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Award type：Honored in official journal of a scientific society, scientific journal  Country：United States

Award type：Honored in official journal of a scientific society, scientific journal  Country：United States

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Authorship：Last author   Language：English   Publishing type：Research paper (scientific journal)   Publisher：Springer Science and Business Media LLC  

Abstract

Interplanetary transport of payloads of unprecedented mass, as envisaged beyond the lunar gateway, will require thrusters with high specific impulse as well as high thrust. To achieve this, innovations in propulsion are critical. Many classes of electric thruster utilise a magnetic applied field module to accelerate charged particles. Magnetoplasmadynamic thrusters exhibit improved performance with increasing field, at least up to the limit of around 0.5 T able to be provided by permanent magnets or copper electromagnets. However, superconducting magnets can generate much stronger magnetic fields. In this study, we utilised a space-relevant cryocooled high temperature superconducting magnet as the applied field module for a central cathode electrostatic thruster (CC-EST). A convex anode enabled ignition at high magnetic fields, and in this configuration the thruster’s performance was characterised in the power range of 1 kW to 2.5 kW and at steady applied fields ranging from 0.6 T to 0.8 T, representing a significant advance in achievable field strength. In combination, these operating parameters enabled the achievement of a magnet-inclusive thruster efficiency of 19%, while the cryocooled magnet was demonstrated to be thermally stable in the presence of the kW-scale plasma, demonstrating the viability of such a design for space flight applications.

DOI： 10.1007/s44205-023-00060-z

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Other Link： https://link.springer.com/article/10.1007/s44205-023-00060-z/fulltext.html

Authorship：Corresponding author   Language：English   Publishing type：Research paper (scientific journal)   Publisher：AIP Publishing  

Hollow cathodes are important devices used for spacecraft electric propulsion. The hollow cathode has two operational modes. One mode is a stable mode called the spot mode, and the other is an unstable mode called the plume mode. Operation in plume mode should be avoided since the instability causes high-energy ions that sputter-erode the cathode parts. In this study, the relationship between discharge oscillations and ion energy distribution in plume mode was investigated using a triple Langmuir probe and retarding potential analyzer for a 40-A class xenon hollow cathode with a lanthanum hexaboride emitter. The triple probe can measure unsteady electron temperature and plasma density oscillations. The electron temperature was not so high, 1 to 2 eV. Some instabilities were observed in the plume mode. The ionization instability with a low frequency oscillation of 30 kHz was the dominant mode. A broad spectrum around 330 kHz due to ion acoustic turbulence was observed. In addition, in the downstream plume region, oscillations around 120 kHz were observed owing to temporal change in anomalous resistivity. The 95% ion population voltage found to be 20 and 30 eV in spot and plume modes, respectively. The magnitude of the low frequency ionization oscillation was found to be inversely proportional to ion energy in plume mode. This indicates that the resonant energy transfer from the oscillation to the ion energy through Landau damping probably plays an important role in high energy ion generation in plume mode. A clear correlation between discharge current and electron temperature waveforms was found. The larger the electron temperature fluctuation, the stronger the correlation between discharge current and electron temperature, and the larger the phase difference deviation from 180°.

DOI： 10.1063/5.0139089

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Authorship：Corresponding author   Language：English   Publishing type：Research paper (scientific journal)  

A 100-W class low power arcjet thruster using a hollow cathode was proposed, with the goal of reducing cathode erosion. The hollow cathode's emitter was lanthanum hexaboride expecting its excellent thermionic emission performance. The experiment used argon as a propellant and compared two different emitter lengths. The estimated emitter temperature was almost the same regardless of the emitter length, however, due to of the efficient heat exchange with the longer arc plasma region, the longer emitter provided better performance with higher discharge voltage. Additionally, the hollow cathode was compared to the conventional rod cathode, and the hollow cathode was found to have some advantages. The specific impulse of the hollow cathodes was higher than that of the rod cathode at low flow rate. The results showed that the hollow cathode is best for high specific power and high specific impulse operation. The unstable transition between low- and high-voltage modes has been observed in the rod cathode, but not in the hollow cathode configurations. Helium was put to the test as a propellant and compared to argon for simulate the future use of hydrogen as a propellant. The thrust efficiency was lower due to the high discharge voltage and significant heat loss, but the specific impulse was higher as expected. No severe recession or erosion of the emitter was observed after the test campaign, but some surface color change was seen which may affect in the long-term operation.

DOI： 10.1016/j.actaastro.2023.02.015

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Authorship：Corresponding author   Language：English   Publishing type：Research paper (scientific journal)  

Dry ice has been proposed as a propellant for electric propulsion, as its use may facilitate the realization of a low-cost propulsion system compared to that based on a xenon propellant. In addition, dry ice can be obtained from space (e.g., from the atmosphere of Mars) and human exhalation. This study thus conducted experiments with dry ice storage and carbon dioxide gas supply using triple-point storage. With this method, a constant tank pressure and flow rate was achieved without the application of regulators or control valves. However, the triple-point state was unexpectedly halted, even though solid dry ice remained in the tank. An internal visual observation revealed that the solid phase was surrounded by the liquid phase, which prevented the solid phase from melting. Therefore, heat conduction fins were installed in the tank to enhance the melting process, which successfully extended the triple-point duration. The supply system was applied successfully to the operation of a 1 kW Hall thruster. Furthermore, a mathematical model for estimating the phase characteristics in the tank was developed, with an in-orbit operation analyzed based on the experimental results. The heat input to the tank and the mass flow rate or enthalpy to the thruster determined the phase characteristics in the tank, which means that implementing a thermal design that considers the flow rate to the thruster is crucial. The simulation demonstrated that dry ice can be consumed by controlling the three phases in the tank using an electrical heater. A gas phase separator concept consisting of the heat conduction fins and capillary screen mesh was proposed for orbital microgravity operation.

DOI： 10.1016/j.actaastro.2022.10.034

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Authorship：Corresponding author   Language：English   Publishing type：Research paper (scientific journal)   Publisher：The Japan Society of Mechanical Engineers  

Vapor cooling shield is a key technology for long-term cryogenic propellant storage in space. Cooling channels with a triply periodic minimal surface embedded inside are expected to improve its cooling performance. Herein, a cooling channel with a triangle cross-section embedded gyroid structure, a type of the triply periodic minimal surface structure, was additively manufactured. The pressure drop and heat transfer performances of the channel were experimentally measured using liquid nitrogen vapor. Furthermore, in addition to the gyroid-embedded channel, three channels with different cross-sections were fabricated for comparison: circular, triangle, and triangle with a step/groove on the bottom. The gyroid-embedded channel exhibited unique characteristics, with a thermal conductance that was approximately 40% higher than that of the channel with a simple triangle cross-section, but an excessive pressure drop of approximately 50 times higher than that of the other cross-sections. This denotes that strong vortex and turbulence and the flow separation cause excess pressure drop in the gyroidembedded channel. The pressure drop characteristic of the gyroid-embedded channel against the Reynolds number completely differed from that of the other channels, and the pressure drop of the gyroid-embedded channel can be estimated assuming analogy with particle packed beds.

DOI： 10.1299/mej.23-00015

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Language：Japanese   Publishing type：Research paper (scientific journal)   Publisher：THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES  

DOI： 10.14822/kjsass.70.7_145

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Authorship：Lead author   Language：English   Publishing type：Research paper (scientific journal)  

A cooling system using a cryocooler and boil-off gas was proposed for cryogenic propellant storage on the lunar surface. The cryogenic tank was covered with an advanced non-interlayer-contact multi-layer-insulation and coolant from the cryocooler and boil-off gas flow in the layers to remove the heat load. The proposed system included optimization factors for fluid tubes location, cryocooler temperature, and flow rate. A simple uniform-temperature thermal model was developed to optimize boil-off rates and system weights, and the results revealed a Pareto front. A three-dimensional model was developed to discuss the effect of temperature nonuniformity. Finally, the proposed concept was compared to other candidates. The results showed that the proposed system offered a well-balanced system in terms of boil-off rate, system mass, and electric power consumption, which drastically reduced the boil-off rate when compared to only multi-layer insulations, which reduced the weight and power consumptions as compared to zero-boil-off systems.

