研究者詳細 - 杵淵　紀世志

論文 - 杵淵　紀世志

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Vapor-film removal via liquid-jet impingement in small channels for improved cooling performance 査読有り Open Access

Kondo S., Beppu R., Yoneda K., Umemura Y., Kobayashi H., Kinefuchi K.

Applied Thermal Engineering 270 巻頁： 126218 - 126218 2025年7月

担当区分：最終著者,　責任著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Applied Thermal Engineering

High-temperature structures in aerospace and industrial applications require efficient cooling systems. While increasing internal wetted area through small channels, such as those in porous materials, enhances cooling performance, vapor film formation during liquid coolant boiling impedes heat transfer. This study proposes a novel cooling system incorporating liquid-jet nozzles into small channels to remove vapor films using jet inertial forces and enhance heat transfer efficiency. Experiments were conducted in two phases. Initial visualization studies using colored water and nitrogen gas without phase change enabled quantitative evaluation of liquid contact ratio on the impingement wall. An indicator, the Weber number, was newly proposed for the problem, revealing a minimum Weber number of 0.1 for effective gas-film removal. Subsequent experiments with heated walls and low-boiling-point liquids demonstrated that the critical Weber number for film rupture remained consistent regardless of phase change or coolant type, validating the applicability of Weber number as a universal parameter for evaluating film removal feasibility. Analysis of boiling curves revealed that increasing jet flow rates or Weber numbers promoted earlier transitions from film to nucleate boiling even at higher wall superheats. Considering the inherent tradeoff between Weber number and internal wetting surface area, this study establishes crucial design guidelines for optimizing liquid jet cooling systems in high-temperature structures.

DOI： 10.1016/j.applthermaleng.2025.126218

Hall Thruster System Using Dry Ice Propellant at the Triple Point 査読有り

Nosaka, S; Cho, S; Watanabe, H; Gunji, K; Matsui, M; Kinefuchi, K

JOURNAL OF PROPULSION AND POWER 2025年3月

Estimation method for dry ice formation under triple point through tank depressurization 査読有り Open Access

Nosaka S., Cho S., Watanabe H., Kinefuchi K.

Applied Thermal Engineering 262 巻頁： 125270 - 125270 2025年3月

担当区分：最終著者,　責任著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Applied Thermal Engineering

To achieve a carbon–neutral society, efficient CO2 transportation is necessary. Ship transportation is a potential candidate for long-distance transportation, wherein storing CO2 in a high-density, low-pressure state in tanks is desirable. Storing liquid CO2 near the triple point is regarded as an effective approach. However, the solidification of the liquid when the tank pressure reduces for any reason is cited as a risk. Dry ice formation should be prevented because it can hinder CO2 unloading, cause pipeline blockages, and damage equipment. To establish high-performance, efficient CO2 transportation ships that store CO2 at or near the triple point, understanding the phase changes within tanks during depressurization is crucial. This study first derived theoretical equations that can be applied straightforwardly in practical situations, based on mass, volume, and energy conservation laws to calculate the solid and liquid phase changes caused by heat input and CO2 release outside the tank. To verify these equations, experiments were conducted to visualize the solid and liquid phase changes inside a tank at a triple point by venting CO2 gas at a constant flow rate. The experiments were performed in a vacuum chamber under controlled heat input conditions. The validity and effectiveness of the theory were demonstrated by comparing the derived phase-change prediction equations with the experimental results.

DOI： 10.1016/j.applthermaleng.2024.125270

MPD thruster with a strong applied field up to 0.5 T using cryogenic fluid-cooled magnetic coil 査読有り Open Access

Maeshima D., Mori R., Inoue S., Tsuchiya Y., Kinefuchi K.

Journal of Electric Propulsion 4 巻 ( 1 ) 2025年2月

担当区分：最終著者,　責任著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Journal of Electric Propulsion

To achieve an electric thruster with both high efficiency and high thrust density, we propose an applied-field magnetoplasmadynamic thruster that utilizes liquid hydrogen both as a propellant and as a coolant for high-temperature superconducting magnets. In this study, as an initial step toward this goal, we developed an applied filed magnetoplasmadynamic thruster with a strong magnetic field applied by a conventional normal-conducting copper coil cooled with cryogenic liquid nitrogen. This paper primarily focuses on the engineering aspects of integrating a novel thruster system with a cryogenic fluid-cooled coil. Additionally, it provides initial insights into the thruster’s operating modes under a strong magnetic field. The cryogenic coil successfully achieved a magnetic field of 0.5 T by cooling the coil with liquid nitrogen, thanks to the appropriate thermal and sealing designs. The developed cryogenic cooling system can also be applied to high-temperature superconductors. Discharge experiments using argon as a propellant revealed that the thruster operated in an electrothermal mode, exhibiting both high-voltage and low-voltage modes akin to arcjet thrusters. The specific impulse in this mode was limited to around 340 s. However, when the cathode position was shifted downstream, a mode with strong blue emission along the central axis, resembling electromagnetic acceleration, was occasionally observed. The current-voltage characteristics also appeared to indicate characteristics of electromagnetic acceleration. Further investigation into the transition to electromagnetic acceleration mode is necessary for future work.

