETSAT) and propulsion system for PETSAT, are presented in this paper. Based on PETSAT ethos, design policy of the propulsion system is provided. According to the policy, we designed propulsion system and concretely estimated and assembled mono-propellant and bi-propellant systems, and it indicated that mono-propellant propulsion with 50-60 seconds of specific impulse and 1 N of thrust is probable. In the case of bi-propellant, 120-150 seconds of specific impulse is valid even based on the design policy. We conducted captive tests of mono-propellant and bi-propellant propulsions with a breadboard model and flight model of propulsion system for PETSAT, and obtained good operations and performances. Now we are proceeding with its further improvements.

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A rocket-ramjet combined-cycle engine was tested in ramjet-mode. The combustor model had two rockets in the combustor section. They were used as an igniter in this operation mode. In the preliminary tests, the downstream combustion ramjet-mode was demonstrated with a 1.4-degree of divergent duct condition. In this study, the upstream and downstream combustion ramjet-mode operations were applied to the combined cycle engine model with large angle of divergent duct condition. In the case of upstream combustion ramjet-mode, the combustion condition at the exit of the combustor showed high combustion efficiency.

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Pre-cooled turbojet engines have been investigated aiming at realization of reusable space transportation systems and hypersonic airplanes. Evaluation methods of these engine performances have been established based on ground tests. There are some plans on the demonstration of hypersonic propulsion systems. JAXA focused on hypersonic propulsion systems as a key technology of hypersonic transport airplane. Demonstrations of Mach 5 class hypersonic technologies are stated as a development target at 2025 in the long term vision. In this study, systems analyses of hypersonic turbojet experiment (HYTEX) with Mach 5 flight capability is performed. Aerodynamic coefficients are obtained by CFD analyses and wind tunnel tests. Small Pre-cooled turbojet is fabricated and tested using liquid hydrogen as fuel. As a result, characteristics of the baseline vehicle shape is clarified, . and effects of pre-cooling are confirmed at the firing test.

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In the light of recently revived interest in high energy density fuels for aerospace applications^1,2), a new look is being given at unconventional fuels. Among the latter are hydrides, because their hydrogen content and density. Among hydrides silanes are of interest because of their combustion and energetic properties.
Silanes are silicon hydrides organized in molecular chains similar to those of hydrocarbons; at STP, lower silanes (SiH₄, Si₂H₆) are gaseous and extremely pyrophoric; with increasing chain length, silanes become liquid from trisilane (Si₃H₈) on, and therefore easily pumped. Another important feature of silanes is the large amount of hydrogen theoretically available by thermal decomposition: in fact at moderate temperatures (about 500 K) the chains begin to break and at 700 K their decomposition is complete, yielding silicon and gaseous hydrogen, useful for propulsion in combination with air nitrogen and oxygen. This last feature, if confirmed, could identify silanes not only as energy carriers but also components in bi-fuel systems. To assess their theoretical performance, simulations were conducted assuming silanes and/or their thermal decomposition products in combination with various oxidizers and air. Preliminary results are suggestive of their potential for some specialized applications, especially where compactness is at premium.

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The regenerative cooled combustion chamber of a cryogenic liquid rocket engine is exposed to a large temperature difference between the hot gas (about 3500K) and the liquid hydrogen (about 20K). This induces thermal stress, and strain is accumulated in the chamber wall throughout the cyclic firing tests. Evaluation of the stress and the strain distribution in a chamber wall is essential since chamber life is usually related to such stress and strain. In this study, the residual strain in a regenerative cooled combustion chamber wall was measured by applying the neutron diffraction method and the X-ray diffraction method. The measured data were compared with the numerical data by finite element analysis, and the feasibility of the neutron diffraction method for the regenerative cooled combustion chamber of a cryogenic liquid rocket engine was evaluated.

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Quasi-one-dimensional, both equilibrium and kinetic analyses were carried out to evaluate vitiation effects on thrust performance of a typical engine under M6 flight conditions. With incoming-flow parameters varied, it was found that matching static pressure and static temperature (or total enthalpy) of the incoming flow gave minimum deviation on engine performance between the non-vitiated and vitiated cases. However, change in thermal properties due to vitiation resulted in few percent deviation even in this case. This deviation could be corrected by using equilibrium analysis results even for kinetic calculation. Chemical kinetics was found to be dominant in relatively narrow range in fuel equivalence ratio.