DOI： 10.1016/j.cryogenics.2022.103494

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This study numerically reveals effect of pressurant gas species on thermal behavior in a cryogenic propellant tank for space propulsion systems. The simulation was conducted for a ground experiment using liquid nitrogen in which gaseous helium and the same species as the liquid, gaseous nitrogen, were used as prepressurants. The tank was sealed after prepressurization to observe self-pressurization (pressure rise with evaporation) and thermal stratification (high-temperature liquid near the surface). A higher evaporation rate and a thicker thermal layer were observed for helium prepressurization in the experiment. To clarify the underlying physics, a simulation considering phase change and conjugate heat transfer was developed. The simulated pressure and temperature closely matched the experimental results. Conjugate heat transfer played an important role; that is, the heat flow through the ullage part of the tank skin into the cryogenic liquid was a dominant heat transfer process, causing evaporation to occur mainly at the contact point of the wall and the gas–liquid interface. Afterward, nitrogen vapor rose along the tank wall due to buoyancy under nitrogen prepressurization. However, buoyancy in the ullage is lower under helium prepressurization since light helium stayed in the upper part of the ullage, and a radial vapor flow is produced from the contact point, leading to a higher heat flux from the tank wall to the liquid nitrogen. This mechanism shows that evaporation rate and heat flow into the liquid are higher for helium prepressurization than for nitrogen prepressurization.

DOI： 10.2514/1.A35061

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Authorship：Corresponding author   Language：Japanese   Publishing type：Research paper (scientific journal)   Publisher：THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES  

For future space transportation missions beyond earth orbit, propulsion systems applying cryogenic propellants can achieve efficient and rapid in-space mass transfer with high specific impulse. Refilling of cryogenic propellant generated on Moon surface also enhances the advantage. In this study, we propose a new cryogenic liquid rocket engine system for efficient multiple ignitions for such missions. Two engine operation modes, high-specific-impulse idle mode and low-pump-rotation throttling mode, are discussed using newly developed cycle analysis tool. State-of-the-art electric motor driven valves which continuously control the flow rates can realize these operation modes. The cycle analysis was conducted based on the expander bleed cycle considering combustion and turbopump instabilities and chamber wall temperature limit. The analytical results show the high-specific-impulse idle mode can offer higher specific impulse comparing to the conventional operation. During the low-pump-rotation throttling mode, the discharged propellant is used for tank pre-pressurization; therefore, helium consumption for tank pre-pressurization can be drastically reduced. Possible operational points with high specific impulse were found in the low-pump-rotation throttling mode with additional valve and piping to maintain the pump flow coefficient. These new engine operation modes offer less chill-down and helium consumption before ignitions and result in efficient in-space orbital transportation.

DOI： 10.2322/jjsass.70.110

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Language：Japanese   Publishing type：Research paper (scientific journal)  

To achieve high-thrust-density operation, we propose electrostatic-magnetic hybrid ion acceleration in which the empirical thrust density limit of the electrostatic acceleration is surpassed without violent plasma oscillation by combing the collisional momentum transfer mechanism, which is the ion acceleration mechanism of the electromagnetic acceleration. To achieve hybrid ion acceleration, we experimentally obtained two design criteria: one near anode propellant injection and another at the on-axis hollow cathode location. The thrust characteristics of three thrusters composed of a slowly diverging magnetic field between an on-axis hollow cathode and a coaxially set ring anode were examined. By injecting xenon propellant along the anode inner surface, the electron impact ionization process was enhanced, and generated ions are electrostatically accelerated through the radial-inward potential gradient perpendicular to the axial magnetic lines of force. The hybrid ion acceleration characteristics were obtained only if these two criteria were satisfied and the obtained thrust was consistent with the thrust formula derived for steady-state, quasi-neutral plasma flows. In addition to the criteria, strengthening the magnetic field and enhancing the propellant mass flux were effective for improving thrust density without deteriorating thrust efficiency. Among the experimental conditions in this study, the maximum thrust density was 70 N/m2 with an anode specific impulse of 1200 s, which cannot be achieved in a purely electrostatic thruster with thrust density 6.3 times than that of a typical Hall thruster.

DOI： 10.1063/5.0066083

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Language：English   Publishing type：Research paper (scientific journal)   Publisher：American Institute of Aeronautics and Astronautics (AIAA)  

DOI： 10.2514/1.b38480

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In this study, a novel single-piece thin multi-layer tungsten resistive heater was successfully fabricated using additive manufacturing and tested as an electrothermal thruster. The heater has 12 resistive layers, with each layer having a thickness and height of 0.15 and 81 mm, respectively, and can provide high heating efficiency. A single-piece or monolithic heater was manufactured via additive manufacturing technique, which drastically improved its reliability and decreased its manufacturing cost. In the heating and thrust measurement tests that used nitrogen gas as a propellant, the heater reached a gas temperature of ∼2000 K at a 140-A heater current without experiencing any failure. The tungsten-heater resistance linearly increased with an increase in temperature due to the temperature dependence of tungsten's resistivity. The specific impulse and thrust increased with the heater temperature in accordance with the theoretical prediction. Even including a voltage drop due to a contact resistance, the achieved heater efficiency reached 63% at a 100-A heater current even without a thermal insulation around the thruster. The heater efficiency decreased with an increase in the heater temperature due to heat loss to the surroundings. The heat-loss analysis indicated that both thermal conduction and radiation heat losses were crucial for improving the heater performance at a high-temperature operation of over 2000 K.

DOI： 10.1063/5.0068926

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Language：English   Publishing type：Research paper (scientific journal)   Publisher：American Institute of Aeronautics and Astronautics (AIAA)  

A high-efficiency concentric tubular-type resistojet with potential application to short-term orbit-raising maneuvers has been fabricated by 3-D printing and demonstrated. The propellant flows through multiple layers of cylindrical shells, with this structure also functioning as a single-piece heater. A 6-cm-high cylinder was realized with a wall thickness of 0.2 mm, using Inconel 718. A nodal thermal analysis was performed to identify the upper-limit current at a temperature limit of the wall material, and it was revealed that an outlet gas temperature of 871 K can be achieved with 77 A of current at 0.2 g/s of mass flow rate. The designed heater was combined with a boron nitride insulator and a stainless-steel housing, and thrust was measured in a vacuum chamber with nitrogen as the propellant. At a mass flow rate of 0.2 g∕s and 75 A of current, an outlet temperature of 747 K, a specific impulse of 108 s, and a heater efficiency of 72% were achieved. These results with nitrogen propellant were used to predict the performance of a tungsten-made resistojet with a hydrogen propellant, and a specific impulse of over 700 s can be expected at a heater temperature of 2000 K.

DOI： 10.2514/1.b38187

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Language：English   Publishing type：Research paper (scientific journal)   Publisher：American Institute of Aeronautics and Astronautics (AIAA)  

In this paper, the authors investigate whether airbreathing engines have useful application to vertical takeoff and vertical landing systems, which currently represent the mainstream for reusable launch vehicles. A theoretical analysis has been performed to determine the net impact of the specific impulse benefits and weight penalties of a vertical takeoff reusable launch vehicle fitted with an airbreathing propulsion system. To maximize the thrust-to-weight ratio of an airbreathing engine, the authors propose using a conventional fan driven by a separate gas generator in lieu of a conventional core airbreathing combustor, as well as combined rocket–airbreathing operation to reduce engine size. The proposed engine system and its propulsive performance are described herein. Furthermore, a new sounding rocket is described as a practical application for the proposed vertical takeoff and vertical landing airbreathing engine. Conventional horizontal takeoff and horizontal landing airbreathing engine concepts tend to focus on maximizing specific impulse and airspeed, which necessitates further development, specifically in the areas of heat-resistant materials and structural technology. By proposing a vertical takeoff and vertical landing airbreather in this study, a different approach is taken and the potential to significantly improve launch capability and reliability is demonstrated in comparison to the current technology level for reusable launch vehicles.

DOI： 10.2514/1.a34988

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Language：English   Publishing type：Research paper (scientific journal)   Publisher：PERGAMON-ELSEVIER SCIENCE LTD  

This paper reports measurements of the xenon ground state and excited state densities inside a mu 10 microwave ion thruster. This thruster exhibits a 40% thrust enhancement upon changing from the waveguide to the discharge chamber propellant injection mode. In the present work, the associated mechanism was quantitatively evaluated using two-photon laser induced fluorescence (TALIF) spectroscopy to monitor the thruster waveguide. The 834.7 nm emission from excited state xenon was investigated with a 224.3 nm dye laser to excite the Xe I 5p61 S0 6pMODIFIER LETTER PRIME [3/2]2 state, compared with the emission without the laser. The resulting data confirm that the neutral density exhibits a linear relationship with the propellant flow rate in the cold gas and ionized state, while the ion acceleration decreases the neutral density by the same order of magnitude as the propellant utilization efficiency is changed. As the propellant flow rate increases, the collisions of neutrals that generate excited states occur in the waveguide and, when this process plateaus, the ground state emission suddenly increases. Propellant injection from the discharge chamber is evidently effective at suppressing collisions with electrons in the waveguide that generate excited states and that potentially interfere with microwave propagation.