DOI： 10.1007/s44205-025-00110-8

その他リンク： https://link.springer.com/article/10.1007/s44205-025-00110-8/fulltext.html

Visualization of suppressed shock wave/turbulent boundary layer interaction using cryogenic wall cooling 査読有り

Miki Y., Ando L., Acuña M., Kinefuchi K.

Physics of Fluids 37 巻 ( 1 ) 2025年1月

記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Physics of Fluids

To investigate the wall-cooling effects on shock wave/turbulent boundary layer interaction (SWTBLI) with limited experimental data, a supersonic wind tunnel wall was cooled using liquid nitrogen as the cryogenic coolant. Under the condition of the mainstream Mach number, in the range of 2.02-2.04, the wall temperature was cooled to 88-92 K, corresponding to a wall-to-recovery temperature ratio of 0.31-0.33. The flow structures with and without wall cooling were observed using the schlieren method. The reflected shock motion or interaction length in the schlieren image suggested that boundary layer separation was suppressed under the cooling condition in relation to that under the non-cooling condition, and the suppressed ratio of the cooling-to-uncooling interaction length was approximately 0.60-0.72. Additionally, while the mainstream state and wall temperature near the separation were constant, a gradual change in the separated flow field was observed under the wall-cooling condition. This was due to the slow wall temperature increase in the upstream wall of the separation region, where the incoming boundary layer developed. Each flow field of SWTBLI in the present experiment, using liquid nitrogen as the cryogenic coolant, was consistent with the classical Chaman's free interaction theory.

DOI： 10.1063/5.0250796

Characterization of heated volume generation by nanosecond pulsed plasma actuator with various pressure environments 査読有り

Matsunaga T., Iwamoto M., Miki Y., Kinefuchi K.

Journal of Physics D: Applied Physics 57 巻 ( 37 ) 2024年9月

担当区分：最終著者,　責任著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Journal of Physics D: Applied Physics

Nanosecond dielectric barrier discharge (NS-DBD) has emerged as a promising technique for controlling high-speed flows, generating a heated volume that generates strong density and viscosity gradients, thereby perturbing flow dynamics. Since its potential application in low-pressure, high-speed flows, understanding how the size and growth of the heated volume correlate with surrounding pressure is crucial. In this study, we employed typical schlieren and background-oriented schlieren (BOS) techniques to investigate the heated volume’s sensitivity to surrounding pressure in quiescent air. The observed heated volume’s size variations with surrounding pressure likely stemmed from the increase in thermal diffusivity at lower pressures. BOS findings unveiled a nearly linear decrease in heated volume’s core density with energy input. Meanwhile, the heated volume’s size augmented with energy input but exhibited gradual saturation, attributable possibly to shear stresses impeding volume expansion as temperature and viscosity rose, or to consumption of energy in vibration excitation and other reactions. In the cases of 100 and 50 kPa, the sensitivity of the heated volume’s size to the reduced electric field appeared to be similar. However, at 10 kPa, where the reduced electric field is higher compared to that of the 100 and 50 kPa cases due to the lower air density, the size sensitivity drastically decreased. This suggested a transition in discharge mode from filamentary to diffusive behavior at lower pressures.

DOI： 10.1088/1361-6463/ad5699

Onboard Cryogenic Liquid-Propellant Subcooler Based on Thermodynamic Vent for Upper-Stage Propulsion System 査読有り

Banno Y., Kinefuchi K.