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The improvement of the performance of composite propellant has been required due to the development of a wide variety of rockets and missiles. Glycerin has three oxygen atoms in its molecular structure, and it is a safe and low-viscosity material. Therefore, it is expected that glycerin would be one of candidates as a plasticizer of an AP/HTPB based-propellant. The burning characteristics and the mechanical properties of AP/HTPB propellant containing glycerin were investigated in this study. The viscosity of the uncured propellant containing glycerin was lower than that without glycerin. The tensile properties were improved with the small added quantity of glycerin. The burning rate of the propellant at 70%AP increased with the addition of glycerin and the propellant containing glycerin had a plateau burning characteristics. It was found that the addition of glycerin as a plasticizer would be an effective way to improve the burning characteristics, mechanical properties, and processability of the AP/HTPB propellant.

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A small model aerodynamic tests of the combined cycle engine were carried out to evaluate its performance in subsonic and supersonic conditions. In this regime of the flow speed, the combined cycle engine operates as an ejector-jet or ramjet. The nitrogen gas was exhausted as the substitution for the actual rocket gas. In a subsonic condition, there appeared local pressure rise at the kink point of the ramp, increasing the pressure drag. Both wall pressure and the pitot pressure distribution at the exit of the model suggested that the flow structure is “two-layered” ; one is subsonic induced air flow, and the other is the supersonic rocket exhaust. A slit was carved on the topwall inside the isolator section, expecting a better suction performance in the ejector-jet mode. The modification actually had an effect to enhance the lower limit of the rocket pressure at which the choking of the induced air is achieved.

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It is generally acknowledged that bearings and axial seals have a tendency to go wrong compared with other rocket engine elements. And when those components have malfunction, missions scarcely succeed. However, fundamental performance (maximum rotational speed, minimum flow rate, power loss, durability, etc.) of those components has not been grasped yet. Purpose of this study is to grasp a critical performance of mechanical seal and hybrid ball bearing of turbopump. In this result, it was found that bearing outer race temperature and bearing coolant outlet temperature changed along saturation line of liquid hydrogen when flow rate was decreased under critical pressure. And normal operation of bearing was possible under conditions of more than 70,000 rpm of rotational speed and more than 0.2 liter/s of coolant flow rate. Though friction coefficient of seal surface increased several times of original value after testing, the seal showed a good performance same as before.

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Mixing and combustion of supersonic reacting flows are currently under investigation for new generation RBCC vehicles. Because of the speed within the SCRJ combustor, the length required for fuel and air to mix and react is commonly thought to be a difficult compromise if friction drag is to be kept to a minimum. LES simulations can be an useful tool to better understand the physics of supersonic combustion. By simulating combustion and analyzing numerical results, it is possible to draw fundamental conclusions suggesting how to design a more efficient combustor. Thus the subgrid scale (SGS) model to be implemented as closure of the filtered Navier Stokes equations is critical in reproducing experimental results and giving credibility to numerical predictions. In the present paper, a new SGS model for supersonic reacting flows is being proposed. This model accounts for compressibility effects on mixing and combustion by introducing a complete equations for the subgrid kinetic energy transport. This equation accounts for the effect of heat release due to combustion, dilatation due to the compressibility of the fluid, and the pressure gradient that are responsible for the baroclinic term and then for streamwise vorticity and dissipation. The reaction rate accounts for the effect of the Mach number by means of the SGS Mach number squared. The effect of the small scales scalars fluctuations on turbulent mass diffusivity is included by modeling the local Schmidt number.

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A series of static firing tests was conducted to investigate the fuel regression characteristics of a Cascaded Multistage Impinging-jet (CAMUI) type hybrid rocket motor. A CAMUI type hybrid rocket uses the combination of liquid oxygen and a fuel grain made of polyethylene as a propellant. The collision distance divided by the port diameter, H/D, was varied to investigate the effect of the grain geometry on the fuel regression rate. As a result, the H/D geometry has little effect on the regression rate near the stagnation point, where the heat transfer coefficient is high. On the contrary, the fuel regression rate decreases near the circumference of the forward-end face and the backward-end face of fuel blocks. Besides the experimental approaches, a method of computational fluid dynamics clarified the heat transfer distribution on the grain surface with various H/D geometries. The calculation shows the decrease of the flow velocity due to the increase of H/D on the area where the fuel regression rate decreases with the increase of H/D. To estimate the exact fuel consumption, which is necessary to design a fuel grain, real-time measurement by an ultrasonic pulse-echo method was performed.