DOI： 10.1016/j.vacuum.2021.110269

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DOI： 10.1016/j.actaastro.2020.12.047

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Abstract: An advanced background-oriented schlieren (BOS) method, named as the speckle beam-oriented schlieren technique, was newly developed to measure the distribution of refraction angles in transparent media. A speckle pattern is generated by passing a coherent laser beam through a holographic diffuser, a pinhole, and a lens, generating a collimated background image that is projected directly onto the image sensors. Since the intensity of the background image is maintained at a high level, this method is, in principle, useful for diagnosing fast and/or low signal-to-noise phenomena, such as high-temperature gasses with radiation emission. Moreover, by splitting the background beam into two imaging paths with different focal lengths, the refraction angles can be measured for a schlieren object with uncertain location, and the depth position of the refraction angles can be resolved. This technique was demonstrated by measuring the refraction angle and the depth position distribution in a sonic jet with different injected locations. Graphic abstract: [Figure not available: see fulltext.]

DOI： 10.1007/s00348-020-03113-3

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Authorship：Lead author   Language：English   Publishing type：Research paper (scientific journal)   Publisher：American Institute of Aeronautics and Astronautics (AIAA)  

This Paper explores keeper ignition and discharge characteristic of a hollow cathode center mounted on a Hall thruster to discuss the optimum startup sequence. A strong magnetic field along the hollow cathode in the centermount configuration is considered to play an important role in the ignition and discharge; however, the characteristics in this configuration has never been reported. The effect of cathode mass flow rate on the ignition and discharge with 3Akeeperignitionis examined both with and without magnetic field. The influence of cathodeheater power and flow from the anode is also investigated. The lower flow rate causes strong ionization instability, while the higher flow rate seems to induce ion acoustic turbulence. Applied magnetic field improves the ionization instability due to the decrease in electron temperature through plasma production. These results indicate ignition sequences over 1.5 mg=s flow rate with magnetic field appear promising. On the other hand, the confined electrons around the magnetic field line prevent the initiation of the discharge. The ignition and instability are improved with high heat-up power because more thermionic emission is obtained. The anode flow provides slight improvement to the characteristics. The Fourier transform also demonstrates the presence of ionization instability and ion acoustic turbulence. The behavior of peak frequencies of ionization instability can be explained based on the predator-prey theory.

DOI： 10.2514/1.B37814

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Language：Japanese   Publishing type：Research paper (scientific journal)   Publisher：Transactions of the Japan Society for Aeronautical and Space Sciences  

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DOI： 10.2322/tjsass.64.189

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Authorship：Lead author   Language：English   Publishing type：Research paper (scientific journal)  

The thermal behavior of cryogenic propellant tanks is crucial issue in the operation of cryogenic propulsion systems. Herein, ground experiments were conducted in a 600-mm-diameter cryogenic tank filled with liquid nitrogen. As pre-pressurants, gaseous helium and gaseous nitrogen (of the same species as the liquid), were used to investigate its effect in accordance with actual propulsion systems. The tank was sealed after pre-pressurization to observe the self-pressurization. The evaporation rate and heat flow in the tank were estimated based on pressure and temperature measurements. In addition, the axial liquid temperature distribution was obtained through the liquid draining from the tank bottom, and a thermal stratification model was developed. These results demonstrated that the type of pre-pressurant significantly affected the thermal behavior in the tank. A higher evaporation rate and higher liquid internal energy rise rate with a thicker thermal layer were observed in the helium pre-pressurization case. The lower nitrogen partial pressure in the helium case enhanced the vaporization and growth of the thermal layer. Estimation of the power balance in the tank demonstrated that not only the ullage but also the heat mass of the tank provided heat for the evaporation and thermal layer. The evaporation occurs mainly at the contact point between the tank skin and liquid surface. The nitrogen vapor rises in a thin layer along the tank skin because of buoyancy under nitrogen pre-pressurization; however, buoyancy is lower under helium pre-pressurization, and a radial vapor flow is probably produced from the contact point instead, leading to a higher heat flux to the liquid.

DOI： 10.1016/j.cryogenics.2020.103196

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A hollow cathode is an efficient electron source in the self-heating mode utilizing the discharge power. However, in sub-ampere currents, it needs keeper power to maintain the thermionic electron discharge, which could decrease the thrust efficiency. To address this problem, we propose using a microwave cathode, which is based on the flight model of a microwave ion thruster neutralizer cathode, as an alternative to a hollow cathode. First, we redesigned the magnetic field of a microwave cathode discharge chamber and tested it in the diode mode configuration. The electron emission current is doubled compared to the original performance. Next, we coupled the improved microwave cathode with a 200-W class Hall thruster and compared the characteristics and performance with a hollow cathode. We confirmed that the magnetic field polarity affects the ignition characteristics. We measured the thrust by an inverted thrust stand, the ion energy distribution functions by a retarding potential analyzer, and the beam profiles by an ion collector. The thrust and thrust efficiency are equivalent for both types of cathode. The specific impulse is 10% higher in the case of the microwave cathode. Since the potential difference between the microwave cathode and ground rapidly increased at currents above 600 mA, this could be taken to be the trade-off point against the hollow cathode.

DOI： 10.1016/j.actaastro.2020.06.049

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DOI： 10.1016/j.cryogenics.2019.102996

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A low-cost and reliable propellant flow controller is necessary to offer competitive electric propulsion systems. This paper first reviews xenon flow control methods for Hall thrusters and ion engines. Through a trade study, a simple and low-cost xenon flow controller applying a heated capillary is proposed. The test article is fabricated based on a developed theory and tested with xenon. The estimation based on laminar flow shows slightly different characteristics from those of the experiment, although the Reynolds number indicates a laminar condition. The turbulent flow assumption is in good agreement with the experiment at a high mass flow rate while the flow characteristics are between laminar and turbulent at a low mass flow rate. That implies that the surface roughness and tube curvature induce the laminar–turbulent transition even in the low mass flow rate case. To obtain a wider flow rate range, maintaining laminar flow in the capillary is desirable because of the high sensitivity of the friction coefficient with temperature in the laminar flow. Component improvements achievable through miniaturization and reduction of the inner wall surface roughness and capillary curvature are considered to be necessary to prevent the turbulent transition.

DOI： 10.2514/1.B37726

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Language：English   Publishing type：Research paper (scientific journal)   Publisher：JAPAN SOC AERONAUT SPACE SCI  

To investigate the neutral xenon density distribution of electric thrusters such as ion and Hall thrusters, two-photon absorption laser-induced fluorescence (TALIF) spectroscopy was applied to a microwave cathode. First, the background pressure of the vacuum chamber was measured by TALIF. In the present measurements, the ground state was excited by a 224.29 nm laser, and 834.68 nm fluorescence was detected. The first measurement confirmed that the fluorescence intensity linearly increases with respect to the ground state number density. Based on this result, the density of neutral ground-state xenon was measured at the exit of the nozzle of the microwave cathode. The variation in the density with the microwave power was successfully measured at xenon flow rates of 0.029 and 0.098 mg/s. The measured densities varied from 2.3 x 10(19) to 8.4 x 10(19) m(3) with a maximum error of +/- 20% due to the plasma fluorescence.

DOI： 10.2322/tjsass.63.281

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DOI： 10.1016/j.actaastro.2019.09.025

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This paper reports the first study to measure xenon neutral ground state particle density of microwave cathode by two-photon laser induced fluorescence spectroscopy (TALIF). Xenon is commonly used as a propellant in electric propulsion like Hall thrusters, ion thrusters, and their cathodes. For electric propulsion, information about neutral particles is important such as the ionization degree and the charged exchange collisions (CEX). The measurement target is XeI 5p S 6p[3/2] , which absorbs at a wavelength of 224.29 nm and emits fluorescence of 834.7 nm. The measurement system was demonstrated for three cases: cold gas, without electron extraction, with electron extraction. From three cases, the measurement system can detect a neutral ground state particle density of 10 m order without and with a plasma. In a cold gas, the neutral ground state particle density is (8.4±0.4)×10 m at 0.098 mg/s. Without electron extraction, the neutral ground density decreases by ionization and excitation With electron extraction, the density varied from 0.6 to 2.3 times compared to without electron extraction depending on anode voltage. 61 19 −3 19 −3 0 2

DOI： 10.1016/j.vacuum.2019.108846

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Electrostatic plasma thrusters such as ion engines and Hall thrusters commonly use xenon as a propellant and several measurement techniques for xenon ions and metastable neutrals have been applied to evaluate the characteristics of the thrusters. Although density measurements of ground state neutral xenon can provide crucial information on the ionization characteristics and help explain charge exchange phenomenon, much less research is available due to its technical difficulty. Two-photon absorption laser induced fluorescence is promising because it allows access to ground state xenon atoms by using around 220–260 nm wavelength lasers which have become more readily available lately. In this study, observation of the fluorescence following two-photon excitation from a room temperature (cold) xenon gas cell is conducted with 249 and 252 nm wavelength excitation at xenon pressures of 0.1 and 10 Torr. The fluorescence signals are obtained against a wide range of laser intensities, and the resulting fluorescence response comprises of a few regimes – weak-excitation, saturation, and an intermediate regime. The natural lifetime and quenching rate are evaluated by analyzing the fluorescence decay, and the result is consistent with published literature. Finally, actual application to ground tests of Hall thrusters is discussed based on the experimental results, especially with respect to their fluorescence responses.

