永田 晴紀 (ナガタ ハルノリ)
工学研究院 機械・宇宙航空工学部門 宇宙航空システム | 教授 |
広域複合災害研究センター | 教授 |
高等教育推進機構 | 教授 |
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- 2023年03月, 日本機械学会宇宙工学部門, 業績賞
永田 晴紀 - 2019年08月, 米国航空宇宙学会, 2018 AIAA Hybrid Rockets Best Paper Award
Investigation of Graphite Nozzle Erosion in Hybrid Rockets Using O2/C2H4
Landon Kamps;Shota Hirai;Kazuhito Sakurai;Tor Viscor;Yuji Saito;Laymond Guan;Hikaru Isochi;Naoto Adachi;Mitsunori Ito;Harunori Nagata - 2018年08月, International Astronautical Federation, Academy member
永田 晴紀 - 2018年04月, 日本機械学会, フェロー会員
永田 晴紀 - 2017年07月, 米国航空宇宙学会, 2016 Hybrid Rockets Student Best Paper Award
Experimental and Analytical Investigation of Effect of Pressure on Regression Rate of Axial-Injection End-Burning Hybrid Rockets
Yuji Saito;Toshiki Yokoi;Hiroyuki Yasukochi;Kentaro Soeda;Tsuyoshi Totani;Masashi Wakita;Harunori Nagata - 2013年03月, 日本航空宇宙学会, フェロー会員
永田 晴紀 - 2008年, 日本航空宇宙学会, 日本航空宇宙学会賞(技術賞)
CAMUI型ハイブリッドロケット技術
永田 晴紀, 日本国 - 2006年, SI2006優秀講演賞
日本国 - 2005年05月, 日本機械学会北海道支部, 研究技術賞
CAMUI型ハイブリッドロケットの開発
永田 晴紀 - 2005年04月, 日本機械学会宇宙工学部門, 一般表彰フロンティアの部
CAMUI型ハイブリッドロケットの開発
永田 晴紀 - 2003年12月, 日本燃焼学会, 奨励賞
新形式ハイブリッドロケットの研究
永田 晴紀
論文
- GEOspace X-ray imager (GEO-X)
Yuichiro Ezoe, Ryu Funase, Harunori Nagata, Yoshizumi Miyoshi, Hiroshi Nakajima, Ikuyuki Mitsuishi, Kumi Ishikawa, Masaki Numazawa, Yosuke Kawabata, Shintaro Nakajima, Ryota Fuse, Ralf C. Boden, Landon Kamps, Tomokage Yoneyama, Kouichi Hagino, Yosuke Matsumoto, Keisuke Hosokawa, Satoshi Kasahara, Junko Hiraga, Kazuhisa Mitsuda, Masaki Fujimoto, Munetaka Ueno, Atsushi Yamazaki, Hiroshi Hasegawa, Takefumi Mitani, Yasuhiro Kawakatsu, Takahiro Iwata, Hiroyuki Koizumi, Hironori Sahara, Yoshiaki Kanamori, Kohei Morishita, Daiki Ishi, Aoto Fukushima, Ayata Inagaki, Yoko Ueda, Hiromi Morishita, Yukine Tsuji, Runa Sekiguchi, Takatoshi Murakawa, Kazuma Yamaguchi, Rei Ishikawa, Daiki Morimoto, Yudai Yamada, Shota Hirai, Yuki Nobuhara, Yownin Albert M. Leung, Yamato Itoigawa, Ryo Onodera, Satoru Kotaki, Shotaro Nakamura, Ayumi Kiuchi, Takuya Matsumoto, Midori Hirota, Kazuto Kashiwakura
Journal of Astronomical Telescopes, Instruments, and Systems, 9, 03, SPIE-Intl Soc Optical Eng, 2023年09月12日
研究論文(学術雑誌) - One-dimensional modelling of the nozzle cooling with cryogenic oxygen flowing through helical channels in a hybrid rocket
Giuseppe Gallo, Landon Kamps, Shota Hirai, Carmine Carmicino, Harunori Nagata
Acta Astronautica, 210, 176, 196, 2023年09月, [査読有り], [最終著者]
研究論文(学術雑誌), A one-dimensional model of the cooling process occurring in a hybrid-rocket nozzle working with cryogenic oxygen flowing through helical channels is addressed in this work. Although the regenerative cooling system in rocket engines involves further complexity, it is here investigated as an option for the suppression of graphite-nozzle erosion in hybrid rockets. The model is based on the solution of the conservation equations of mass, momentum, and energy of the coolant. The Modified Benedict-Webb-Rubin equation of state was used for its accuracy in both the fluid phases of oxygen. The conservation equations for the coolant flow are coupled with a one-dimensional heat transfer model for the evaluation of the nozzle wall temperature at both the cold and hot sides, which has allowed assessing the developed model prediction capabilities by means of the data collected from three engine hot firings. Thus, parametric analyses were carried out to show the effect of the number of helical channels and evaluate the performance improvement obtained in supercritical conditions. - Numerical Analysis of Nozzle Erosion in Hybrid Rockets and Comparison with Experiments
Daniele Bianchi, Mario Tindaro Migliorino, Marco Rotondi, Landon Kamps, Harunori Nagata
JOURNAL OF PROPULSION AND POWER, 38, 3, 389, 409, AMER INST AERONAUTICS ASTRONAUTICS, 2022年05月
英語, 研究論文(学術雑誌), © 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. The motor performance loss due to nozzle throat erosion in hybrid rocket engines is an important feature to be considered during the motor design. In fact, a peculiar characteristic of hybrid rockets operating conditions is a greater concentration of oxidizing species in the combustion products with respect to solid rockets, which emphasizes the throat erosion phenomenon. Hence, the need for an accurate physical understanding of this process is fundamental for technology advancement. The aim of this work is to investigate the graphite nozzle erosion in 2kN-class hybrid rockets burning high-density polyethylene (HDPE) and oxygen. The main focus of the work is the evaluation of nozzle throat erosion and wall temperature as a function of motor operating conditions by performing numerical simulations from a proven in-house CFD-solver. The results obtained from a detailed parametric analysis are discussed and then used to derive a regression law for nozzle throat erosion. The regression law here proposed is then validated through comparison with the experimental results obtained from lab-scale 2kN-thrust class rocket firing tests performed at Hokkaido University. - Fuel Regression Characteristics of Axial-Injection End-Burning Hybrid Rocket Using Nitrous Oxide
Ryota Okuda, Kodai Komizu, Ayumu Tsuji, Takumi Miwa, Mai Fukada, Shuichi Yokobori, Kentaro Soeda, Landon Kamps, Harunori Nagata
JOURNAL OF PROPULSION AND POWER, 1, 14, AMER INST AERONAUTICS ASTRONAUTICS, 2022年04月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), © 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. This study is an investigation of fuel regression characteristics in an axial-injection end-burning hybrid rocket using nitrous oxide. Previous studies of end-burning hybrid rockets used gaseous oxygen as the oxidizer. Nitrous oxide may be more suitable than the gaseous oxygen for use in space-based missions because of the weight savings associated with the oxidizer storage vessels, supply system, and motor mass. Previous research using nitrous oxide on a multiport fuel showed that stabilized combustion was not possible. In this study, two types of nozzle closures were employed to increase the initial chamber pressure and promote the formation of stabilized combustion in multiport fuels. The results of 12 firing tests showed that the regression rates when using nitrous oxide as the oxidizer were as high as that from previous research (0.61-4.5 mm/s at 0.25-0.75 MPa) using gaseous oxygen as the oxidizer. These high regression rates were nearly five times higher than that of experiments using single port fuels. It is clear from a visualization experiment that fuel flakes break off and travel downstream in solid form during firing, which could cause the fuel regression rate of multiport fuels to be higher than that of single port fuels. - GEO-X (GEOspace X-ray imager)