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Within the frame of a broader activity towards research into the application of methane in cryogenic liquid rocket engines, the efficiency of film cooling was studied in a LOX/CH₄ fired subscale sized model combustor. Aiming at booster as well as upper stage applications, the combustion pressure levels have been varied between 4 MPa and 7 MPa. The effectiveness of the ambient temperature film was determined in axial and circumferential directions by measuring temperature gradients in the copper liner material. The experiments revealed remarkable circumferential differences of the film cooling efficiency which remain existent far downstream. However, circumferential film cooling varieties are more pronounced at close proximity of the point of film coolant injection.

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Laser ignition condition was studied for a new reaction control system using green propellant of GH2/GO2 by the use of laser-produced ablation plasma. Dependence of the laser pulse width on the minimum ignition laser energy was measured experimentally. Temporal behavior of the laser pulse gave an explanation of the dependence in 10 to 22 ns regime. Heating slope threshold consisting of slope threshold and duration threshold was newly introduced for the interpretation. This new concept could describe not only the explanation of the dependence but also a possibility of more efficient ignition condition.

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Efforts have been paid to realize a new composite propellant using thermoplastics as a fuel binder and lithium as a metallic fuel. Thermoplastics binder makes it possible the storage of solid propellant in small blocks and to provide propellants blocks into rocket motor case at a quantity needed just before use, which enables the production facility of solid propellant at a minimum level, thus, production cost significantly lower. Lithium has been a candidate for a metallic fuel for the ammonium perchlorate based composite propellants owing to its capability to reduce the hydrogen chloride in the exhaust gas, however, never been used because lithium is not stable at room conditions and complex reaction products between oxygen, nitrogen, and water are formed at the surface of particles and even in the core. However, lithium particles whose surface shell structure is well controlled are rather stable and can be stored in thermoplastics for a long period. Evaluation of several organic thermoplastics whose melting temperatures are easily tractable was made from the standpoint of combustion characteristics, and it is shown that thermoplastics propellants can cover wide range of burning rate spectrum. Formation of well-defined surface shell of lithium particles and its kinetics are also discussed.

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The acoustics of combustion chambers and the interaction of acoustics and combustion is investigated in a model combustor operated with LOX/H₂ and LOX/CH₄. Acoustic excitations are induced by a siren during hot fire tests and the response of atomization and combustion is recorded with dynamic pressure sensors and high speed OH-imaging is applied. By analyzing the temporal and spatial distribution of the flame response the question is addressed, whether in the experiments the coupling of acoustics to combustion is via pressure or velocity sensitive processes. In the experiments it appeared that the acoustic eigenmodes of the combustor are characteristically deviating from cylinder modes. The results obtained can be explained by a modal analyses of the combustor geometry. These investigations have been extended to study the influence of absorbers and absorber rings on the acoustics of combustion chambers and resonance frequencies predicted by modal analysis and experimental results are presented.

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Sputter monitoring system using continuous-wave cavity ring-down spectroscopy (cw-CRDS) was built for both lifetime assessment and contamination effects in Hall thrusters. First, we have performed proof of principle measurements of sputtered boron atoms from BN targets by argon ions using pulsed CRDS. The measurement strategy is based upon detection of boron atoms via an absorption line from ground state at a wavelength of 249.773 nm. The path-integrated number density is 1.8 × 10¹³ m^-2 at ion beam current of 45 mA and argon ion energy of 1,000 eV. The number density is proportional to the ion beam current, as expected. These results show the validity of the boron sensor for detecting sputtered boron atoms. Next, in order to achieve the required detection sensitivity and time response, we implement CRDS with a continuous-wave (cw) laser for enhanced sensitivity. The target was changed to manganese (λ=403.07 nm) instead of boron. The results show that a detection limit of per-pass absorbance of 0.6 ppm for a 1-s measurement time. Scaling the experimental results and accounting for changes in laser system and mirror reflectivity indicate that the BN detection system should have sufficient signal-to noise for expected Hall thruster conditions.