DOI： 10.1016/j.actaastro.2019.03.018

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DOI： 10.1088/1757-899X/502/1/012062

Authorship：Lead author   Language：Japanese   Publishing type：Research paper (scientific journal)   Publisher：一般社団法人 日本航空宇宙学会  

DOI： 10.14822/kjsass.67.6_227

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The sharp leading edges of a hypersonic vehicle are promising for the enhancement of its aerodynamic performance. However, they need a cooling mechanism because they are exposed to extremely high heat flux during reentry or hypersonic cruise, which easily heats up the surface materials beyond their temperature limits. This study explores the concept of a leading edge with planar heat-pipe cooling for which the whole inner surfaces are directly covered with a capillary wick. Its isothermal characteristic and simple configuration can provide robustness; however, due to the uniqueness of the planar type, the thermofluid characteristics should be investigated. Therefore, two kinds of heating tests (the inside visualization test and high-temperature test) were conducted using water as a working fluid. The gravity effects were surveyed, changing the test article attitude. The results showed that the planar concept worked as expected; no large temperature distribution was observed along both the chord and span directions, even in a nose-up attitude or a top heat configuration; the heat transfer performance was improved in nose-down attitude. During the heat-pipe operation, a heat transfer coefficient approximately 100 times higher than that of an evacuated operation was obtained. Visualizing inside the test article revealed that the screen mesh wick was properly covered with condensed water and that typical bubble formations of nucleate boiling were observed around the tip in some cases.

DOI： 10.2514/1.A34274

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Language：Japanese   Publishing type：Research paper (scientific journal)   Publisher：一般社団法人 日本航空宇宙学会  

&lt;p&gt;Many space vehicles are powered by liquid hydrogen and liquid oxygen. Such fuel are cryogenic fluids, so they are easy to boil and become gas-liquid two phase flow. The LE-5B-3 engine has the capability of the idle mode firing same as the LE-5B-2 engine. Assessment of flow condition at the inlet of fuel turbo pump is important to operate the engine, because the fuel may flow in saturated condition under the idle mode in principle. In a two-phase flow state, void fraction is one of the most important parameters to assess the flow. Although many types of void fraction sensors were proposed, the capacitive technique has advantages to mount on the engine from the viewpoint of size, weight, toughness. In this study, plural circular electrodes capacitive void fraction sensor is developed for LE-5B-3 engines&#039; ground firing test. The sensor was designed based on electric field analysis, and the specification was assessed prior to the ground test. The sensor was used in qualification test, and it was succeeded in achieving stable measurement and it helped to understand the fluid state during the engine operation. The sensor design technique, the assessment results and the ground test results are discussed in this paper.&lt;/p&gt;

DOI： 10.2322/astj.JSASS-D-18-00017

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Torealize high-performance cryogenic propulsion systems, the chilldown sequence has to be improved. Because the chilldown is carried out under low gravity, the effect of gravity on the two-phase flow, especially at low flow rate, should be investigated. To understand the physics under low gravity, an experiment was conducted using a sounding rocket. Two identical test sections with different mass flow rates simulated part of a turbopump, each of which has a complex flowpath including slits and a dead end. Using liquid nitrogen, the flight experiment obtained data of temperatures, pressures, void fractions, and video frames of liquid motion. Then, the flight experiment data were compared to the ground data taken under normal gravity, revealing that the slits played an important role in the chilldown process and that the test sections were quickly chilled down under low gravity. The slits of the test sections formed liquid jets, and their behaviors were different from those in the ground experiment. In the flight experiment, the jets easily reached the dead end of the test sections and cooled down the whole walls due to the increase in inertia and wettability; however, such behaviors were hardly observed in the ground experiment. The difference between the ground and flight is significant at lower flow rate.

DOI： 10.2514/1.A34222

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This study numerically explores the flow physics associated with nanosecond pulsed plasma actuators that are designed to control shock-wave induced boundary-layer separation in a Mach 2.8 supersonic flow. By using two dielectric barrier surface discharge actuator configurations, parallel and canted with respect to the flow velocity vector, a previous experiment suggested that the actuator worked in two ways to influence the interaction: boundary layer heating and vorticity production. The heating effect was enhanced with the parallel electrode and made the boundary-layer separation stronger, while the canted electrode produced vorticity and suppressed the boundary-layer separation due to the momentum transfer from the core flow. Because the detailed physical processes are still unclear, in this paper a numerical investigation is undertaken with a large eddy simulation and an energy deposition model for the plasma actuation, in which the dielectric barrier discharge produced plasma is approximated as a high temperature region. The flow characteristics without the plasma actuation correspond to the experimental observation, indicating that the numerical method successfully resolves the shock-wave/boundary-layer interaction. With the plasma actuation, complete agreement between the experiment and calculation has not been obtained in the size of the shock-wave/boundary-layer interaction region. Nevertheless, as with the experiment, the calculation successfully demonstrates definite difference between the parallel and canted electrodes: the parallel electrode causes excess heating and increases the strength of the interaction, while the canted electrode leads to a reduction of the interaction strength, with a corresponding thinning of the boundary layer due to the momentum transfer. The counter flow created by the canted actuator plays an important role in the vortex generation, transferring momentum to the boundary layer and, consequently, mitigating the shock induced boundary layer separation.

DOI： 10.1063/1.5051823

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Authorship：Lead author   Language：English   Publishing type：Research paper (scientific journal)   Publisher：American Institute of Physics Inc.  

Space propulsion systems use screen mesh devices as filters to block contaminants and as propellant management devices to settle the propellants. The bubble point pressure indicates the basic capillary performance for liquid acquisition of screen meshes. Actively controlling the bubble point pressure can result in flexible and efficient operation of the propulsion systems. High-performance cryogenic propellants, such as liquid hydrogen and oxygen, exhibit magnetic properties. Therefore, a method to actively control the bubble point pressure of cryogenic propellants by applying a magnetic field is proposed in this study. The magnetic pressures affect the pressure balance around the gas-liquid interface separated by the screen mesh, which can thereby control the bubble point pressure. To demonstrate the concept and theoretical basis, a bubble point experiment is conducted using a ferrofluid and solenoid. This experiment proves that the magnetic field actively controls the bubble point pressure and performs both suppression and enhancement of the liquid acquisition performance of the screen mesh. The theory related to magnetic pressures is observed to successfully predict the experimental results. The feasibility of the active control of the bubble point pressure of liquid oxygen is discussed based on the validated theory, and two applications of this technique in cryogenic propulsion systems are depicted.

DOI： 10.1063/1.5034222

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Language：Japanese   Publishing type：Research paper (international conference proceedings)   Publisher：プラズマ・核融合学会編集委員会  

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Language：Japanese   Publishing type：Research paper (scientific journal)   Publisher：一般社団法人 日本航空宇宙学会  

&lt;p&gt;Reducing the amount of propellant for re-cooling is an important issue for the rocket propulsion system using cryogenic fuel. Immediately after the start of the engine, the liquid fuel boils and becomes two-phase flow. In the state of two-phase flow, the void fraction, which is the gas-liquid ratio, is one of the important value for flow control. For above problem, we are developing void fraction measurement system for the cryogenic fluid. These devices were attached to the S310-43 sounding rocket for the purpose of &quot;measuring two-phase flow behavior and heat transfer characteristics during coasting flight.&quot; These devices withstood the vibration shock test of 40G and succeeded to measure the void fraction of liquid/gas nitrogen two phase flow under vacuumed and microgravity circumstance. This report explains development and experiment results of the void fraction sensor and a capacitance amplifier. &lt;/p&gt;

DOI： 10.2322/jjsass.66.147

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Authorship：Lead author   Language：English   Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics Inc.  