Yuichiro Ezoe, Ryu Funase, Harunori Nagata, Yoshizumi Miyoshi, Hiroshi Nakajima, Ikuyuki Mitsuishi, Kumi Ishikawa, Yosuke Kawabata, Shintaro Nakajima, Landon Kamps, Masaki Numazawa, Tomokage Yoneyama, Kouichi Hagino, Yosuke Matsumoto, Keisuke Hosokawa, Satoshi Kasahara, Junko Hiraga, Kazuhisa Mitsuda, Masaki Fujimoto, Munetaka Ueno, Atsushi Yamazaki, Hiroshi Hasegawa, Takefumi Mitani, Yasuhiro Kawakatsu, Takahiro Iwata, Hiroyuki Koizumi, Hironori Sahara, Yoshiaki Kanamori, Kohei Morishita
SPACE TELESCOPES AND INSTRUMENTATION 2022: ULTRAVIOLET TO GAMMA RAY, 12181, SPIE-INT SOC OPTICAL ENGINEERING, 2022年
英語, 研究論文(国際会議プロシーディングス), GEO-X (GEOspace X-ray imager) is a small satellite mission aiming at visualization of the Earth's magnetosphere by X-rays and revealing dynamical couplings between solar wind and magnetosphere. In-situ spacecraft have revealed various phenomena in the magnetosphere. In recent years, X-ray astronomy satellite observations discovered soft X-ray emission originated from the magnetosphere. We therefore develop GEO-X by integrating innovative technologies of the wide FOV X-ray instrument and the microsatellite technology for deep space exploration. GEO-X is a 50 kg class microsatellite carrying a novel compact X-ray imaging spectrometer payload. The microsatellite having a large delta v (>700 m/s) to increase an altitude at 40-60 R-E from relatively low-altitude (e.g., Geo Transfer Orbit) piggyback launch is necessary. We thus combine a 18U Cubesat with the hybrid kick motor composed of liquid N2O and polyethylene. We also develop a wide FOV (5x5 deg) and a good spatial resolution (10 arcmin) X-ray (0.3-2 keV) imager. We utilize a micromachined X-ray telescope, and a CMOS detector system with an optical blocking filter. We aim to launch the satellite around the 25th solar maximum. - Large-Scale CAMUI Type Hybrid Rocket Motor Scaling, Modeling, and Test Results
Tor Viscor, Landon Kamps, Kazuo Yonekura, Hikaru Isochi, Harunori Nagata
Aerospace, 9, 1, 1, 1, MDPI AG, 2021年12月
研究論文(学術雑誌), An understanding of the scalability of hybrid rocket regression models is critical for the enlargement and commercialization of small-scale engines developed within universities and similar research institutions. This paper investigates the fuel regression rates of recent 40 kN thrust-class motor experiments, which were designed based on fuel regression rate correlations of 2.5 kN thrust-class motors from previous research. The results show that fuel regression rates of the 40 kN experiments were within 26% of predictions made using correlations based on 2.5 kN experiments. - Burn Time Correction of Start-Up Transients for CAMUI Type Hybrid Rocket Engine
Tor Viscor, Hikaru Isochi, Naoto Adachi, Harunori Nagata
AEROSPACE, 8, 12, MDPI, 2021年12月
英語, 研究論文(学術雑誌), Burn time errors caused by various start-up transient effects have a significant influence on the regression modelling of hybrid rockets. Their influence is especially pronounced in the simulation model of the Cascaded Multi Impinging Jet (CAMUI) hybrid rocket engine. This paper analyses these transient burn time errors and their effect on the regression simulations for short burn time engines. To address these errors, the equivalent burn time is introduced and is defined as the time the engine would burn if it were burning at its steady-state level throughout the burn time to achieve the measured total impulse. The accuracy of the regression simulation with and without the use of equivalent burn time is then finally compared. Equivalent burn time is shown to address the burn time issue successfully for port regression and, therefore, also for other types of cylindrical port hybrid rocket engines. For the CAMUI-specific impinging jet fore-end and back-end surfaces, though, the results are inconclusive. - Hybrid Rockets as Post-Boost Stages and Kick Motors
Landon Kamps, Shota Hirai, Harunori Nagata
AEROSPACE, 8, 9, MDPI, 2021年09月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), Hybrid rockets are attractive as post-boost stages and kick motors due to their inherent safety and low cost, but it is not clear from previous research which oxidizer is most suitable for maximizing Delta V within a fixed envelope size, or what impact O/F shift and nozzle erosion will have on Delta V. A standard hybrid rocket design is proposed and used to clarify the impact of component masses on Delta V within three 1 m(3) envelopes of varying height-to-base ratios. Theoretical maximum Delta V are evaluated first, assuming constant O/F and no nozzle erosion. Of the four common liquid oxidizers: H2O2 85 wt%, N2O, N2O4, and LOX, H2O2 85 wt% is shown to result in the highest Delta V, and N2O is shown to result in the highest density Delta V, which is the Delta V normalized for motor density. When O/F shift is considered, the Delta V decreases by 9% for the N2O motor and 12% for the H2O2 85 wt% motor. When nozzle erosion is also considered, the Delta V decreases by another 7% for the H2O2 85 wt% motor and 4% for the N2O motor. Even with O/F shift and nozzle erosion, the H2O2 85 wt% motor can accelerate itself (916 kg) upwards of 4000 m/s, and the N2O motor (456 kg) 3550 m/s. - Reduction Mechanism of Transition Metal Oxide Particles in Thermally Induced Nanobubbles during Pulsed Laser Melting in Ethanol
Kentaro Suehara, Ryosuke Takai, Yoshie Ishikawa, Naoto Koshizaki, Kazunobu Omura, Harunori Nagata, Yuji Yamauchi
CHEMPHYSCHEM, 22, 7, 675, 683, WILEY-V C H VERLAG GMBH, 2021年04月, [査読有り]
英語, 研究論文(学術雑誌), Pulsed laser melting in liquid (PLML) is a technique to fabricate spherical submicrometer particles (SMPs) wherein nanosecond pulsed laser (several tens to several hundreds of mJ pulse(-1) cm(-2)) irradiates raw particles dispersed in liquid. Raw particles are transiently heated above the melting point to form spherical particles, which enables pulsed heating of surrounding liquid to form thermally induced bubbles by liquid vaporization. These transient bubbles play an important role as a thermal barrier to rapidly heat the particle. Reduced SMPs are generated from raw metal-oxide nanoparticles by PLML process in ethanol. This reduction cannot be explained by high-temperature thermal decomposition, but by mediation of molecules decomposed from ethanol. Computational simulations of ethanol decomposition by pulsed heating for 100 ns at the temperature 1000-4000 K revealed that ethylene is generated as the main product. Gibbs free energies of oxide reduction reactions mediated by ethylene greatly decreased compared to those without ethylene mediation. This explanation can be applied to reductive SMP formation from various transition metal oxides by PLML. - Fuel Regression Characteristics in Hybrid Rockets Using Nitrous Oxide/High-Density Polyethylene
Seiji Ito, Landon Kamps, Harunori Nagata
JOURNAL OF PROPULSION AND POWER, 37, 2, 342, 348, AMER INST AERONAUTICS ASTRONAUTICS, 2021年03月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌) - GEO-X (GEOspace X-ray imager)