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A laser ignition microthruster was developed for microspacecraft propulsion. Small solid-propellant pellets were ignited by a diode laser beam to produce a high impulse from a very small power source. Basic experiments showed that the microthruster produced an impulse of 60 mNs, using 60 mg B/KNO₃ pellets ignited by a 1 W laser beam. A space-based demonstration of the microthruster system is planned onboard a microspacecraft named KKS-1, which has been chosen by the Japan Aerospace Exploration Agency as one of the six small satellites to be launched with the Greenhouse Gases Observing SATellite, which is scheduled to be launched on a H2A rocket in early 2009.

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One of the chief drawbacks of MPD thrusters is their low thrust efficiency. This paper presents a numerical optimization of the thruster nozzle and of the distribution of the self-induced magnetic field that occurs in the nozzle, with the goal of developing nozzle design guidelines. This model accounts for the pressure gradient and the Lorentz force, and optimization was performed using a genetic algorithm. The results showed that (1) large improvements in performance can be obtained in a model representing a thruster of practical size; (2) optimal location of the nozzle throat was at 15 [%] - 30 [%] of nozzle length downstream of the inlet and the throat is an essential feature of the nozzle; and (3) the profile of the upstream portion of the nozzle (where most of the acceleration of the plasma occurs) is little affected by up-scaling of nozzle size.

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Experimental and numerical investigations on thrust performance of micro-single-nozzle thrusters with different area ratios were conducted to characterize and improve thrust performance of each nozzle. In the numerical simulation, the detailed mechanism of thrust improvement was investigated. From the experimental and numerical, studies, it was shown that for micro-single-nozzles with different throat sizes with identical exit height the thrust increased when the throat size was reduced. In the simulation, it was shown that pressure along the nozzle wall and static pressure at nozzle-exit plane became larger for smaller throat size nozzles. In this case, it turned out that contribution of the pressure thrust on the total thrust was more significant.

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The current status of experimental studies of hollow cathodes for the next-generation ion engines in the Aerospace Research and Development Directorate, JAXA is described. One of the topics on the hollow cathode studies is a life test of a discharge cathode. The keeper disk, orifice plate, and cathode tube of this discharge cathode are made of "high density graphite," which possesses much higher tolerance to ion impingement compared with conventional metal materials. The life test had started in March 2006 and the cumulative operation time reached 15,600 hours in April 2008. No severe degradation has been found both in the operation voltages and electrodes so far, and the test is favorably in progress. In addition to the life test of the discharge cathode, some experiments for design optimization of neutralizer cathodes have been performed. A life test of the neutralizer cathode is being started in June 2008.

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To inspect numerical models for precise numerical analysis code development and to research ion thruster mechanisms (its discharge and ion beam optics behavior), a two-dimensional Kaufman-type visualized ion thruster was designed and fabricated. Its operations, probe measurements and spectroscopy measurements indicate that this thruster can visualize a schematic of ion production and extraction as described in conventional textbooks: ion beam extraction influences the ion density distribution in most of the discharge chamber; a close correlation exists between the ion beam focus position and the appropriate beam optics.

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High power ion cyclotron resonance heating is performed in a fast-flowing plasma operated with hydrogen and helium gases. Ion heating is clearly observed in hydrogen plasma as well as in helium plasma. The resonance region of magnetic field is broader and wave absorption efficiency is higher in hydrogen plasma than those in helium plasma. The thermal energy of the heated ions is converted to the kinetic energy of the exhaust plume by passing through a diverging magnetic nozzle set in a downstream region. In the magnetic nozzle energy conversion occurred as keeping the magnetic moment constant, but some discrepancy was observed in larger gradient of magnetic field. The kinetic energy of the exhaust plume is successfully controlled by an input power of radio-frequency wave, which is one of the key technologies for the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) type plasma thruster.

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Discharge current oscillation at the frequency range of 10-100 kHz causes serious problems in using an anode layer type Hall thruster in space. As a new approach to the discharge stabilization, azimuthally nonuniform propellant flows were created in an acceleration channel. A plenum chamber was divided into two rooms to which xenon gas was supplied at different flow rates. As a result, stable discharge was achieved in wider operation conditions with the larger mass flow differential. The oscillation amplitude at the maximum thrust efficiency was decreased by up to 94%. In addition, thrust vector was measured and the deviation angle was found to vary with the differential flow rate and magnetic flux density.