Nanosecond-pulsed surface dielectric barrier discharge plasma actuators are used to control shock-wave/ boundary-layer interactions. Characterization of the separation region without plasma actuation is conducted to understand the interaction based on measurements of pressure distribution, schlieren imaging, and velocimetry by the femtosecond laser electronic excitation tagging technique. The results show a weak separation due to the interaction. Three types of plasma actuators are applied to control the separation. Schlieren images are taken by a high-speed camera to evaluate the magnitude of the interaction. It is shown that the plasma actuators affect the flow in two different ways: heat generation in the boundary layer, and generation of vorticity near the surface.When the first effect is dominant, the shock-wave/boundary-layer interaction becomes stronger and the size of the separation bubble increases. If the vorticity generation prevails, it suppresses the separation due to the momentum transfer from the main flow to the boundary layer. The experimental results of the three actuator's geometries provide the design guidelines for nanosecond-pulse-driven electrodes to control the interaction.An optimal actuation frequency is found through an investigation of the frequency response of the flow.

DOI： 10.2514/1.B36530

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Language：Japanese   Publishing type：Research paper (scientific journal)   Publisher：プラズマ応用科学会  

A 200 W class Hall thruster has been developed for the small satellites main propulsion. In order to overcome the degradation owing to the miniaturization, it has a unique magnetic field configuration, similar to “magnetic shielding”. The thrust performance was measured using a pendulum thrust stand. The thrust, thrust to power ratio, thrust efficiency is achieved to 13.1 mN,56 mN/kW and 0.36, respectively at discharge voltage of 300 V, xenon mass flow rate of 1.0 mg/s. The obtained thrust performance is comparable to the other 200 W class thruster.

DOI： 10.34377/aps.25.2_85

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Language：Japanese   Publishing type：Research paper (scientific journal)   Publisher：一般社団法人 日本航空宇宙学会  

&lt;p&gt;全電化静止衛星では推進系としてイオンエンジン，ホールスラスタ等の電気推進のみを搭載し，軌道保持のみならず軌道上昇も電気推進にて行う．これにより従来の化学推進系に比して，長い遷移期間を要するものの大幅に搭載推薬量を削減することができる．2015年，米にて世界初の全電化静止衛星が打上げられ，イオンエンジンによる約半年間の軌道上昇後，静止軌道に到達した．これに追従すべく，各国で全電化静止衛星およびそれに搭載される電気推進の研究開発が活発化している．ホールスラスタはその高推力電力比により3～4カ月程度の短期間遷移を実現でき，さらに低コスト，ロバスト等の特徴も有し，全電化衛星向け電気推進の最有力候補とされる．本稿では全電化静止衛星の世界動向を俯瞰するとともに，全電化衛星への搭載が計画されているホールスラスタの開発動向についても概説する．&lt;/p&gt;

DOI： 10.14822/kjsass.65.9_274

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Authorship：Lead author   Language：English   Publishing type：Research paper (international conference proceedings)   Publisher：AMER INST AERONAUTICS ASTRONAUTICS  

During rocket flights, ionized exhaust plumes from solid rocket motors may interfere with radio frequency transmission under certain conditions. A computational fluid dynamics and finite difference time-domain method coupling approach was established for predicting interference and radio frequency attenuation levels during an actual rocket flight. The detailed plasma flowfield and radio frequency transmission characteristics were revealed in the calculations. The calculated far-field received levels were compared with the in-flight attenuation data at different look angles (angles between the vehicle axis and the line of sight of the antennas), and the calculated results showed good agreement with the flight data over a wide range of look angles. An adaptation of the model, based on the diffraction theory, proved appropriate both for rough estimation of attenuation and for conducting a preliminary analysis of signal/rocket plume interactions.

DOI： 10.2514/1.A32957

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Language：Japanese   Publishing type：Research paper (scientific journal)   Publisher：THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES  

The Japan Aerospace Exploration Agency launched the S-310-43 sounding rocket from the Uchinoura Space Center on Aug.04, 2014 for the purpose of investigating such behavior as boiling and flow of cryogenic liquid rocket propellant in an environment simulating coasting flight on orbit by using the sounding rocket's sub-orbital ballistic flight. In the low-gravity state, the cryogenic fluid (liquid nitrogen) was introduced into the test sections of similar shapes to the flow channels in the cryogenic propulsion systems. The boiling of liquid nitrogen inside the test-sections and the transition of flow regimes from gas/liquid two-phase flow to liquid mono-phase flow were visualized. The temperatures, pressures and void fractions of each channels were measured as well. Development of the experimental equipment for S-310-43 sounding rocket is described in this paper.

DOI： 10.2322/jjsass.63.188

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Next generation arcjets should have light-weight design and prolonged lifetime. For the former topic, it is shown that the radiator mass can be drastically reduced by the effective use of propellant as a coolant at the lower temperature region on the radiator. Resulting thruster weight of 2.0 kg including the radiator is possible for 15 kWe arcjet. For the latter topic, replaceable cathode system is proposed and some key issues are mentioned.

DOI： 10.2322/tastj.12.to_1_1

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DOI： 10.2514/1.A32223

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Authorship：Lead author   Language：Japanese   Publishing type：Research paper (scientific journal)   Publisher：一般社団法人 日本航空宇宙学会  

For future space transportation system development, efficient liquid propellant acquisition technologies under microgravity could be required to realize long-term missions in orbit. Microgravity environment is generally established through drop tower, parabolic flight by airplane or orbital/suborbital experiment. These methods are large-scale or not flexible so that compact and simple method is needed for the efficient development. To respond to such request, we propose a static experiment on the ground to realize the similar static free surface under microgravity. The ground experimental result was compared with the results of two types of two-phase flow simulation codes for the verification of the methodology and discuss the characteristics of these numerical codes. We also simulated the free surface under the actual flight condition by using the same simulation method and result showed the validity of the experimental method.

DOI： 10.2322/astj.12.73

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Authorship：Lead author   Language：English   Publishing type：Research paper (scientific journal)   Publisher：AMER INST PHYSICS  

Under certain conditions during rocket flights, ionized exhaust plumes from solid rocket motors may interfere with radio frequency transmissions. To understand the relevant physical processes involved in this phenomenon and establish a prediction process for in-flight attenuation levels, we attempted to measure microwave attenuation caused by rocket exhaust plumes in a sea-level static firing test for a full-scale solid propellant rocket motor. The microwave attenuation level was calculated by a coupling simulation of the inviscid-frozen-flow computational fluid dynamics of an exhaust plume and detailed analysis of microwave transmissions by applying a frequency-dependent finite-difference time-domain method with the Drude dispersion model. The calculated microwave attenuation level agreed well with the experimental results, except in the case of interference downstream the Mach disk in the exhaust plume. It was concluded that the coupling estimation method based on the physics of the frozen plasma flow with Drude dispersion would be suitable for actual flight conditions, although the mixing and afterburning in the plume should be considered depending on the flow condition. © 2012 American Institute of Physics.

DOI： 10.1063/1.4762857

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Language：English   Publishing type：Research paper (scientific journal)   Publisher：TAYLOR & FRANCIS LTD  

This study focuses on understanding and modeling the physical phenomena that occur in degraded zones of silica-phenolic (SiFRP) materials under exposure to high-temperature gasses when applied to a liquid rocket engine (LRE) combustor. Although understanding and modeling these phenomena is considered essential in designing an LRE combustor, few studies on these fields can be found in the available literature. Basically, it is well known that when ablators are heated, a pyrolysis reaction proceeds in them, forming three distinct zones: a charred, a decomposed, and a virgin zone. The obtainable information for the thermal response of SiFRP in ground-firing tests is classified in two categories. The first category involves the equilibrium state characteristics after a long time has elapsed following burnout. This refers to the degraded thickness distribution, which reflects 3D information (the combustor's inner surface x the thickness direction) regarding the heat load distribution over the entire combustor's inner surface, owing to the highly insulating nature of SiFRP. The second category involves the transient characteristics of the propagation of the degraded zones in SiFRP, which can be detected using an ultrasonic testing (UT) method. In this paper, the progress of in-depth phenomena of SiFRP and their physical variations were intentionally studied. Our aim was to clarify and specify the quantitative threshold values of the interface points that characterize each degraded zone and the UT reflection point, and then express these values in terms of physical quantities that could appear in a numerical analysis. © 2012 Japan Society for Composite Materials, Korean Society for Composite Materials and Taylor & Francis.