Yuichiro Ezoe, Ryu Funase, Harunori Nagata, Yoshizumi Miyoshi, Satoshi Kasahara, Hiroshi Nakajima, Ikuyuki Mitsuishi, Kumi Ishikawa, Junko S. Hiraga, Kazuhisa Mitsuda, Masaki Fujimoto, Munetaka Ueno, Atsushi Yamazaki, Hiroshi Hasegawa, Yosuke Matsumoto, Yasuhiro Kawakatsu, Takahiro Iwata, Hironori Sahara, Yoshiaki Kanamori, Kohei Morishita, Hiroyuki Koizumi, Makoto Mita, Takefumi Mitani, Masaki Numazawa, Landon Kamps, Yusuke Kawabata
SPACE TELESCOPES AND INSTRUMENTATION 2020: ULTRAVIOLET TO GAMMA RAY, 11444, SPIE-INT SOC OPTICAL ENGINEERING, 2021年, [査読有り]
英語, 研究論文(国際会議プロシーディングス), GEO-X (GEOspace X-ray imager) is a 50 kg-class small satellite to image the global Earth's magnetosphere in X-rays via solar wind charge exchange emission. A 12U CubeSat will be injected into an elliptical orbit with an apogee distance of similar to 40 Earth radii. In order to observe the diffuse soft X-ray emission in 0.3-2 keV and to verify X-ray imaging of the dayside structures of the magnetosphere such as cusps, magnetosheaths and magnetopauses which are identified statistically by in-situ satellite observations, an original light-weight X-ray imaging spectrometer (similar to 10 kg, similar to 10 W, similar to 10x10x30 cm) will be carried. The payload is composed of a ultra light-weight MEMS Wolter type-I telescope (similar to 4x4 deg(2) FOV, <10 arcmin resolution) and a high speed CMOS sensor with a thin optical blocking filter (similar to 2x2 cm(2), frame rate similar to 20 ms, energy resolution <80 eV FWHM at 0.6 keV). An aimed launch year is 2023-25 corresponding to the 25th solar maximum. - Stabilized combustion of circular fuel duct with liquid oxygen
Ayumu Tsuji, Yuji Saito, Landon Kamps, Masashi Wakita, Harunori Nagata
PROCEEDINGS OF THE COMBUSTION INSTITUTE, 38, 3, 4845, 4855, ELSEVIER SCIENCE INC, 2021年, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), This research is an investigation of the flame spread opposed to a liquid oxidizer flow in a solid fuel duct. Several firing tests were conducted using liquid oxygen as the oxidizer and solid poly methyl methacrylate (PMMA) as the fuel. The results indicate that the flame spread rate decreased with increasing oxidizer port velocity and decreasing port diameter. This study reveals through visual confirmations and empirical correlations of the flame spread rate that the flame spread opposed to liquid oxygen in a solid fuel duct can be classified as stabilized combustion. Extinction and abnormal regression were observed when oxidizer port velocity was high and port diameter was small. Furthermore, the cooling of the solid fuel by the liquid oxygen flow had a strong effect on the transition between normal regression and extinction, or abnormal regression. A model of the flame spread rate which considers the heat balance at the fuel surface assuming a fully developed thermal boundary layer is introduced and shown to agree well with the experimental results. Lastly, it is revealed that the difference in kinematic viscosity between liquid oxygen and gaseous oxygen is the main reason dependency of port diameter on flame spread rates differs between the liquid oxygen tests in this study and gaseous oxygen tests in previous studies.(c) 2020 The Combustion Institute. Published by Elsevier Inc. All rights reserved. - Investigation of Graphite Nozzle Erosion in Hybrid Rockets Using Oxygen/High-Density Polyethylene
Landon Kamps, Shota Hirai, Kazuhito Sakurai, Tor Viscor, Yuji Saito, Raymond Guan, Hikaru Isochi, Naoto Adachi, Mitsunori Itoh, Harunori Nagata
JOURNAL OF PROPULSION AND POWER, 36, 3, 423, 434, AMER INST AERONAUTICS ASTRONAUTICS, 2020年05月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), A recently developed reconstruction technique is used to investigate graphite nozzle erosion in two scales of hybrid rocket motors, 30-N-thrust class and 2000-N-thrust class, using oxygen as the oxidizer and high-density polyethylene as the fuel. Thermocouple measurements taken from within the nozzles are used to estimate nozzle throat wall temperature. Forty-four static firing tests were conducted under varying experimental conditions to confirm the validity of the reconstruction technique results, to investigate the conditions at the onset of erosion, and to formulate an empirical predictive model of nozzle erosion rate. Results show that a single formula that treats the combustion gas as a single oxidizing agent for which heterogenous rate constants are functions of equivalence ratio can satisfactorily replicate the erosion rate of graphite by a combustion gas containing multiple oxidizing species. Furthermore, the chemical-kinetic-limited conditions of the onset of nozzle erosion are specified by a novel empirical correlation, which shows that erosion begins at lower temperature and pressure in oxidizer-rich combustion gas than in fuel-rich combustion gas. - Experimental Investigation of the Continuous Transition of Flame-Spreading near the Blow-Off Limit
K. Komizu, Y. Saito, A. Tsuji, H. Nagata
Journal of Combustion, 2020, 2020年, [査読有り], [最終著者]
研究論文(学術雑誌), © 2020 K. Komizu et al. This study investigates the continuous transition from flame-spreading to stabilized combustion near the blow-off limit in opposed forced flow by using expanding solid fuel duct that makes distribution of oxidizer velocity in the axial direction. The stabilized combustion is a diffusion flame that appears in the Axial-Injection End-Burning Hybrid Rocket. The boundary between flame-spreading and stabilized combustion has not been investigated in detail. Polymethyl methacrylate (PMMA) rectangular ducts were used as a fuel, and gaseous oxygen was used as an oxidizer. All firing tests were conducted at atmospheric pressure. The diffusion flame traveled in the opposed-flow field where the oxidizer velocity increases continuously in the upstream direction. The combustion mode changed when oxidizer velocity at the flame tip exceeded a certain value. The oxidizer velocity used in this experiment ranges from 0.6 to 32.8 m/s. Experimental results show that a threshold oxidizer velocity of the transition can be determined. In this study, the threshold velocity was 26.4 m/s. - Controlling heat release of crystallization from supercooling state of a solid-solid PCM, 2-amino-2-methyl-1,3-propanediol
Ryohei Gotoh, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata