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We performed observations of the low-frequency discharge current oscillation of a Hall thruster by widely varying the external control parameters, and successfully clarified the range of parameter conditions required for oscillation-free stable operation. Furthermore, to control the parameters inside the stable operation region even during transient operations, we developed a method of synchronous control of the anode power conditioner and magnetic coil power conditioners. Upon applying this method of control to the newly developed power processing unit, Low-frequency oscillation-free operation was realized even for transient operations such as turning on the thrusters, changing the thrust condition, and turning off the thrusters. Also, the electromagnetic interference noise in the primary bus power lines was demonstrated to be sufficiently small.

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A new cathode with inductively coupled plasma for ion engines was designed and then evaluated experimentally in this study. The fabricated cathode, which named “ICP/C (Inductively Coupled Plasma Cathode)”, employed electromagnetic fields induced by RF (Radio Frequency) current at 13.56 MHz, and had the new ignition method that spark discharge occurred by induced voltage was used. Through the experimental investigations, the new method using “C-shaped” type electrode enabled the ICP/C to ignite instantaneously without other electron source. And the cathode achieved high neutralization performance, which was over 1700 mA of extracted electron current at 80 W of RF input power, and 3.0 sccm of Xe flow rate, when target voltage simulating ion beam was 20 V. Additionally, the application of ICP/C as a main cathode of direct current ion thruster was confirmed. As these results, ICP/C may permit instantaneous ignition and simple handling to ion engines, and it may even allow a longer lifetime of ion engines.

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A space propulsion system based on the acceleration of fusion ash is discussed by use of the energy balance equation and a hypothetic ash extraction and acceleration system. The fusion reactions D-T, D-³He, p-¹¹B and ³He-³He are investigated under the condition of thermal generation of high energy ions and equal plasma system conditions in terms of T_i/T_e relation and plasma beta. External plasma heating is defined by an equal efficiency concerning thermal energy conversion and energy transfer back into the plasma. There is no additional external heating applied to the fusion system. Power losses are based on neutrons, bremsstrahlung, synchrotron radiation and convection. We compare the plasma pressures, volumetric power densities, magnetic field strengths, heat waste, exhaust velocities and thrust density levels depending on the temperature and the hot ion mode. We show that, based on the fusion products, the exhaust velocity may reach several percent of speed of light in the case of ³He-³He. The temperature driven radiation losses of the ³He-³He reaction puts the purely aneutronic property into perspective. The mass flow rate densities of the considered fusion products are very low leading to very low thrust power densities. Considering the supposed system masses of a fusion based space vessel the thrust density levels are negligible and reach the order of 1 N/m³ near the optimum in the case of ³He-³He. We conclude that a propulsion system based on the acceleration of fusion products or ash is unfeasible for typical manned missions e.g. to Mars.

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A Pure Magnetic Sail is a deep space propulsion system consisting of a coil mounted on a spacecraft. In order to predict the thrust characteristics of a Pure Magnetic Sail in space, thrust measurement and magnetic field measurement were conducted using a scale model in a laboratory. To simulate the solar wind, a magnetoplasmadynamic arcjet provides a high density (2×10¹⁹ m^-3) and high velocity (47 km/s) plasma flow that impinges on a 20-turn 25-mm-radius coil simulating a Pure Magnetic Sail. When the magnetic cavity size of the scale model (L) is increased from 0.12 to 0.17 m, the thrust increases from 0.47 to 0.92 N. Scaling up, this experiment corresponds to a 300-km diameter Pure Magnetic Sail in space. The thrust also depends on the coil tilt angle, which is the angle between the direction of the solar wind flow and the coil axis. The maximum thrust of 1.5 N is obtained for a tilt angle of 90 degrees.