DOI： 10.1080/09243046.2012.743715

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Language：Japanese   Publishing type：Research paper (scientific journal)   Publisher：The Japan Society for Composite Materials  

This study focuses on the understanding and modeling of the physical phenomena occurring in the degraded zones of Silica Phenolic (hereafter referred as SiFRP) under exposure to high temperature gases when applied to liquid rocket engine (LRE) combustor. Although the understanding and modeling of the phenomena is supposed to be essential in designing LRE combustor, a few works done in this fields appear in the open literatures. Basically, it is well known that when heated, the pyrolysis reaction proceeds in the SiFRP, forming 3 distinct zones of charred, decomposed and virgin zone, respectively. The obtainable information for the thermal response of SiFRP at ground firing tests is classified in 2 categories. The first is the equilibrium state characteristics after long time elapsed from the burnout, namely, the degraded thickness distribution, which reflects 3-D information (combustor inner surface×thickness-direction) of heat load distribution over the entire combustor inner surface thanks to the highly insulating nature of SiFRP. The second is the transient characteristics with regard to the degraded zones propagation in the SiFRP, which can be detected by application of ultrasonic testing (UT) method. In this paper, the progress of in-depth phenomena of SiFRP and the physical variation were intentionally studied. We strive to clarify and specify the quantitative threshold values of the interface points which characterizes each degraded zone and UT reflection point and eventually express the threshold values in terms of physical quantities that could appear in numerical analysis.

DOI： 10.6089/jscm.38.228

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Language：Japanese   Publishing type：Research paper (scientific journal)   Publisher：The Japan Society for Composite Materials  

This study focuses on the understanding and modeling of the physical phenomena occurring in the heated surface of Silica Phenolic (hereafter referred as SiFRP) under exposure to high temperature gases when applied to liquid rocket engine (LRE) combustor. Although the understanding and modeling of the phenomena is supposed to be essential in designing LRE combustor, the works done in this fields are seldom seen in open literatures. When we look at the inner surface of SiFRP combustor of LRE after ground firing tests, the signs of white streak, melting and recession are sometimes observed. In this research, the experimental characterization activities on the SiFRP surface phenomena are conducted, based on the various optical measurements (digital camera, SEM (scanning electron microscope), XRF (X-ray fluorescence), spectral reflectivity (wavelength: 400-700 nm)) for the heated samples, namely, the ones from arcjet heating tests under Air or N₂ stream conditions or the ones from SiFRP combustor after LRE.

DOI： 10.6089/jscm.38.200

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Authorship：Lead author   Language：English   Publishing type：Research paper (scientific journal)   Publisher：The Japan Society of Mechanical Engineers  

Ultimate weight reduction and low manufacturing cost are strongly required in the development of next-generation launch vehicles. It is one of effective methods for these purposes to change structural materials from conventional aluminum alloys to composites, such as CFRP (Carbon Fiber Reinforced Plastics). Under these situations, we developed a launch vehicle structure by using CFRP/honeycomb sandwich panels in order to achieve both weight and cost reduction. FEM (Finite Element Method) simulation was utilized to effectively carry out the structural design. Metallic structures are applied for the flanges (joining interfaces between the other structures) in general; however, we used CFRP structures also for the interface flanges for the weight reduction and manufacturing simplicity. The designed CFRP structure was fabricated and a static load test was conducted considering flight loads during actual launch. The strain distribution was obtained by FBG (Fiber Bragg Grating) sensors and three-dimensional displacement was measured by laser tracker as well as normal strain gauges in the limit load test. These measured data were compared with the FEM simulation results and they show good agreement. We will apply these manufacturing, testing, simulation and sensing techniques for the future launch vehicle developments. © 2011 The Japan Society of Mechanical Engineers.

DOI： 10.1299/kikaia.77.81

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Authorship：Lead author   Language：English   Publishing type：Research paper (scientific journal)   Publisher：IEEE-INST ELECTRICAL ELECTRONICS ENGINEERS INC  

The ionized exhaust plumes of solid rocket motors may interfere with RF transmission under certain flight conditions. To understand the important physical processes involved, we measured microwave attenuation and phase delay due to the exhaust plume during sea-level static firing tests for a full-scale solid propellant rocket motor. The measured data were compared with the results of a detailed simulation performed using the frequency-dependent finite-difference time-domain ((FD) TD) method. The numerically derived microwave attenuation was in good agreement with experimental data. The results revealed that either the line-of-sight microwave transmission through ionized plumes or the diffracted path around the exhaust plume mainly affects the received RF level, which depends on the magnitude of the plasma-RF interaction. © 2010 IEEE. 2

DOI： 10.1109/TAP.2010.2055796

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Authorship：Lead author   Language：English   Publishing type：Research paper (scientific journal)   Publisher：AMER INST AERONAUT ASTRONAUT  

The draining process for the circular-plate-type device, using both experiments and numerical simulations was investigated. The effect of turbulence was evaluated by conducting both turbulent and laminar flow calculations. The Navier-Stokes equation was spatially discretized by the finite-volume method, and the convective flux was evaluated by a second-order upwind scheme. The air and water gas-liquid two-phase flow was modeled by the volume-of-fluid (VOF) method, and the surface tension of water was assumed to have a 75° contact angle to the wall. The cell distribution of the normal 2-D grid and radial cross section of the 3-D grid were similar and the minimum grid size was 0.125 mm around the baffle edge and the maximum was approximately 5 mm at the tank equator for both grids. A numerical accuracy study based on Richardson statistics indicated that the fractional errors of the liquid remaining were roughly 19 and 1.2% for the normal and finer-grid systems, respectively.

DOI： 10.2514/1.48398

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Authorship：Lead author   Language：English   Publishing type：Research paper (scientific journal)   Publisher：AMER INST AERONAUT ASTRONAUT  

Under certain conditions during rocket flight, ionized exhaust plumes from solid rocket motors may interfere with RF transmission. To understand the relevant physical processes involved in this phenomenon, measurement of microwave attenuation and phase delay caused by rocket exhaust plumes was attempted in a sea-level static firing test for a full-scale solid propellant rocket motor. The measured data were analyzed by comparing them with simulation results for an exhaust plume flowfield. The results revealed that the change in the shock structure in the plume affects the microwave attenuation level, since it significantly affects the plasma density at the measuring location. When the plasma density in the plume is low, the microwaves can penetrate the plume. The plume plasma properties were successfully estimated for that situation in which the numerically calculated attenuation level agreed well with the experimental results. On the other hand, high-density plasma in the plume does not allow penetration. Therefore, microwaves bypass around the plume and the diffraction effect becomes dominant. Copyright © 2010 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

DOI： 10.2514/1.48173

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Authorship：Lead author   Language：English   Publishing type：Research paper (scientific journal)   Publisher：AMER INST AERONAUT ASTRONAUT  

Experimental velocimetry in the discharge chamber of a two-dimensional magnetoplasmadynamic (MPD) arcjet, fabricated for experimental internal flow measurement, was conducted to investigate the acceleration process for hydrogen propellant. In the experiment, we evaluated the neutral atom velocity and the temperature from the laser absorption spectroscopy using a tunable diode laser. The results using two types of anode, a flared-tvpe anode and a converging-diverging (C-D)-type anode, were compared for the case with a discharge current of 13 kA and a mass-flow rate of 0.65 g/s. It was found that a large velocity slip between the ions and the neutrals prevented the acceleration of the neutral particles. This velocity slip is expected to reduce thrust performance because the flow with ion-neutral slip requires additional electric power compared to the flow without velocity slip. The velocity slip was reduced in the case of the C-D anode compared to the flared anode because of strong ion-neutral momentum coupling in the throat region of the C-D anode. Copyright © 2006 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

DOI： 10.2514/1.17038

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Authorship：Lead author   Language：Japanese   Publishing type：Research paper (scientific journal)   Publisher：THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES  

Velocity and temperature measurements were conducted for a two-dimensional magnetoplasmadynamic arcjet with hydrogen propellant. To obtain the velocities of both atoms and ions, laser absorption spectroscopy was employed for atom, and time-of-flight technique was used for ions. In a quasi-steady operation at 13kA/0.65g/s, larger ions velocity (33km/s) than that of the atoms (13km/s) was found in the case of flared anode configuration, which implies that large mean free path between the ions and atoms prohibited momentum transfer from the ions to the neutral particles. This velocity differen...