Journal of Spacecraft and Rockets 61 巻 ( 4 ) 頁： 919 - 929 2024年7月

担当区分：最終著者,　責任著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Journal of Spacecraft and Rockets

Subcooling of cryogenic liquid propellants offers significant advantages for launch vehicles. Liquid subcooling can reduce the extent of cavitation generation in turbopumps, which provides advantages for upper-stage engine reignition in orbit. Additionally, liquid-propellant subcooling densifies the liquid, which can reduce the propellant tank volume and increase launch capability when it is applied during the propellant-filling phase on the ground. Here, a cryogenic propellant subcooling system derived from thermodynamic vent system (TVS) concepts is proposed. In contrast to prior approaches with TVS, the proposed method only focuses on the liquid subcooling and only requires the installation of a heat exchanger at the tank bottom, making it easily applicable to conventional upper stages and enhancing launch capability. First, ground tests were conducted using liquid nitrogen. We introduced energy efficiency to evaluate the subcooler performance, and an optimum point was found for the coolant supply pressure and Joule–Thomson orifice diameter. To investigate the advantages of the proposed method, a theoretical model for the subcooling process before engine reignition in orbit was developed based on experimental observation. The model assuming liquid hydrogen demonstrated an improved weight penalty in the proposed method compared to the conventional vent/pressurization method. A study that quantitatively addresses the enhancement of launch capability through TVS-induced liquid subcooling is unprecedented, paving the way for a new cryogenic propulsion system.

DOI： 10.2514/1.A35888

Background-oriented schlieren and laser Rayleigh scattering complementary method for accurate density field visualization 査読有り

Iwamoto M., Miki Y., Kinefuchi K.

Experiments in Fluids 65 巻 ( 6 ) 2024年6月

担当区分：最終著者,　責任著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Experiments in Fluids

Gas flow visualization is an essential technique for understanding the gas flow characteristics. Various quantitative distribution measurement methods have been proposed, each with its own advantages and disadvantages. For example, the background-oriented schlieren method provides the quantitative density distribution for wide areas with a simple optical setup, but it disadvantageously requires the appropriate boundary conditions need to be set when integrating the Poisson equation. The laser Rayleigh scattering method also provides quantitative density distribution, but it requires a high-power laser for wide-area measurements because laser intensity directly influences measurement accuracy. This study proposes a method that complements the weak points of the above two methods. First, a wide area is measured using the background-oriented schlieren method, and then, the laser Rayleigh scattering method is applied only for the boundary region to obtain the boundary condition. For a heated turbulent air jet with Reynolds number 3000, the results of the proposed method are compared with the numerical analysis and thermocouple temperature measurements. The results well match, indicating the applicability and usefulness of the proposed method. Furthermore, these results contribute to demonstrating the significance of boundary conditions in the background-oriented schlieren method and the establishment of setting guidelines.

DOI： 10.1007/s00348-024-03772-6

Plume mode instability enhanced by emitter surface poisoning in hollow cathode 査読有り Open Access

Suzuki A., Cho S., Watanabe H., Kinefuchi K.

Journal of Applied Physics 135 巻 ( 10 ) 頁： 103301 2024年3月

担当区分：最終著者,　責任著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Journal of Applied Physics

The unstable plume mode of hollow cathodes should be avoided in practical applications because it severely degrades the overall cathode lifetime. In this study, we investigate the spot-plume transition and plasma stability characteristics of an unused segmented lanthanum hexaboride emitter. The expansion of the unstable plume mode region is observed during a discharge experiment. Subsequently, the segmented emitter is retrieved, the inner surface of the emitter is observed, and the work function on the surface is measured at room temperature. The emitter surface exhibits color variations with oxygen and carbon detection. The downstream edge shows the original purple color and almost no degradation in the work function. The high temperature in this region promotes the desorption of carbon and oxygen. In the spot mode, this region mainly contributes to thermionic electron emission; therefore, the discharge voltage in the spot mode does not change during the discharge experiment. Carbon or carbide is detected in the middle of the axial direction on the emitter surface, where the surface temperature is not sufficiently high to desorb carbon during discharge. Based on the surface analysis results, the dominant substance in the region where carbon is detected was lanthanum carbide. An increase in the work function is indicated in the region, which appears to increase the plasma instability. According to previous studies, an increase in the work function results in a rise in the potential in the emitter, and an increase in the electron temperature in the outside plume region induces the plasma instability. Further investigation is needed to understand the mechanism connecting the rise in the work function and the rise in the electron temperature in the plume region.