INTERNATIONAL JOURNAL OF HEAT AND MASS TRANSFER, 137, 1132, 1140, PERGAMON-ELSEVIER SCIENCE LTD, 2019年07月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), The use of phase change materials (PCMs) for heat storage and as a heat source has become an important aspect for energy management. Some PCMs store energy when in a non-equilibrium state (a supercooling state), and supply energy when released from this state. This means PCMs have the ability to sustain heat energy for long periods and select the heat supply timing. 2-amino-2-methyl-1,3-propanediol (AMP), a solid-solid PCM, stores about 264 J/g of heat energy at the crystal transition temperature of about 78 degrees C. AMP has the attractive characteristic of storing heat energy in its solid supercooling state, similar to solid-liquid PCMs. In addition, AMP crystallizes from the supercooling state and releases heat energy of about 140 J/g during the heating process. These positive attributes make AMP a good candidate to assist in heating a system. This study applied this characteristic to methods handling the exoergic heat energy of the crystallization of AMP. First, the thermal properties are studied by DSC measurement and thermal cycle tests in different mass conditions. Second, the crystallization is investigated by observation of crystal growth. The results show that the supercooling state crystallizes with exoergic heat during the heating process. It turns out that the crystal nucleation rate (1/s) highly depends on the temperature and AMP mass. The crystal growth rate (mu m/s) is acquired in this experiment. By using this information, it is possible to handle the exoergic heat of the crystallization from the supercooling state by changing the AMP mass and minimum temperature during cooling. Moreover, the heat energy that is kept in the supercooling state can be also controlled by crystal nucleus addition or impact. (C) 2019 Elsevier Ltd. All rights reserved. - Comprehensive Data Reduction for N2O/HDPE Hybrid Rocket Motor Performance Evaluation
Landon Kamps, Kazuhito Sakurai, Yuji Saito, Harunori Nagata
AEROSPACE, 6, 4, MDPI, 2019年04月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), Static firing tests of a hybrid rocket motor using liquid nitrous oxide (N2O) as the oxidizer and high-density polyethylene (HPDE) as the fuel are analyzed using a novel approach to data reduction that allows histories for fuel mass consumption, nozzle throat erosion, characteristic exhaust velocity (c) efficiency, and nozzle throat wall temperature to be determined experimentally. This is done by firing a motor under the same conditions six times, varying only the burn time. Results show that fuel mass consumption was nearly perfectly repeatable, whereas the magnitude and timing of nozzle throat erosion was not. Correlations of the fuel regression rate result in oxidizer port mass flux exponents of 0.62 and 0.76. There is a transient time in the c efficiency histories of around 2.5 s, after which c efficiency remains relatively constant, even in the case of excessive nozzle throat erosion. Although nozzle erosion was not repeatable, the erosion onset factors were similar between tests, and greater than values in previous research in which oxygen was used as the oxidizer. Lastly, nozzle erosion rates exceed 0.15 mm/s for chamber pressures of 4 to 5 MPa. - High Pressure Fuel Regression Characteristics of Axial-Injection End-Burning Hybrid Rockets
Yuji Saito, Masaya Kimino, Ayumu Tsuji, Yushi Okutani, Kentaro Soeda, Harunori Nagata
JOURNAL OF PROPULSION AND POWER, 35, 2, 328, 341, AMER INST AERONAUTICS ASTRONAUTICS, 2019年03月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), This study is an investigation of axial-injection end-burning hybrid rockets aimed at revealing fuel regression characteristics under relatively high-pressure conditions. Firing tests are conducted using gaseous oxygen as the oxidizer at chamber pressures and oxidizer port velocities ranging from 0.22 to 1.05 MPa and 31 to 103 m/s, respectively. The results of 15 static firing tests show that the fuel regression rate increases as the chamber pressure increases, and regression rates range from approximately 1.1 mm/s at 0.25 MPa to 5.5 mm/s at 0.90 MPa. Furthermore, it is observed that the pressure exponent of the fuel regression rate is 1.05 and the fuel regression rate is not influenced by the oxidizer port velocity in this study. The model explains that the backfiring problem tends to occur in relatively high-pressure conditions, and it leads to the conclusion that increasing the nozzle throat diameter is an effective means of preventing backfiring from occurring. - Exhaust Heat Characteristics of Single Liquid Droplet Stream for Liquid Droplet Radiator
TAKANASHI Tomohiro, TOTANI Tsuyoshi, SHIMADA Taizo, RYOMON Kento, WAKITA Masashi, NAGATA Harunori
日本伝熱学会論文集, 27, 1, 43, 52, 社団法人 日本伝熱学会, 2019年, [査読有り], [最終著者]
日本語, In order to realize a liquid droplet radiator (LDR), which is an equipment used for waste heat rejection in large space structures, the exhaust heat characteristics of a single liquid droplet stream in vacuum are required. In this study, these characteristics were obtained by a combination of experiments and numerical analyses. Experiments were conducted for measuring the amount of waste heat coming from a single liquid droplet stream of silicone oil as the working fluid, using a radiant flux sensor (RFS). The emissivity of the RFS at low temperature was measured, using a black radiation ball with known emissivity at 26°C. Numerical analyses were also performed to separate the radiation from the background of the experimental apparatus included in the radiation measured by the RFS. It was determined that at approximately -180°C, the emissivity of the RFS falls from the catalog value of 0.8 to 0.5. Moreover, the emissivity of the liquid droplet was found to be approximately in the range 0.42–0.51 and the effective emissivity of the single droplet stream was approximately 0.10. The application of these results can improve the feasibility of LDRs. - Tubular Equivalent Regression Rate in Hybrid Rockets with Complex Geometries
KAMPS Landon, NAGATA Harunori
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 17, 4, 544, 551, 一般社団法人 日本航空宇宙学会, 2019年, [査読有り], [最終著者]
英語,A new performance parameter titled “tubular equivalent regression rate” is introduced to evaluate burning rates in hybrid rockets with geometrically complex solid propellant grains. Tubular equivalent regression rates are calculated for eight previously reported CAMUI-type hybrid rocket firing tests and compared with extrapolations of previously reported empirical correlations for classic, swirl and vortex hybrid rockets. A non-dimensional number titled “CAMUI Number” is introduced to evaluate how CAMUI-like a solid propellant grain is. The CAMUI Number ranges from 0-1: 0 means no CAMUI-type blocks are used, 1 means only CAMUI-type blocks are used. The results show that the tubular equivalent regression rate increases logarithmically with CAMUI Number, and approaches a value of around 3 [mm/s] for a CAMUI Number of 1. This increase in tubular equivalent regression rate is shown to correspond to an increase in performance range from a classic (tubular) hybrid rocket at low CAMUI Numbers (0.1) to surpassing a vortex hybrid rocket for high CAMUI Numbers (>0.7). Furthermore, through the block-by-block analysis of tubular equivalent regression rate in a fuel grain with a CAMUI Number of 0.71, it is shown that maximum burning rates were achieved in blocks under slightly oxidizer rich conditions.