DOI： 10.2322/jjsass.53.215

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Language：Japanese   Publisher：宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)  

令和3年度宇宙輸送シンポジウム（2022年1月13日-14日. オンライン開催)
Space Transportation Symposium FY2021 (January 13-14, 2022. Online Meeting)
資料番号: SA6000173030
STCP-2021-031

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Language：Japanese   Publisher：宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)  

令和3年度宇宙輸送シンポジウム（2022年1月13日-14日. オンライン開催)
Space Transportation Symposium FY2021 (January 13-14, 2022. Online Meeting)
著者人数: 12名
資料番号: SA6000173016
STCP-2021-016

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Language：Japanese   Publisher：宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)  

令和3年度宇宙輸送シンポジウム（2022年1月13日-14日. オンライン開催)
Space Transportation Symposium FY2021 (January 13-14, 2022. Online Meeting)
著者人数: 16名
資料番号: SA6000173015
STCP-2021-015

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Language：Japanese   Publisher：Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

Space Transportation Symposium FY2020 (January 14-15, 2021. Online Meeting)

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Language：Japanese   Publisher：Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

Space Transportation Symposium FY2020 (January 14-15, 2021. Online Meeting)

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Language：Japanese   Publisher：Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

Space Transportation Symposium FY2020 (January 14-15, 2021. Online Meeting)

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Language：Japanese   Publisher：Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

Space Transportation Symposium FY2020 (January 14-15, 2021. Online Meeting)

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Magnetoplasmadynamic thrusters are a form of electric propulsion for space applications that use magnetic and electric fields to accelerate plasma from a spacecraft to generate thrust. It has been shown experimentally and theoretically that applying a strong magnetic field to an MPD thruster can improve thrust and efficiency and lower the required discharge current. This work presents concept design and modelling of a 1 T high-Temperature superconducting applied field module cooled by a miniaturised cryocooler targeting an existing thruster. Using a 3D finite element modelling approach, thermal and electromagnetic predictions of the mechanical assembly are performed, which include temperature dependent thermal properties of the mechanical components and temperature and field dependent critical current anisotropy of the superconductor. The model was used to generate design curves to determine the operational temperature required to achieve central fields up to 2.5 T.

DOI： 10.1109/TENCON54134.2021.9707308

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The Japan Aerospace Exploration Agency, in partnership with academia and industry, are developing the Air Turbo Rocket for Innovative Unmanned Mission (ATRIUM) engine: an air turboramjet + rocket combine cycle propulsion system intended to replace conventional liquid rocket engines in Vertical Takeoff Vertical Landing applications, such as reusable sounding rockets. A subscale Flight Test Bed (FTB) vehicle is also being developed to demonstrate the ATRIUM engine in a flight environment. In this paper, the ATRIUM engine and FTB vehicle are introduced, and current progress in their development is summarized. Future test plans and practical applications are also discussed.

DOI： 10.2514/6.2021-4197

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Language：Japanese   Publisher：宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)  

令和元年度宇宙輸送シンポジウム（2020年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県
Space Transportation Symposium FY2019 (January 16-17, 2020. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan
資料番号: SA6000147068
レポート番号: STEP-2019-019

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Language：Japanese   Publisher：宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)  

令和元年度宇宙輸送シンポジウム（2020年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県
Space Transportation Symposium FY2019 (January 16-17, 2020. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan
資料番号: SA6000147012
レポート番号: STCP-2019-012

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Language：Japanese   Publisher：The Japan Society of Mechanical Engineers  

DOI： 10.1299/jsmemecj.2020.J19101

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Future NASA architectures have baselined cryogenic propulsion systems as well as cryogenic fluid management to support lunar missions and ultimately to support future missions to Mars. These missions will require chilling hardware down prior to engine restart as well as chilling lines and tanks prior to transferring and refueling these propulsion elements in orbit. In lieu of expensive tests conducted on-orbit, accurate predictive computational models of these chilldown processes can be used to reduce system and propellant mass as well as mission risk. To gain confidence in these computational models, appropriate anchoring and validation to experimental data in a relevant environment needs to be performed. Recent ground and sub-orbital flight experiments conducted by the Japan Aerospace Exploration Agency (JAXA) investigated chilldown of a complex channel resembling a turbopump bearing cavity at low flow rates. This work presents Computational Fluid Dynamics (CFD) model development of the chilldown experiment employing two-phase flow boiling models available in commercial CFD software STAR-CCM+ using Volume of Fluid (VOF) and the traditional Euler-Euler multiphase flow solvers. Comparisons of the numerical and experimental results under normal and low-gravity conditions are presented. An assessment of solid wall temperatures and phase distribution yielded important insights into multiphase solver choice, dependence on gravity environment, and challenges associated with cryogenic flow boiling prediction and validation.

DOI： 10.2514/6.2020-3818

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This paper presents the development of a two-phase CFD model that is used to study sloshing of a cryogenic fluid (LN2) undergoing phase change generated by lateral acceleration and its effects on the ensuing heat and mass transfer and pressure drop in a small-scale tank. In this model the interface is captured using the volume-of-fluid (VOF) method [1]. Kinetic-theory based Schrage relation [2] is used for calculating phase change mass transfer at the liquid-vapor interface. Computational results are compared to the data provided by a non-isothermal sloshing experiment with phase change conducted by the University of Tokyo and JAXA in 2018 using liquid Nitrogen in a transparent tank. The effect of different thicknesses of a liquid temperature stratification layer created at the interface on the pressure drop during sloshing was studied experimentally. CFD results for interface movement and tank pressure are presented and compared in this paper to experimental data for the cases with the smallest and largest liquid temperature stratification thickness.

DOI： 10.2514/6.2020-3794

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The efficient storage of cryogenic propellants is among the key technologies for long-duration space exploration missions. For the orbital transfer vehicle, the required thermal insulation performance is more than ten times higher than that of conventional spray-on foam insulation. Conventional multi-layer insulation blankets are used as excellent insulation for spacecraft in the vacuum environment, but are not usable in the atmospheric environment. A new type of insulation-A load-bearing, non-interlayer-contact spacer MLI (LB-NICS MLI)-has been developed. The insulation performance in both air and vacuum environments is measured with a boil-off calorimeter. According to the test results, LB-NICS MLI is about 3 times superior under 1 atmosphere and 13 times superior under a vacuum, compared to foam insulation. &copy; Published under licence by IOP Publishing Ltd.

DOI： 10.1088/1757-899X/502/1/012062

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Language：Japanese   Publisher：Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

Space Transportation Symposium FY2018 (January 17-18, 2019. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan

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A 3D printed Inconel resistojet is proposed as an option for short time and high fuel efficiency orbit transfers. The current thruster is presented as a proof-of-concept for high performance high temperature variants. Experiments on N2 propellant have been conducted, and the measured performance parameters are presented. Finally, the extra application of the 3D printed resistojet as part of a hybrid electro-chemical thruster is presented.

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In order to improve the cryogenic propellant management technologies for a liquid hydrogen rocket with high specific impulse, JAXA, the University of Tokyo, and the NASA Glenn Research Center have jointly organized a multi-agency model validation collaboration project. As part of this project, JAXA's boiling simulation was validated with NASA's experimental data on vertical pipeline chill-down. Simulation results were in good agreement with the experimental data obtained using an improved boiling model to reproduce the spray flow. This activity achieved liquid hydrogen turbo-pump simulation at JAXA for grasping the boiling flow phenomenon from engine cut-off to re-ignition. This joint research resulted in an international cooperative relationship for discussing the cryogenic propellant management technologies necessary to develop next-generation liquid rockets.

DOI： 10.2514/6.2019-4439

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A 3D-printed resistojet, wherein the heater and flow path were united, was produced. The heater consists of multi-layer shells and is difficult to break down. In this paper, the thermal design and the results of the thrust measurement are reported. The measured electrical resistance matched with the predicted one. Both the heater and the thrust reached 70 percent efficiency when paired with a nitrogen propellant at a mass flow rate of 0.2 g/s.

DOI： 10.2514/6.2018-4907

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Japan Aerospace Exploration Agency, JAXA, has been developing an all-electric geostationary satellite named Engineering Test Satellite 9 (ETS-9). It has domestic 6-kW class Hall thruster for both orbit raising and station keeping. Both the 6 kW thruster input power and xenon mass flow rate, 20 mg/s, are the largest ever in Japanese electric propulsion development, therefore, a new test facility with high pumping speed and appropriate cooling system is required. The construction of the new test facility, HTDT (Hall Thruster Development Test) facility, has been completed in Nov 2017. The 3 m in diameter and 8 m long vacuum chamber with 21 cryopumps has successfully achieved 220 kL/s pumping speed with xenon. To remove the 6 kW heat load from the Hall thruster, the facility uses a closed-loop water cooling system and no liquid nitrogen and Freon, hence, the operation cost can be reduced compared with conventional facilities for high power electric propulsion. The requirement to test the 6-kW Hall thruster has been achieved through a 30-hour long qualification test. The facility effect – differences in thruster performance and discharge current oscillation characteristics between in ground facilities and in orbit – is regarded as an important risk in the Hall thruster development. To characterize the facility effect, the 6-kW Hall thruster has been tested in both HTDT facility and JAXA’s another test facility. Preliminary results of the facility effect investigation, comparison of thruster performance and current oscillation against background pressure, are discussed.