DOI： 10.1063/5.0188080

Operation of a plasma thruster featuring a 1.1 T high temperature superconducting magnet 査読有り

Chris R. Acheson, Kiyoshi Kinefuchi, Daisuke Ichihara, Daiki Maeshima, Ryoyu Mori, Ryota Nakano, Ryohei Takagi, Konstantinos Bouloukakis, Jakub Glowacki, Max Goddard-Winchester, Nicholas J. Long, Jamal R. Olatunji, Betina Pavri, Randy Pollock, Cameron Shellard, Nick M. Strickland, Stuart C. Wimbush

Journal of Electric Propulsion 3 巻 ( 17 ) 2024年

Performance Evaluation of a Plasma Thruster Using a High-Temperature Superconducting Magnet 査読有り

Kiyoshi Kinefuchi, Stuart Wimbush, Daisuke Ichihara, Chris Acheson, Ryota Nakano, Daiki Maeshima, Ryohei Takagi, Ryoyu Mori, Jamal Olatunji, Max Goddard-Winchester, Randy Pollock, Nick Strickland, Jakub Glowacki, Betina Pavri

Transactions of the JSASS, Aerospace Technology Japan 22 巻 ( AJCPP-2023 ) 頁： aj1 - aj6 2024年

Investigation on the Hybrid NS-DSMC Simulation of a Nozzle Flow Ionization in a Rarefied Atmosphere using a Post-computation Approach 査読有り

Virgile CHARTON, Julien LABAUNE, Kiyoshi KINEFUCHI

Journal of Evolving Space Activities 2 巻 ( 0 ) 頁： 153 2024年

Operational demonstration and experimental characterisation of a central cathode electrostatic thruster equipped with a high temperature superconducting magnet 査読有り

Chris R. Acheson, Jakub Glowacki, Ryota Nakano, Daiki Maeshima, Dominik Saile, Betina Pavri, Ryohei Takagi, Ryoyu Mori, Randy Pollock, Jamal R. Olatunji, Max Goddard-Winchester, Nicholas M. Strickland, Daisuke Ichihara, Stuart C. Wimbush, Kiyoshi Kinefuchi

Journal of Electric Propulsion 2 巻 ( 1 ) 2023年12月

担当区分：最終著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：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

その他リンク： https://link.springer.com/article/10.1007/s44205-023-00060-z/fulltext.html

Energetic ion and plasma oscillation measurements during plume mode operation of a hollow cathode 査読有り

Suzuki A., Kinefuchi K., Ichihara D., Cho S., Watanabe H., Kubota K.

Physics of Plasmas 30 巻 ( 7 ) 2023年7月

担当区分：責任著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Physics of Plasmas

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

Low power arcjet thruster using LaB<inf>6</inf> hollow cathode 査読有り

Takahashi T., Kinefuchi K.

Acta Astronautica 206 巻頁： 89 - 99 2023年5月

担当区分：最終著者,　責任著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Acta Astronautica

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

Dry ice propellant for electric propulsion with triple-point storage 査読有り Open Access

Maki T., Kinefuchi K., Cho S., Watanabe H.

Acta Astronautica 202 巻頁： 283 - 291 2023年1月

担当区分：責任著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Acta Astronautica

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

Comparison of vapor cooling characteristics of a triply periodic minimal surface and other channel geometries 査読有り

Fukuzaki, T; Kinefuchi, K; Umemura, Y; Okita, K; Sakai, H

Mechanical Engineering Journal 10 巻 ( 3 ) 頁： 23-00015 - 23-00015 2023年

基幹ロケット高度化推進系の開発査読有り

杵淵紀世志, 更江渉, 小林弘明, 梅村悠, 杉森大造, 藪崎大輔, 藤田猛, 沖田耕一, 西村真二, 石川佳太郎, 北山治, 姫野武洋, 佐藤哲也

日本航空宇宙学会誌 70 巻 ( 7 ) 頁： 145 - 152 2022年7月

記述言語：日本語掲載種別：研究論文（学術雑誌）出版者・発行元：一般社団法人日本航空宇宙学会

<p>基幹ロケット高度化はH-IIAロケット第2段を改良することにより国際競争力や打上げの柔軟性向上を狙い，2009年度から概念設計を開始し2011年度に開発へ移行した．推進系に対してはロングコーストおよび再々着火実現のための技術開発が課せられ，2013年度に開発を完了した．第2段の性能は輸送能力に直結するため極限の高性能化が求められるが，微小重力や真空といった軌道上環境により地上検証が困難となる．開発では地上要素試験と解析技術を組み合わせ，重要な要素についてはH-IIAロケット上で衛星分離後の機会を利用した軌道上実験を行うなど日本の独自性を生かし産学官共同で開発を進めた．開発の成果は商用通信衛星の受注に繋がり，2機の地球観測衛星の異なる軌道への同時打上げ，惑星間軌道投入にも適用された．獲得した推進系技術はH3ロケットを始め，将来の月探査や再使用ロケットにも貢献する．</p>

DOI： 10.14822/kjsass.70.7_145

Cooling system optimization of cryogenic propellant storage on lunar surface 査読有り Open Access

Kinefuchi K., Miyakita T., Umemura Y., Nakajima J., Koga M.