- Error Analysis for CAMUI Type Hybrid Rocket Regression Simulation
TOR Viscor, NAGATA Harunori
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 17, 4, 519, 524, 一般社団法人 日本航空宇宙学会, 2019年, [査読有り], [最終著者]
英語,This paper describes the error and uncertainty analysis of the CAMUI hybrid rocket regression simulator. Simulation errors compared to test firings are described and followed by an analysis of the potential uncertainties causing this error. For each uncertainty identified, a sensitivity analysis is then performed with the help of a custom-built simulator to evaluate its impact on the simulator accuracy. It was found that uncertainties in LOX travel time, Reynolds number grouping and model assumptions for the first upstream burning surface have the largest impact on the simulator accuracy and are identified as the main focus points for further research.
- 端面燃焼式ハイブリッドロケットの推力制御時におけるヒステリシス特性に関する研究
君野 正弥, 齋藤 勇士, 奥谷 勇士, 津地 歩, 添田 建太郎, 永田 晴紀
日本航空宇宙学会論文集, 67, 4, 119, 125, 一般社団法人 日本航空宇宙学会, 2019年, [査読有り], [最終著者]
日本語,In this study, the authors conduced ten firings to investigate a hysteresis characteristics in Axial-Injection End-Burning hybrid rockets under throttling operation. Oxidizer mass flow rate and chamber pressure were throttled by two methods, actuating valves in a fluid circuit consisting of two oxidizer supply lines and a motor controlling. Chamber pressure and oxidizer mass flow rate were measured during each firing. The results show that two types of hysteresis characteristics were observed when throttling operation is repeated. One is a hysteresis with respect to increase and decrease of the oxidizer mass flow rate. Another is a hysteresis for the cycle. It is considered that the former hysteresis has the influence of a chamber pressure response time. In addition, the latter hysteresis is not necessarily observed even in the same chamber pressure region.
- Ultralightweight x-ray telescope missions: ORBIS and GEO-X
Yuichiro Ezoe, Yoshizumi Miyoshi, Satoshi Kasahara, Tomoki Kimura, Kumi Ishikawa, Masaki Fujimoto, Kazuhisa Mitsuda, Hironori Sahara, Naoki Isobe, Hiroshi Nakajima, Takaya Ohashi, Harunori Nagata, Ryu Funase, Munetaka Ueno, Graziella Branduardi-Raymont
JOURNAL OF ASTRONOMICAL TELESCOPES INSTRUMENTS AND SYSTEMS, 4, 4, SPIE-SOC PHOTO-OPTICAL INSTRUMENTATION ENGINEERS, 2018年10月, [査読有り]
英語, 研究論文(学術雑誌), Toward an era of x-ray astronomy, next-generation x-ray optics are indispensable. To meet a demand for telescopes lighter than the foil optics but with a better angular resolution <1 arcmin, we are developing micropore x-ray optics based on micromaching technologies. Using sidewalls of micropores through a thin silicon wafer, this type can be the lightest x-ray telescope ever achieved. Two Japanese missions, ORBIS and GEO-X, will carry this telescope. ORBIS is a small x-ray astronomy mission to monitor supermassive blackholes, while GEO-X is a small exploration mission of the Earth's magnetosphere. Both missions need an ultralight-weight (<1 kg) telescope with moderately good angular resolution (<10 arcmin) at an extremely short focal length (<30 cm). We plan to demonstrate this type of telescope in these two missions around 2020. (C) The Authors. Published by SPIE under a Creative Commons Attribution 3.0 Unported License. - Preliminary Thermal Design for Microsatellites Deployed from International Space Station's Kibo Module
Delburg P. Mitchao, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata
JOURNAL OF THERMOPHYSICS AND HEAT TRANSFER, 32, 3, 789, 798, AMER INST AERONAUTICS ASTRONAUTICS, 2018年07月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), A preliminary thermal design was proposed by determining all the possible combinations of solar absorptivity and infrared emissivity on the panel surfaces of Earth-pointing satellites deployed from Japan's experimental module small-satellite orbital deployer. The three most common internal panel configurations in a 50 kg satellite with body-mounted solar cells and dimensions of 550 x 350 x 550 mm were considered. Conductive insulation was applied between the inner and outer structures to decrease the temperature change of inner components. The worst hot-and cold case conditions were estimated based on the beta angle of the orbit and the Earth's distance from the sun. The analyses were carried out using a simple tool created in MATLAB (c). The tool output combinations of optical properties that satisfied the predefined allowable temperature range of structures and components. These optical properties were subsequently verified using Thermal Desktop (c)'s SINDA/FLUINT with the RadCAD module. Using these combinations, the thermal design for microsatellites in a low Earth and non-sun-synchronous orbit may be shortened. - Investigation of Throttling Response Characteristics of Axial-Injection End-Burning Hybrid Rockets
Yuji SAITO, Masaya KIMINO, Ayumu TSUJI, Kazunobu OMURA, Hiroyuki YASUKOCHI, Kentaro SOEDA, Tsuyoshi TOTANI, Masashi WAKITA, Harunori NAGATA
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 16, 1, 9, 18, Japan Society for Aeronautical and Space Sciences, 2018年
英語, 研究論文(学術雑誌) - 端面燃焼式ハイブリッドロケットの超小型衛星への応用
齋藤 勇士, 君野 正弥, 添田 建太郎, 戸谷 剛, 永田 晴紀
日本航空宇宙学会誌, 66, 10, 291, 295, 一般社団法人 日本航空宇宙学会, 2018年, [招待有り], [最終著者]
日本語,超小型衛星の運用の高機能化および深宇宙探査には,推進器がますます不可欠な存在となる.化学エネルギを用いて大推力を得ることのできる化学ロケットは数km/sの増速が与えられるキックモータになる.主衛星に相乗りする形で打ち上げられる超小型衛星には厳格な安全基準が求められるため,プラスチック等を燃料とするハイブリッドロケットが注目を浴びている.その中でも,端面燃焼式ハイブリッドロケットは,従来型ハイブリッドロケットを凌駕する燃焼特性および推力制御特性が期待されてきた.端面燃焼式ハイブリッドロケットは,燃料製作に困難さを有していたが,高精度3Dプリンタの台頭によって,2014年に実証に成功した.これまでの間,筆者らは数多くの燃焼実験を実施し,研究成果を国内外の学会で発表してきた.本論文では今まで得られた端面燃焼式ハイブリッドロケットの知見をまとめ,今後の課題を紹介する.