DOI： 10.2514/6.2018-4510

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Language：Japanese   Publisher：Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

Space Transportation Symposium FY2016 (January 19-20, 2017. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan

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Language：Japanese   Publisher：Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

Space Transportation Symposium FY2016 (January 19-20, 2017. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan

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Publisher：53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017  

DOI： 10.2514/6.2017-4761

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Language：Japanese   Publisher：日本航空宇宙学会  

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Language：Japanese   Publisher：一般社団法人 日本機械学会  

&lt;p&gt;The payload capacity of launch vehicles must be increased in order to allow the exploration and development of space to be extended from low-Earth orbit into the solar system. A propellant system using a cryogenic fluid must reduce the amount of unusable propellant due to evaporation and boiling. However, in space exploration and development, where safety and reliability of missions are critical, predictions of the boiling heat transfer of current technology are not sufficiently reliable for thermal management design due to a lack of knowledge and relevant research. Therefore, the objective of this research is to understand and accurately predict boiling heat transfer by developing numerical simulation tools for two-phase flows that consider phase change. In this paper, recent research activity toward the development of chill-down process simulation technology is presented.&lt;/p&gt;

DOI： 10.1299/jsmeted.2016.c113

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In this study, nanosecond pulse surface dielectric barrier discharge plasma actuators are used to control shock wave boundary layer interaction in Mach 2-8 supersonic air flow. An oblique shock wave is generated by a 14 degree wedge-shaped shock generator, and interacts with the boundary layer. First, characterization of the separation region without plasma actuation is conducted to understand the interaction region based on measurement of pressure distribution, schlieren imaging and velocimetry by the femtosecond laser electronic excitation tagging technique. The results show a weak separation due to the interaction occurs in this configuration. Then, three types of plasma actuator are fabricated and applied to control the separation. Schlieren images are taken with several us camera speeds and the size of interaction region or location of reflected shock wave are measured to evaluate the magnitude of the interaction. It was shown that the nanosecond surface dielectric barrier discharge plasma actuators work in two different ways: the heat generation in the boundary layer, and generation of the vorticity near the surface. In the first case the shock wave - boundary layer interaction becomes stronger and the size of separation bubble increases. In the second case the vorticity production successfully suppresses the boundary layer separation due to momentum transfer from the main flow to the boundary layer. The experimental results of the three actuator configurations provide design guidelines for nanosecond pulse driven electrodes to control the shock wave - boundary layer interaction. An optimal nanosecond pulse frequency was found through an investigation of the frequency response of the flow.

DOI： 10.2514/6.2016-5070

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Language：Japanese   Publisher：日本航空宇宙学会  

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Language：Japanese   Publisher：宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)  

平成26年度宇宙輸送シンポジウム（2015年1月15日-16日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000036020レポート番号: STCP-2014-020

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Publisher：51st AIAA/SAE/ASEE Joint Propulsion Conference  

DOI： 10.2514/6.2015-4213

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Language：Japanese   Publisher：Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

Space Transportation Symposium FY2014 (January 15-16, 2015. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan

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The payload capacity of launch vehicles must be increased in order to allow the exploration and development of space to be extended from low-Earth orbit into the solar system. A propellant system using a cryogenic fluid such as liquid oxygen or liquid hydrogen must reduce the amount of unusable propellant due to evaporation and boiling. However, in the space exploration and development where safety and reliability of missions are critical, predictions of the boiling heat transfer of the present technology are not sufficiently reliable for thermal management design due to a lack of knowledge and relevant research. Therefore, the objective of this research is to understand and accurately predict boiling heat transfer by developing numerical simulation tool for two-phase flows that consider phase change. In this paper, some recent research activities toward the development of chill-down process simulation technology are presented.

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Language：Japanese   Publisher：宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)  

平成25年度宇宙輸送シンポジウム（2014年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000016113レポート番号: STEP-2013-040

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Language：English   Publisher：American Institute of Aeronautics and Astronautics Inc.  

High power electric propulsion system is strongly required for future orbital space transportation. MPD (Magneto-Plasma-Dynamic) thrusters and DC (Direct Current) arcjets with hydrogen as a propellant are promising candidates for the missions because of their high performance and adaptability to high power operation. However, to use hydrogen for long term orbital missions, its storage in orbit is crucial issue to be considered. Firstly, we proposed a hydrogen storage and feed system for electric thrusters by applying our technologies derived from the liquid hydrogen launch vehicles. Secondly, we present R&D activities of hydrogen MPD thruster and DC arcjet, especially focusing on the improvement of their performance and durability. Then, development strategy of hydrogen electric thrusters is also discussed. Finally, advantages of hydrogen electric thruster were shown compared with conventional xenon thrusters through mission analyses of lunar orbit insertion and GTO-GEO transportation.

DOI： 10.2514/6.2014-3507

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Language：Japanese   Publisher：宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)  

平成25年度宇宙輸送シンポジウム（2014年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000016099レポート番号: STEP-2013-026

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Language：Japanese   Publisher：宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)  

平成25年度宇宙輸送シンポジウム（2014年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000016097レポート番号: STEP-2013-024

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Publisher：49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference  

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Language：Japanese   Publisher：宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)  

平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県形態: カラー図版あり形態: PDF資料番号: AA0061856104レポート番号: STEP-2012-021

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Language：Japanese   Publisher：宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)  

平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県形態: カラー図版あり形態: PDF著者人数: 12名資料番号: AA0061856048レポート番号: STCP-2012-048

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Language：Japanese   Publisher：宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)  

平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県形態: カラー図版あり形態: PDF資料番号: AA0061856113レポート番号: STEP-2012-030

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Publisher：The Japan Society of Mechanical Engineers  

For the prediction of heat transfer coupled with sloshing phenomena in the propellant tanks of launch vehicle, the pressure drop induced by heat transfer and the dynamic motion of cryogenic liquid in sub-scale vessels were experimentally observed. The correlation between the pressure drop and liquid motion was confirmed in the experiment. Results of the test suggests that pressure is sensitive when gaseous species are same as liquid. In addition, even if a gaseous temperature is different (about 3 OK), the amount of pressure drop does not change.

DOI： 10.1299/jsmemecj.2013._G051014-1

Language：Japanese   Publisher：49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference  

During rocket flights, ionized exhaust plumes from solid rocket motors may interfere with radio frequency (RF) transmission under certain conditions. To clarify the physical process involved and to establish the estimation methodology, a plume–RF interference experiment during a sea-level static firing test of a full-scale solid rocket motor was conducted. The result of the ground experiment was adequately matched by a computational fluid dynamics (CFD) model of the plume flow field coupled to a finite-difference time-domain (FDTD) model of RF transmission. The CFD/FDTD coupling method was then refined for predicting interference and RF attenuation levels during an actual rocket flight. The calculated far-field received levels were compared with the in-flight attenuation data at different look angles (angles between the vehicle axis and the line-of-sight of the antennas). The calculated results showed good agreement with the flight data over a wide range of look angles. An adaptation of the model, based on the diffraction theory, proved appropriate both for rough estimation of attenuation and for conducting a preliminary analysis of signal/rocket plume interactions.

DOI： 10.2514/6.2013-3790

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For the prediction of heat transfer coupled with sloshing phenomena in the propellant tanks of reusable launch vehicle, the pressure drop induced by heat transfer and the dynamic motion of liquid in sub-scale vessels were experimentally observed and numerically investigated. The correlation between the pressure drop and liquid motion was confirmed in the experiment. The mechanisms enhancing heat transfer were discussed based on the computation. It was suggested that splash and wavy surface induced by violent motion of liquid cause the pressure drop in the closed vessel. In addition, as the preliminary investigation, non-isothermal sloshing of liquid nitrogen and liquid hydrogen were successfully visualized and pressure drop depending on the gaseous species was discussed. © 2011 by Takehiro Himeno.

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In rocket flights, ionized exhaust plumes from solid rocket motors may interfere with RF transmission under some conditions. In order to clarify the important physical process involved, microwave attenuation and phase delay due to rocket exhaust plumes were measured during sea-level static firing tests conducted on two types of full-scale solid propellant rocket motors. The measured data were analyzed by comparing them with numerical results such as flowfield simulations of exhaust plumes and by employing a detailed analysis of microwave transmission by using a frequency-dependent finite-difference time-domain (FD2TD) method. The results revealed that either the line-of-sight microwave transmission through ionized plumes or the diffracted path around the exhaust plume mainly affects the received RF level, which depends on the magnitude of the plasma RF interaction. For the actual launch vehicle flight, the transmission process is dominated by the diffraction effect so that we applied a two-dimensional diffraction theory to analyze the communication between a vehicle and a ground station. The attenuation levels estimated using diffraction theory agree with the data recorded in-flight. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

DOI： 10.2514/6.2009-1386

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Applicant：独立行政法人 宇宙航空研究開発機構, 国立大学法人東北大学

Application no：特願2013-111961  Date applied：2013.5

Announcement no：特開2014-232007  Date announced：2014.12

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