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A method of thermal analysis for a solar thermal thruster was created to aid in the thermal design of the thruster. The method consists of two types of thermal analysis: an analysis program for propellant flow, and an analysis of the temperature distribution of the thruster wall using Pro/Engineer. The numerical results were compared with experimental results to confirm the validity of the method, and there was good agreement between them. A thermal design was created using this thermal analysis method to estimate the performance of a solar thermal thruster for the orbital transfer of piggyback satellites mounted on an H2A rocket. When the thruster is made from heat-resistant steel and the propellant is water, the analytical results showed that the Isp is 203 s, the thrust is 16.6 mN, and the maximum temperature of the thruster is 1088 K. The diameter of the concentrator also was calculated, and it was found to be small enough for the concentrator to be mounted on piggyback satellites.

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Microfabrication of a 3 x 3 micro-plasmajet array with ultra-violet lasers and their thrust performance tests were conducted for nozzle elements with exit height of 0.5 mm and length of 1 mm. To evaluate the thrust characteristics of the nozzle-array, its thrust performance was compared with a single-nozzle. The thrust was measured by a calibrated cantilever-type thrust stand in vacuum. From the thrust performance tests, significant improvement of the thrust of each nozzle element with mass flow was obtained with the nozzle-arrays than those of only using a single-nozzle through multi-jet interaction phenomena.

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In a previous work based on numerical analyses and experiments, the effects of anode orifice configuration on the 20 kHz-range discharge current oscillation amplitude and the thrust performance have been analyzed. In the present work, for the purpose of more detail investigation of the influence of the orifice parameters on the current oscillation and the thrust performance, the ionization and acceleration process in the acceleration channel for various orifice positions were analyzed using a hybrid-PIC code. The results show that the amplitude of the oscillation, the ion loss to the channel walls, and the electron current has the largest value for the orifice set in the middle radial position.

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In this study, we are proposing a novel miniaturized ion engine system µ1. Recently microspacecraft and propulsion system to be installed there have attracted a lot of attentions. To accomplish the miniaturization of spacecraft component, multifunctionalization of devices are key technologies. The ion engine we are proposing here is distributed on microspacecraft and give a number of functions and strong redundancy to the spacecraft. To realize this concept, we introduced a novel idea for an ion engine system. That is to use single plasma source as both ion beam source and neutralizing electron source only by electrical connection. This ion engine system is released from the necessity of a number of neutralizers. Our concept requires a plasma source driven by very low power microwave. Here we proposed an antenna design method for a small plasma source using microwave discharge, and developed a miniaturized ion engine. As a result, the performance of the miniaturized ion engine was improved up to the ion production cost of 240 V and propellant utilization efficiency of 40 % at the input microwave power of 1.0 W and mass flow rate of 0.15 sccm.

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The total electrode fall voltage in a channel of magnetoplasmadynamic (MPD) thruster was determined by using "zero-limit approximating method", which is one of the classical methods widely used in the arc welding field. A new five-channel parallel-plate MPDT was designed and operated in a quasi-steady mode. This paper presents the measurement of the discharge voltage vs. the electrode gap for gaps from 1 mm to 4 mm. The extrapolated zero-gap intercept resulted in 18 V, which was considered as the total electrode-fall voltage. The electrode-fall voltage did not depend on the discharge current unless the discharge current exceeded onset threshold. It is considered that most of the electrode fall is deposited on the cathode side since the space potential at the inter-electrode region was almost equal to the anode potential.

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Simulation on the control of the discharge current oscillation is performed by introducing the cross-field transport of Xenon ions in the channel. The ion current oscillation and the thrust efficiency are shown by comparing our simulation with the observation. It is shown for the first time that the ion current becomes stable and the thrust efficiency is improved when the ion mobility decreases in the channel of Hall thrusters, by comparing with the experimental results of TALT-2a Hall thruster.

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To extend the application of the 20 mN class xenon ion engine which was used for north-south station keeping of a geosynchronous satellite--the engineering test satellite-VIII-, the ion thruster's thrust throttling performance was evaluated. The candidate target application is aerodynamic drag compensation of a super-low-altitude earth observation spacecraft. Performance test results show that the thruster, which passed through 12,000 h life test, exhibited a thrust throttling performance from 6 mN to 28 mN, although the thruster showed performance degradation from its initial performance before the life test. A power processing control unit is under development to meet the thrust throttling performance.