- Small satellites with MEMS X-ray telescopes for X-ray astronomy and solar system exploration
Yuichiro Ezoe, Yoshizumi Miyoshi, Satoshi Kasahara, Tomoki Kimura, Kumi Ishikawa, Masaki Fujimoto, Kazuhisa Mitsuda, Hironori Sahara, Naoki Isobe, Hiroshi Nakajima, Takaya Ohashi, Harunori Nagata, Ryu Funase, Munetaka Ueno, Graziella Branduardi-Raymont
SPACE TELESCOPES AND INSTRUMENTATION 2018: ULTRAVIOLET TO GAMMA RAY, 10699, SPIE-INT SOC OPTICAL ENGINEERING, 2018年, [査読有り]
英語, 研究論文(国際会議プロシーディングス), Toward a new era of X-ray astronomy, next generation X-ray optics are indispensable. To meet a demand for telescopes lighter than the foil optics but with a better angular resolution less than 1 arcmin, we are developing micropore X-ray optics based on micromaching technologies. Using sidewalls of micropores through a thin silicon wafer, this type can be the lightest X-ray telescope ever achieved. Two new Japanese missions ORBIS and GEO-X will carry this optics. ORBIS is a small X-ray astronomy mission to monitor supermassive blackholes, while GEO-X is a small exploration mission of the Earth's magnetosphere. Both missions need a ultra light-weight (<1 kg) telescope with moderately good angular resolution (<10 arcmin) at an extremely short focal length (<30 cm). We plan to demonstrate this optics in these two missions around 2020, aiming at future other astronomy and exploration missions. - Fuel Regression Characteristics of a Novel Axial-Injection End-Burning Hybrid Rocket
Yuji Saito, Toshiki Yokoi, Hiroyuki Yasukochi, Kentaro Soeda, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata
JOURNAL OF PROPULSION AND POWER, 34, 1, 247, 259, AMER INST AERONAUTICS ASTRONAUTICS, 2018年01月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), The regression characteristics of axial-injection end-burning hybrid rocket were experimentally investigated using a laboratory-scale motor. The axial-injection end-burning type fuel grains were made by high-accuracy three-dimensional printing. Firing tests were conducted using gaseous oxygen as the oxidizer at a chamber pressure range of 0.10 to 0.43MPa. Results of 15 static firings tests show that fuel regression rate increases as the chamber pressure rises, and fuel regression rate decreases as the oxidizer port velocity increases. A data reduction method was developed to avoid the difficulty in calculating oxidizer-to-fuel ratio. A simplified fuel regression model based on the granular diffusion flame model is developed to investigate regression characteristics. The trend in results as calculated using the granular diffusion flame model agrees with that in experimentally observe values. However, this does not hold true in tests with varying oxidizer port velocity. A granular diffusion flame model only takes into account simple solid propellant regression. Therefore, modification of the model is needed for calculating the fuel regression rate of an end-burning hybrid rocket. - Verification Firings of End-Burning Type Hybrid Rockets
Harunori Nagata, Hayato Teraki, Yuji Saito, Ryuichiro Kanai, Hiroyuki Yasukochi, Masashi Wakita, Tsuyoshi Totani
JOURNAL OF PROPULSION AND POWER, 33, 6, 1473, 1477, AMER INST AERONAUTICS ASTRONAUTICS, 2017年11月, [査読有り], [筆頭著者]
英語, 研究論文(学術雑誌), The authors have previously proposed the concept of end-burning-type hybrid rockets, which would use cylindrical fuel grains consisting of an array of many small ports running in the axial direction, through which oxidizer gas would flow. Because of difficulty in manufacturing a fuel grain that satisfied requirements such as high volumetric filling rate (above 0.95) and microsized port intervals, the end-burning hybrid rocket had yet to be achieved. This paper reports the results of verification firing tests of a novel end-burning-type hybrid rocket made possible for the first time by recent progress in three-dimensional printing technology. The results clearly distinguish the initial transient and steady periods of the end-burning mode and prove that no oxidizer-to-fuel ratio shift occurs during firing. Because the initial transient is a period for the exit end face to attain a steady-state shape, an initial end-face shape being close to the steady-state shape can shorten this period. A firing test with fuel having tapered ports is shown to attain a steady-state shape in less than 1s, which is much shorter than the nontapered case of about 6 s. - Method for Determining Nozzle-Throat-Erosion History in Hybrid Rockets
Landon Kamps, Yuji Saito, Ryosuke Kawabata, Masashi Wakita, Tsuyoshi Totani, Yusuke Takahashi, Harunori Nagata
JOURNAL OF PROPULSION AND POWER, 33, 6, 1369, 1377, AMER INST AERONAUTICS ASTRONAUTICS, 2017年11月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), The authors of this paper introduce a new reconstruction technique titled nozzle-throat reconstruction technique to estimate nozzle-throat-erosion history and oxidizer-to-fuel-mass-ratio history in hybrid rockets. Nine static-firing tests were carried out on a 2kN-class cascaded multistage impinging-jet-type hybrid-rocket motor under varying oxidizer-flow rates to evaluate the accuracy of reconstructed results. Nozzle-throat-erosion histories calculated by the nozzle-throat reconstruction technique agreed well with measured values for initial nozzle-throat radius, and successfully reconstructed the case, in which no measurable amount of nozzle-throat erosion occurred. For equivalence ratios 0.6-1.4, the relationship between nozzle-throat-erosion rate and equivalence ratio of reconstructed results displays a trend consistent with chemical-kinetic-limited heterogeneous-combustion theory, as well as predictions made by previous researchers. - Investigation of axial-injection end-burning hybrid rocket motor regression
Yuji Saito, Toshiki Yokoi, Lukas Neumann, Hiroyuki Yasukochi, Kentaro Soeda, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata
ADVANCES IN AIRCRAFT AND SPACECRAFT SCIENCE, 4, 3, 281, 296, TECHNO-PRESS, 2017年05月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), The axial-injection end-burning hybrid rocket proposed twenty years ago by the authors recently recaptured the attention of researchers for its virtues such as no xi (oxidizer to fuel mass ratio) shift during firing and good throttling characteristics. This paper is the first report verifying these virtues using a laboratory scale motor. There are several requirements for realizing this type of hybrid rocket: 1) high fuel filling rate for obtaining an optimal xi ; 2) small port intervals for increasing port merging rate; 3) ports arrayed across the entire fuel section. Because these requirements could not be satisfied by common manufacturing methods, no previous researchers have conducted experiments with this kind of hybrid rocket. Recent advances in high accuracy 3D printing now allow for fuel to be produced that meets these three requirements. The fuel grains used in this study were produced by a high precision light polymerized 3D printer. Each grain consisted of an array of 0.3 mm diameter ports for a fuel filling rate of 98%. The authors conducted several firing tests with various oxidizer mass flow rates and chamber pressures, and analysed the results, including xi history, using a new reconstruction technique. The results show that xi remains almost constant throughout tests of varying oxidizer mass flow rates, and that regression rate in the axial direction is a nearly linear function of chamber pressure with a pressure exponent of 0.996. - Orbit manipulation by use of lunar swing-by on a hyperbolic trajectory
Shuntaro Suda, Yasuhiro Kawakatsu, Shujiro Sawai, Harunori Nagata, Tsuyoshi Totani
Advances in the Astronautical Sciences, 160, 4027, 4041, 2017年