Cryogenics 124 巻 2022年6月

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Cryogenics

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

Numerical Study of Effect of Pressurant Gas Species on Thermal Behavior in Cryogenic Tank 査読有り

Kinefuchi K., Umemura Y.

Journal of Spacecraft and Rockets 59 巻 ( 4 ) 頁： 1262 - 1275 2022年

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Journal of Spacecraft and Rockets

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

効率的多数回着火を実現する極低温液体ロケットエンジンとサイクル解析査読有り

近藤奨一郎, 杵淵紀世志, Richardson Mathew, 坂本勇樹, 小林弘明

日本航空宇宙学会論文集 70 巻 ( 4 ) 頁： 110 - 118 2022年

Electrostatic-magnetic hybrid ion acceleration for high-thrust-density operation 査読有り

Ichihara D., Nakano R., Nakamura Y., Kinefuchi K., Sasoh A.

Journal of Applied Physics 130 巻 ( 22 ) 2021年12月

記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Journal of Applied Physics

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

Additive-manufactured single-piece thin multi-layer tungsten heater for an electrothermal thruster 査読有り

Kinefuchi K., Nakata D., Coral G., Suyalatu , Sakai H., Tsukizaki R., Nishiyama K.

Review of Scientific Instruments 92 巻 ( 11 ) 頁： 114501 2021年11月

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Review of Scientific Instruments

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

Thrust density enhancement in an electrostatic–magnetic hybrid thruster 査読有り

Ichihara D., Oka K., Higo A., Nakamura Y., Kinefuchi K., Sasoh A.

Journal of Propulsion and Power 37 巻 ( 6 ) 頁： 973 - 976 2021年11月

Thermal Design and Experimental Verification of a Three-Dimensional-Printed Resistojet 査読有り

Nakata Daisuke, Kinefuchi Kiyoshi, Sakai Hitoshi, Suyalatu

JOURNAL OF PROPULSION AND POWER 38 巻 ( 1 ) 頁： 1 - 9 2021年9月

記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Journal of Propulsion and Power

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

Conceptual design study of a vertical takeoff and landing airbreather 査読有り

Kobayashi H., Maru Y., Richardson M.P., Kinefuchi K., Sato T.

Journal of Spacecraft and Rockets 58 巻 ( 5 ) 頁： 1279 - 1292 2021年9月

記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Journal of Spacecraft and Rockets

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

Neutral atom density measurements of xenon plasma inside a μ10 microwave ion thruster using two-photon laser-induced fluorescence spectroscopy 査読有り

Tsukizaki R., Yamashita Y., Kinefuchi K., Nishiyama K.

Vacuum 190 巻 2021年8月

記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Vacuum

This paper reports measurements of the xenon ground state and excited state densities inside a μ10 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 6pʹ [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

Performance of a miniature Hall thruster and an in-house PPU 査読有り

Naoji Yamamoto, Masatoshi CHONO, Ryudo TSUKIZAKI, Takato MORISHITA, Kenichi KUBOTA, Shinatra CHO, Kiyoshi KINEFUCHI, Toru TAKAHASHI

Transactions of the Japan Society for Aeronautical and Space Sciences 64 巻 ( 3 ) 2021年5月

Design and testing of additively manufactured high-efficiency resistojet on hydrogen propellant 査読有り

Coral Giulio, Kinefuchi Kiyoshi, Nakata Daisuke, Tsukizaki Ryudo, Nishiyama Kazutaka, Kuninaka Hitoshi

ACTA ASTRONAUTICA 181 巻頁： 14 - 27 2021年4月

Speckle beam-oriented schlieren technique 査読有り

Nakamura Y., Suzuki T., Kinefuchi K., Sasoh A.

Experiments in Fluids 62 巻 ( 1 ) 2021年1月

記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Experiments in Fluids

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

Keeper ignition and discharge characteristics of hollow cathode center-mounted on hall Thruster 査読有り

Kinefuchi K., Cho S., Fukatsu T., Tsukizaki R., Funaki I., Hirano Y., Tashiro Y., Shiiki T.

Journal of Propulsion and Power 37 巻 ( 2 ) 頁： 223 - 230 2021年

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Journal of Propulsion and Power

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

Performance of a miniature hall thruster and an in-house PPU 査読有り

Chono M., Yamamoto N., Tsukizaki R., Morishita T., Kubota K., Cho S., Kinefuchi K., Takahashi T.