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In order to adapt to a wide variety of the space flights, such as small geosynchronous satellites and deep space explorers, feasibility study of performance enhancement options for the microwave discharge ion thruster µ10 is underway. Authors are considering the following five options: 1. Lower insertion loss DC blocks; 2. Direct monopole antenna insertion to the discharge chamber without using a circular waveguide part; 3. Optimization of gas injector layout which was originally located deep in the waveguide; 4. Additional magnet rings aiming ion loss reduction to the side wall of the discharge chamber; 5. New ion optics consists of a thinner screen grid and a smaller-hole accelerator grid. Not all but most of them have already been tested and reported in this article. The original models for Hayabusa asteroid explorer generated 8 mN at maximum. Larger thrust generation was impossible even if propellant flow rates and microwave powers were increased. It turned out to be feasible to increase the maximum thrust to a range of 10 - 11 mN with above mentioned options by supplying more flow rates and/or more microwave powers.

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Thrust performances in a microwave discharge ion thruster using argon as the propellant were investigated for possibility of alternative propellant of xenon. First, the ion beam currents using a 10 cm class were measured. The ion beam currents with argon propellant are larger than those with xenon propellant at P_i > 24 W in the same configuration. The propellant utilization and ion beam production cost are 0.50 and 440 W/A, respectively at 2.45 GHz microwave input power of 37 W and argon gas flow rate of 2.28 sccm. Next the ion beam currents of 2 cm class ion thruster were measured. The optimum mass flow rate is 0.4 sccm, which is larger than that using xenon. This would be due to the smaller ionization cross section and lighter mass than those of xenon. The propellant utilization efficiency and ion beam production cost are 0.73, 750 W/A, respectively at argon mass flow rate of 0.4 sccm and input power of 16 W. This performance of ion thrusters using argon competes against that using xenon.

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The Aerospace Research and Development Directorate of the Japan Aerospace Exploration Agency (JAXA) is considering a demonstration of electrodynamic tether (EDT) systems in low Earth orbit (LEO). Carbon nanotubes (CNTs) have some advantages as electron sources compared to conventional Spindt type emitters, and so are expected to be useful in EDT systems. Experiments to investigate the durability of CNT cathodes in a space environment had been conducted in a diode mode, but it was found that electron extraction tests, in which the cathode with a gate electrode is used, are necessary to evaluate the endurance of CNTs more accurately. In this paper, we conducted long duration operating tests of a cathode with a gate. It was found that there was almost no change in cathode performance at current densities below 100 A/m² even after the cathode was operated for over 500 hours in the high vacuum environment.

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The discharge oscillation in an anode-layer type Hall thruster is one of the most serious problems to be overcome. In general, a hollow anode is used in order to stabilize the operation. Recent numerical simulation results show the stable operation depends on the ionization rate in the hollow anode. In order to verify the simulation result, it is needed to measure the plasma profile in the hollow anode.
In this paper, as a first step, the number density and the velocity distribution of single-charged xenon ions in an anode-layer type Hall thruster were investigated by means of Laser-Induced Fluorescence (LIF) for the 5d²F_7/2→6p²D⁰_5/2 excitation transition at 834.7 nm detecting the non-resonant line to the 6s²P_3/2 state at 541.9 nm. As a result, we obtained the relative ion number density and the velocity distributions. The ion velocity distribution function shows both ionization and acceleration occur in the acceleration channel, while only ionization occurs in the hollow anode.

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Numerical analysis has been made for the plasma and flow properties in the microplasma source and micronozzle by developing a self-consistent numerical model, where a two-temperature fluid model was applied to the entire region through the plasma source to nozzle. Numerical results with 4-GHz microwaves and propellant gas of Ar 0.98 mg/s (33 sccm) revealed a spatially nonuniform distribution of plasma and flow parameters in the microplasma source: the absorbed power density, plasma density, and gas temperature were maximum at around the top of the microwave antenna, or near the inlet of the micronozzle, which is preferred for a thruster of electrothermal type using a converging-diverging nozzle.

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A scheme of numerical analyses available for the LSD wave has been constructed. As a typical analysis model, which is expected to realize the laser propulsion systems, axisymmetric-nozzle-geometry is applied. In the numerical analyses on LSD, especially, laser absorption model and thermal-nonequilibrium model are very important to describe physically detail processes. For thermal-nonequilibrium model, 2-temperature (heavy particle's temperature and electron temperature) model has been applied because the electronic mode is firstly excited in laser absorption and a thermal non-equilibrium state easily arises. In this paper, attention is paid for the numerical method for LSD waves propagating through a diatomic gas: hydrogen, nitrogen or oxygen, which is the representation of the gases existing in the nature and is expected to have an economical advantage for a propellant.