研究論文(国際会議プロシーディングス), In the modern space development, small-scale deep space mission should be realized to promote frequent and challenging deep space mission. Therefore, the efficient and quick design method to construct Earth escape trajectory with high flexibility in the boundary condition such as escape velocity, direction and timing is strongly demanded. In this paper, the families of Moon-to-Moon transfers with sequential lunar swing-by on a hyperbolic orbit are computed and stored in a database. These families are useful to enhance the Earth escape energy and to change escape direction which could lead a spacecraft to further destinations. - Hybrid Propulsion Technology Development in Japan for Economic Space Launch
Toru Shimada, Saburo Yuasa, Harunori Nagata, Shigeru Aso, Ichiro Nakagawa, Keisuke Sawada, Keiichi Hori, Masahiro Kanazaki, Kazuhisa Chiba, Takashi Sakurai, Takakazu Morita, Koki Kitagawa, Yutaka Wada, Daisuke Nakata, Mikiro Motoe, Yuki Funami, Kohei Ozawa, Tomoaki Usuki
CHEMICAL ROCKET PROPULSION: A COMPREHENSIVE SURVEY OF ENERGETIC MATERIALS, 545, 575, SPRINGER-VERLAG BERLIN, 2017年, [査読有り]
英語, 研究論文(国際会議プロシーディングス), The demand for the economic and dedicated space launchers for vast amount of lightweight, so-called nano-/microsatellites, is now growing rapidly. There is a strong rationale for the usage of the hybrid propulsion for economic space launch as suggested by the assessment conducted here. A typical concept of development of such an economic three-stage launcher, in which clustering unit hybrid rocket engines are employed, is described with a development scenario. Thanks to the benefits of hybrid rocket propulsion, assuring and safe, economic launcher dedicated to lightweight satellites can be developed with a reasonable amount of quality assurance and quality control actions being taken in all aspects of development such as raw material, production, transportation, storage, and operation. By applying a multi-objective optimization technique for such a launch system, examples of possible launch systems are obtained for a typical mission scenario for the launch of lightweight satellites. Furthermore, some important technologies that contribute strongly to economic space launch by hybrid propulsion are described. They are the behavior of fuel regression rate, the swirling-oxidizerflow- type hybrid rocket, the liquid oxygen vaporization, the multi-section swirl injection, the low-temperature melting point thermoplastic fuel, the thrust and O/F simultaneous control by altering-intensity swirl-oxidizer-flow-type (A-SOFT) hybrid, the numerical simulations of the internal ballistics, and so on. - ORBIT MANIPULATION BY USE OF LUNAR SWING-BY ON A HYPERBOLIC TRAJECTORY
Shuntaro Suda, Yasuhiro Kawakatsu, Shujiro Sawai, Harunori Nagata, Tsuyoshi Totani
SPACEFLIGHT MECHANICS 2017, PTS I - IV, 160, 4027, 4041, UNIVELT INC, 2017年, [査読有り]
英語, 研究論文(国際会議プロシーディングス), In the modern space development, small-scale deep space mission should be realized to promote frequent and challenging deep space mission. Therefore, the efficient and quick design method to construct Earth escape trajectory with high flexibility in the boundary condition such as escape velocity, direction and timing is strongly demanded. In this paper, the families of Moon-to-Moon transfers with sequential lunar swing-by on a hyperbolic orbit are computed and stored in a database. These families are useful to enhance the Earth escape energy and to change escape direction which could lead a spacecraft to further destinations. - 端面燃焼式ハイブリッドロケットの推力制御特性に関する研究
齋藤 勇士, 横井 俊希, 津地 歩, 尾村 和信, 安河内 裕之, 添田 建太郎, 戸谷 剛, 脇田 督司, 永田 晴紀
日本航空宇宙学会論文集, 65, 4, 157, 167, 一般社団法人 日本航空宇宙学会, 2017年, [査読有り], [最終著者]
日本語, In this study, the authors conducted twice experiments to verify the throttling characteristics of axial-injection end-burning hybrid rockets. Oxidizer mass flow rate and chamber pressure were throttled by actuating valves in a fluid circuit consisting of two oxidizer supply lines. Chamber pressure and oxidizer mass flow rate were measured during each firing. The results show that oxidizer to fuel ratio remains constant for similar values of oxidizer mass flow rate. However, two weak points were identified in these throttling firing tests. First, a pressure transient was observed when oxidizer mass flow rate was increased (turn-up operation). The pressure transient consisted of two distinguishable first order lags, a fast lag followed by a slow lag, which are treated by separate curve fitting functions. The fast response is explained by a thermal lag in the solid fuel, whereas the slow response requires further inquiry. Second, the chamber pressure history exhibited hysteresis characteristics of oxidizer mass flow rate due to the increasing fuel regression rate. Therefore, in the throttling tests where oxidizer flow rate was turned-up and returned to the initial condition twice back-to-back, the chamber pressure history was higher in the second iteration than in the first. - Modified Regression Rate Formula of PMMA Combustion by a Single Plane Impinging Jet
Tsuneyoshi Matsuoka, Kyohei Kamei, Yuji Nakamura, Harunori Nagata
INTERNATIONAL JOURNAL OF AEROSPACE ENGINEERING, 2017, HINDAWI LTD, 2017年, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), A modified regression rate formula for the uppermost stage of CAMUI-type hybrid rocket motor is proposed in this study. Assuming a quasi-steady, one-dimensional, an energy balance against a control volume near the fuel surface is considered. Accordingly, the regression rate formula which can calculate the local regression rate by the quenching distance between the flame and the regression surface is derived. An experimental setup which simulates the combustion phenomenon involved in the uppermost stage of a CAMUI-type hybrid rocket motor was constructed and the burning tests with various flow velocities and impinging distances were performed. A PMMA slab of 20mm height, 60mm width, and 20mm thickness was chosen as a sample specimen and pure oxygen and O-2/N-2 mixture (50/50 vol.%) were employed as the oxidizers. The time-averaged regression rate along the fuel surface was measured by a laser displacement sensor. The quenching distance during the combustion event was also identified from the observation. The comparison between the purely experimental and calculated values showed good agreement, although a large systematic error was expected due to the difficulty in accurately identifying the quenching distance. - Development of Wall Regression Model of Hybrid Rocket Solid Fuel
NAWATA Kazuya, SASAKI Shunya, SAITO Tatsuya, OSHIMA Nobuyuki, WAKITA Masashi, TOTANI Tsuyoshi, NAGATA Harunori
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 14, ists30, 67, 72, 一般社団法人 日本航空宇宙学会, 2016年, [査読有り], [最終著者]
英語, 研究論文(学術雑誌),The purpose of this study is to develop a computational method to predict fuel regressions for hybrid rocket solid fuels. The shape of the flow field changes depending on the regression and vaporization of the solid fuel. This shape change and the heat flux from the combustion gas to the fuel are mutually dependent. Therefore a computational method that can accurately predicts both of the fuel regression and the heat flux is necessary to clarify the mutual dependence of them. In this study, we have developed a computational method to predict the regression phenomenon including the effect of the shape change of the flow field. The developed code predicts the regression phenomena by repeating gas-phase calculations and regression-phase calculations. The wall consisting of grids permits the flow field to be an arbitrary shape. As the first step, the complex chemical reaction was not included and numerical results were compared with a sublimation phenomenon of naphthalene in a non-combustion flow. Numerical results successfully predicted Nusselt number change due to regression qualitatively.