Transactions of the Japan Society for Aeronautical and Space Sciences 64 巻 ( 3 ) 頁： 189 - 192 2021年

Experimental analysis of thermal behavior in cryogenic propellant tank with different pressurants 査読有り

Kinefuchi K., Kawashima H., Sugimori D., Umemura Y., Okita K., Kobayashi H., Himeno T.

Cryogenics 112 巻 2020年12月

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Cryogenics

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

Application of a microwave cathode to a 200-W Hall thruster with comparison to a hollow cathode 査読有り

Morishita T., Tsukizaki R., Yamamoto N., Kinefuchi K., Nishiyama K.

Acta Astronautica 176 巻頁： 413 - 423 2020年11月

記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Acta Astronautica

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

Cryogenic propellant recirculation for orbital propulsion systems 査読有り

Kinefuchi Kiyoshi, Kawashima Hideto, Sugimori Daizo, Okita Koichi, Kobayashi Hiroaki

CRYOGENICS 105 巻 2020年1月

Characterization of a capillary flow controller for electric propulsion 査読有り

Kinefuchi K., Cho S., Tsukizaki R.

Journal of Propulsion and Power 36 巻 ( 4 ) 頁： 586 - 592 2020年

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Journal of Propulsion and Power

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

Application of two-photon laser-induced fluorescence spectroscopy to microwave cathode 査読有り Open Access

Tsukizaki R., Yamashita Y., Kinefuchi K., Nishiyama K.

Transactions of the Japan Society for Aeronautical and Space Sciences 63 巻 ( 6 ) 頁： 281 - 283 2020年

記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Transactions of the Japan Society for Aeronautical and Space Sciences

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 © 1019 to 8.4 © 1019 m3 with a maximum error of «20% due to the plasma fluorescence.

DOI： 10.2322/tjsass.63.281

In-flight S-band telemetry attenuation by ionized solid rocket motor plumes at high altitude 査読有り

Kinefuchi Kiyoshi, Yamaguchi Hiroyuki, Minami Mineko, Okita Koichi, Abe Takashi

ACTA ASTRONAUTICA 165 巻頁： 373-381 2019年12月

Neutral ground state particle density measurement of xenon plasma in microwave cathode by two-photon laser-induced fluorescence spectroscopy 査読有り

Yamashita Y., Tsukizaki R., Kinefuchi K., Koda D., Tani Y., Nishiyama K.

Vacuum 168 巻 2019年10月

記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Vacuum

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 5p61S06p[3/2]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 1019m−3 order without and with a plasma. In a cold gas, the neutral ground state particle density is (8.4±0.4)×1019m−3 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.

DOI： 10.1016/j.vacuum.2019.108846

Two-photon absorption laser induced fluorescence with various laser intensities for density measurement of ground state neutral xenon 査読有り

Kinefuchi K., Nunome Y., Cho S., Tsukizaki R., Chng T.L.

Acta Astronautica 161 巻頁： 382 - 388 2019年8月

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Acta Astronautica

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

Development of a new MLI for orbital cryogenic propulsion systems-Thermal performance under one atmosphere to a vacuum 査読有り

Takeshi Miyakita, Kazuya Kitamoto, Kiyoshi Kinefuchi, Masanori Saitoh, Tomoyuki Hirai, Hiroyuki Sugita

IOP Conference Series: Materials Science and Engineering 502 巻 ( 1 ) 2019年6月

Investigation of cryogenic chilldown in a complex channel under low gravity using a sounding rocket 査読有り

Kinefuchi K., Sarae W., Umemura Y., Fujita T., Okita K., Kobayashi H., Nonaka S., Himeno T., Sato T.

Journal of Spacecraft and Rockets 56 巻 ( 1 ) 頁： 91 - 103 2019年

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（国際会議プロシーディングス）出版者・発行元：Journal of Spacecraft and Rockets

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

全電化衛星査読有り

杵淵紀世志

日本航空宇宙学会誌 67 巻 ( 6 ) 頁： 227(J‐STAGE) - 227 2019年

Thermal performance and flow visualization of a planar heat-pipe leading edge 査読有り

Kinefuchi K., Tamba J., Saito I.