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Research and development on pulsed plasma thruster (PPT) for small satellites have been conducting in Tokyo Metropolitan University (TMU). In this study, improvement of total impulse on coaxial PPT was conducted in order to apply it to propulsion systems for small satellites. As a result, total impulse was improved from 3.5N-s to 17.5N-s by modifying thruster head configurations. Moreover, a PPT with propellant feed mechanism named Disk Feed PPT, which changes a used cylindrical propellant for a new propellant, was developed. Disk Feed PPT achieved total impulse of 54.6N-s. Furthermore, latest experiment of Disk Feed PPT showed the total impulse of over 150N-s.

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Several experimental and numerical simulation results for the development of miniature microwave discharge ion thrusters have been reported. The important points to improve the performance of the ion thruster are the shape of the antenna for microwave emission and the magnetic configuration in its thruster head. The extracted ion beam current of 12mA is obtained in conditions of input 2.45GHz microwave power of 8W and propellant (xenon) flow rate of 0.2 sccm with a star type antenna. A numerical simulation code coupling the particle-in-cell method and the finite difference time domain method was used to know plasma behavior in the discharge chamber of an ion thruster.

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The micro-Newton thrust measurement was conducted for a low intensity continuous wave laser thruster for aluminum foil propellant target. To evaluate the thrust characteristics, a torsion-balance thrust stand capable for the measurement of the thrust level down to micro-Newton range was developed. In the case of the aluminum foil target, the maximum thrust level was 6 μN when the laser power was 20 W. It was also found that the laser intensity, or laser power per unit area, irradiated on the target was significantly important on the control of the thrust even under the low-intensity level.

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A miniature microwave discharge neutralizer has been developed in order to develop the miniature ion engine system. Extracted electron currents from the microwave discharge neutralizer are measured for two neutralizers, cylindrical discharge chamber with monopole antenna and the same discharge chamber as miniature microwave discharge ion thruster developed at Kyushu University for two orifice diameters. The former shows that the extracted electron current is 14 mA at incident microwave power of 4 W, 6 mA at incident microwave power of 2 W at mass flow rate of 0.005 mg/s. This small extracted current would be due to the large loss on the discharge chamber wall. On the other hand, the best extracted electron current of the latter, is 11 mA at mass flow rate of 0.005 mg/s and incident microwave power of 4 W, by optimizing the orifice diameter or position of the orifice plate. From the poor performance at incident microwave power of 2 W, the plasma- microwave coupling is worse than that of the former.

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In this paper, identification of an impact force acting on a CFRP stiffened panel is conducted by two different impact force identification methods, i.e. the model-based identification method and the experimental identification method. Both methods use the relations between the impact force and strain responses at measurement positions. The two methods differ, depending on whether the relations are determined analytically using the finite element method or experimentally from impact tests. By employing the relations, the location and force history of the impact force are identified by adjusting the corresponding strain responses to the measured strain responses. The validity and effectiveness of the two identification methods are verified by comparing the identification results with the actual impact force acting on the CFRP stiffened panel.

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For damage detection in CFRP laminates, health monitoring techniques using Lamb waves are effective. In this research, the authors constructed a new system combining a macro fiber composite (MFC) actuator and a fiber Bragg grating (FBG) sensor to employ broadband Lamb waves from DC to 1MHz. Then this system was applied to a CFRP quasi-isotropic laminate to examine the broadband characteristic. As a result, this system proved to be able to send and receive more broadband ultrasonic wave than conventional systems using bulk PZTs. Moreover, we investigated the frequency dispersion characteristic of Lamb waves propagating in the CFRP laminate in more detail by sending narrowband ultrasonic waves with different center frequencies. As a result, we could identify the modes of Lamb waves, comparing the measured frequency dispersion curves with the theoretical dispersion curves. After that, the dispersion curves were measured for the CFRP laminate that included an artificial delamination. Since the result measured in the damaged laminate had different dispersion from the result in the intact laminate, the delamination could be detected from the frequency dispersion curves. Hence this system can obtain more effective information for damage detection.

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