Journal of Spacecraft and Rockets 56 巻 ( 3 ) 頁： 771 - 779 2019年

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（国際会議プロシーディングス）出版者・発行元：Journal of Spacecraft and Rockets

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

LE-5B-3地上燃焼試験用静電容量型ボイド率計の開発査読有り

坂本勇樹, 小林弘明, 東和弘, 長尾直樹, 杉森大造, 杵淵紀世志, 佐藤哲也

航空宇宙技術(Web) 18 巻 ( 0 ) 頁： 19‐28(J‐STAGE) - 28 2019年

記述言語：日本語掲載種別：研究論文（学術雑誌）出版者・発行元：一般社団法人日本航空宇宙学会

<p>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' 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.</p>

DOI： 10.2322/astj.JSASS-D-18-00017

Numerical investigation of nanosecond pulsed plasma actuators for control of shock-wave/boundary-layer separation 査読有り Open Access

Kinefuchi K., Starikovskiy A.Y., Miles R.B.

Physics of Fluids 30 巻 ( 10 ) 2018年10月

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Physics of Fluids

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

Theoretical and experimental study of the active control of bubble point pressure using a magnetic field and its applications 査読有り

Kinefuchi K., Kobayashi H.

Physics of Fluids 30 巻 ( 6 ) 2018年6月

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Physics of Fluids

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

電気推進ロケットエンジンを用いたミッションの現状と今後 (小特集電気推進ロケットエンジン技術の現状と展望) 査読有り

西山和孝, 杵淵紀世志

プラズマ・核融合学会誌 = Journal of plasma and fusion research 94 巻 ( 2 ) 頁： 60 - 65 2018年2月

Control of shock-wave/boundary-layer interaction using nanosecond-pulsed plasma actuators 査読有り Open Access

Kinefuchi K., Starikovskiy A.Y., Miles R.B.

Journal of Propulsion and Power 34 巻 ( 4 ) 頁： 909 - 919 2018年

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（国際会議プロシーディングス）出版者・発行元：Journal of Propulsion and Power

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

微小重力環境下における極低温二相流のボイド率計測査読有り

小林弘明, 小林弘明, 坂本勇樹, 杵淵紀世志, 佐藤哲也

日本航空宇宙学会論文集 66 巻 ( 6 ) 頁： 147‐152(J‐STAGE) - 152 2018年

記述言語：日本語掲載種別：研究論文（学術雑誌）出版者・発行元：一般社団法人日本航空宇宙学会

<p>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 "measuring two-phase flow behavior and heat transfer characteristics during coasting flight." 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. </p>

DOI： 10.2322/jjsass.66.147

200W級ホールスラスタの開発査読有り

長野公勇, 森田太智, 山本直嗣, 窪田健一, 藤井剛, 杵淵紀世志

プラズマ応用科学 25 巻 ( 2 ) 頁： 85‐88 - 88 2017年

記述言語：日本語掲載種別：研究論文（学術雑誌）出版者・発行元：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, 神奈川県資料番号: SA6000147068レポート番号: STEP-2019-019

DOI： 10.34377/aps.25.2_85

全電化衛星の世界動向査読有り

杵淵紀世志, 杵淵紀世志

日本航空宇宙学会誌 65 巻 ( 9 ) 頁： 274‐279(J‐STAGE) - 279 2017年

記述言語：日本語掲載種別：研究論文（学術雑誌）出版者・発行元：一般社団法人日本航空宇宙学会

<p>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．</p>

DOI： 10.14822/kjsass.65.9_274

Prediction of in-flight radio frequency attenuation by a rocket plume 査読有り

Kinefuchi K., Okita K., Funaki I., Abe T.

Journal of Spacecraft and Rockets 52 巻 ( 2 ) 頁： 340 - 349 2015年3月

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（国際会議プロシーディングス）出版者・発行元：Journal of Spacecraft and Rockets

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

ロケット慣性飛行中の二相流挙動および熱伝達特性の観測実験~実験装置部の開発~ 査読有り

小林弘明, 杵淵紀世志, 更江渉, 梅村悠, 藤本圭一郎, 薮崎大輔, 杉森大造, 姫野武洋, 佐藤哲也, 北古賀智史, 角悠輝, 坂本勇樹, 野中聡, 藤田猛

日本航空宇宙学会論文集 63 巻 ( 5 ) 頁： 188-196 (J-STAGE) - 196 2015年

記述言語：日本語掲載種別：研究論文（学術雑誌）出版者・発行元：一般社団法人日本航空宇宙学会

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

Technical Challenges for Advanced Arcjets 査読有り

NAKATA Daisuke, KINEFUCHI Kiyoshi, HOSODA Satoshi, KINOSHITA Masahiro, KUNINAKA Hitoshi

TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, SPACE TECHNOLOGY JAPAN 12 巻 ( ists29 ) 頁： To_1_1 - To_1_5 2014年

記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：一般社団法人日本航空宇宙学会

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

Prediction of In-Flight Radio Frequency Attenuation by Rocket Plume Applying Diffraction Theories 査読有り

Kiyoshi Kinefuchi, Ikkoh Funaki, Takashi Abe

JOURNAL OF SPACECRAFT AND ROCKETS 50 巻 ( 1 ) 頁： 150-158 2013年1月

ロケットエンジン用リサーキュレーションポンプの試作と基礎的な試験について査読有り

渡邉光男, 橋本知之, 杵淵紀世志, 杉田栄一郎

ターボ機械 41 巻 ( 4 ) 頁： 193 - 200 2013年

微小重力下での液体推薬保持技術開発のための地上模擬実験査読有り

杵淵紀世志, 加納康仁, 齊藤靖博, 沖田耕一, 姫野武洋

航空宇宙技術(Web) 12 巻 ( 0 ) 頁： 73-77 (J-STAGE) - 77 2013年

担当区分：筆頭著者記述言語：日本語掲載種別：研究論文（学術雑誌）出版者・発行元：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

Computational fluid dynamics and frequency-dependent finite-difference time-domain method coupling for the interaction between microwaves and plasma in rocket plumes 査読有り

Kinefuchi K., Funaki I., Shimada T., Abe T.

Physics of Plasmas 19 巻 ( 10 ) 頁： 102112-102112-7 2012年10月

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Physics of Plasmas

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

Experimental investigation on thermochemical phenomena in SiFRP 査読有り

Hirai K., Matsuura Y., Kinefuchi K., Kamita T.

Advanced Composite Materials 21 巻 ( 5-6 ) 頁： 459 - 475 2012年10月

記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Advanced Composite Materials

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

SiFRP炭化変色層内部の物理現象の解明に関する実験研究査読有り

平井研一, 松浦芳樹, 杵淵紀世志, 紙田徹

日本複合材料学会誌 38 巻 ( 6 ) 頁： 228 - 235 2012年

記述言語：日本語掲載種別：研究論文（学術雑誌）出版者・発行元：一般社団法人日本複合材料学会

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

SiFRP加熱表面での物理現象の解明に関する実験研究査読有り

平井研一, 杵淵紀世志, 紙田徹

日本複合材料学会誌 38 巻 ( 5 ) 頁： 200 - 206 2012年

複合材ロケット構造の開発査読有り

杵淵紀世志, 夘沢俊行, 舘伊佐夫, 西元美希, 井川寛隆, 紙田徹

日本機械学会論文集 A編(Web) 77 巻 ( 773 ) 頁： 81 - 89 2011年

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：一般社団法人日本機械学会

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 ...

DOI： 10.1299/kikaia.77.81

Frequency-dependent FDTD simulation of the interaction of microwaves with rocket-plume 査読有り

Kinefuchi K., Funaki I., Abe T.

IEEE Transactions on Antennas and Propagation 58 巻 ( 10 ) 頁： 3282 - 3288 2010年10月

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：IEEE Transactions on Antennas and Propagation

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)2 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.

DOI： 10.1109/TAP.2010.2055796

Experimental and numerical simulation study of liquid-propellant draining from rocket tanks 査読有り

Kinefuchi K., Kamita T., Negishi H., Yamada K., Fujimura M.

Journal of Spacecraft and Rockets 47 巻 ( 5 ) 頁： 860 - 863 2010年

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Journal of Spacecraft and Rockets

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

Experimental investigation on microwave interference in full-scale solid rocket exhaust 査読有り

Kinefuchi K., Funaki I., Shimada T., Abe T.

Journal of Spacecraft and Rockets 47 巻 ( 4 ) 頁： 627 - 633 2010年

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Journal of Spacecraft and Rockets

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

Laser absorption velocimetry of plasma flow in two-dimensional magnetoplasmadynamic arcjet 査読有り

Kinefuchi K., Funaki I., Toki K.

Journal of Propulsion and Power 22 巻 ( 5 ) 頁： 1085 - 1090 2006年

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）出版者・発行元：Journal of Propulsion and Power

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

2次元型MPDアークジェットの速度場と電力収支計測査読有り

杵淵紀世志, 船木一幸, 都木恭一郎, 清水幸夫

日本航空宇宙学会論文集 53 巻 ( 616 ) 頁： 215 - 223 2005年

担当区分：筆頭著者記述言語：日本語掲載種別：研究論文（学術雑誌）出版者・発行元：社団法人日本航空宇宙学会

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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