- Influence of Channel Width on Cylindrical Detonation Wave Propagation
KAMEYAMA Shota, NAWATA Shun, HIMONO Tsunetaro, KIKUCHI Keita, WAKITA Masashi, TOTANI Tsuyoshi, NAGATA Harunori
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 14, ists30, 39, 44, 一般社団法人 日本航空宇宙学会, 2016年, [査読有り], [最終著者]
英語, 研究論文(学術雑誌),To achieve stable detonation wave propagation to large-bore pulse detonation engine (PDE) combustors, we investigated an initiator for PDEs that uses a pre-detonator, reflector, and driver gas. In this initiator, a planar detonation wave from the pre-detonator becomes a cylindrical detonation wave after collision with the reflector. Wakita et al. previously posited two hypotheses regarding the dominant factors that determine the threshold of propagation to the target gas, which is a stoichiometric hydrogen–oxygen mixture diluted with nitrogen. This study reveals whether the threshold is determined by w/λ or λ/r. To analyze the effect of channel width w on the transition of the cylindrical detonation wave, experiments were conducted for w = 10 mm and w = 15 mm. However, the results could not elucidate whether the threshold is determined by λ/r or w/λ. To clearly distinguish between the effects of w/λ and λ/r, a narrow channel width w' = 3 mm was chosen and implemented using a torus-shaped obstacle. The results showed that a cylindrical detonation wave propagates without quenching when the nitrogen concentration is above 40%, corresponding to the cell size λ greater than w. Accordingly, the cylindrical detonation propagation threshold was determined to be independent of w/λ.
- Estimation of Hybrid Rocket Nozzle Throat Erosion History
Yuji SAITO, Tsutomu UEMATSU, Hikaru ISOCHI, Masashi WAKITA, Tsuyoshi TOTANI, Harunori NAGATA
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 14, ists30, Pa_145, Pa_151, 2016年, [査読有り], [最終著者]
英語, 研究論文(学術雑誌) - Heat Storage and Release Tests of Heat Storage Material with Crystal Transformation
Tsuyoshi TOTANI, Takuya KUNI, Toshifumi SATOH, Takuya ISONO, Masashi WAKITA, Harunori NAGATA
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 14, ists30, Pi_1, Pi_6, 2016年, [査読有り], [最終著者]
英語, 研究論文(学術雑誌) - 結晶転移により蓄熱する蓄熱材に及ぼす劣化の影響
國拓也, 戸谷剛, 佐藤敏文, 磯野拓也, 脇田督司, 永田晴紀
熱物性, 29, 4, 173, 178, 日本熱物性学会, 2015年11月, [査読有り], [最終著者]
日本語, 本研究では超小型人工衛星用蓄熱材料として,結晶転移によって蓄熱を行うトランス-1, 4-ポリブタジエン(TPBD)に注目した.同一のサンプルに対して蓄熱量の測定および分子構造の測定を行った.測定結果から,TPBDの分子構造内の-CH2の自動酸化によるC-H結合の変化が,蓄熱量に影響を与えていることが確認できた.連続使用の中での蓄熱量の変化を確認するために熱サイクル試験を行った.熱サイクル試験によって得られた温度履歴から蓄熱量及び放熱量の算出を行った結果,蓄熱量は60 J/gから90 J/gの間に,放熱量は-65 J/gから-100 J/gの間になることが分った.静止軌道衛星約4年半の運用期間に相当する1700回の蓄熱および放熱では蓄熱量の熱サイクル劣化は確認できなかった. - Thermal Analyses of Nano- and Micro- Satellites Pointing to the Earth with Deployable Solar Panel on Sun-synchronous Orbit by Small Number of Nodes
Tilok Kumar DAS, Tsuyoshi TOTANI, Masashi WAKITA, Harunori NAGATA
Mechanical Engineering Research Journal, 9, 79, 85, 2015年03月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌) - 大学によるロケット開発 : 経緯と展望(<特集>NPO大学宇宙工学コンソーシアムUNISECの軌跡と展望 第6回)
永田 晴紀, 湯浅 三郎, 和田 豊, 那賀川 一郎
日本航空宇宙学会誌, 63, 2, 51, 57, 一般社団法人日本航空宇宙学会, 2015年02月, [招待有り], [筆頭著者]
日本語, 大学宇宙工学コンソーシアム(UNISEC)設立当初のロケット関連団体の多くはハイブリッドロケット研究会のメンバーであり,この研究会の活動の一環として打上げられた機体が,我が国初のハイブリッドロケットとなった.2002年以降はUNISECを舞台に大学によるロケット関連活動は着実に広がり,打上げ実験が可能な場所も全国各地に整備された.非燃焼式ロケットや有翼実験機の開発等,活動内容の多様化も見られた.一方,大学研究室が開発したロケットは未だに宇宙に到達していない.自主開発ロケットによる宇宙到達は技術的ハードルが高く,UNISECにおけるロケット開発の延長上に宇宙をイメージするのは容易ではない.大学におけるロケット関連活動を活性化させ,ロケットが本来持っている若手技術者への強い訴求力を生かすためには,UNISECにおけるロケット関連活動の延長上に宇宙をイメージさせるための新たな方策が望まれる. - Scale effect on solid fuel regression in CAMUI-type hybrid rocket motor
Harunori Nagata, Mitsunori Ito
Progress in Scale Modeling, Volume II: Selections from the International Symposia on Scale Modeling, ISSM VI (2009) and ISSM VII (2013), 2, 249, 263, Springer International Publishing, 2015年01月01日, [査読有り], [招待有り], [筆頭著者]
英語, 論文集(書籍)内論文, The objective of this study is to obtain a rule to define a similarity condition under which subscale tests should be conducted to simulate firings of full-scale CAMUI-type hybrid motors. Static firing tests with fuel grains of different scaling have estimated the validity of similarity conditions based on convective heat transfer mechanisms. Fuel grains of all scales consist of four cylindrical polyethylene blocks with two axial ports. Experimental results show that except the fore-end face of the uppermost block and the back-end face of the rearmost block, similarity conditions based on convective heat transfer mechanisms are valid on end faces of fuel blocks. Because there is no end face downstream of the rearmost block, the flow field between fuel blocks with intense turbulence does not exist near the back-end face of the block, resulting in a small convective heat transfer rate. As a result, radiative heat transfer is not negligible on this burning surface and causes an error in the similarity condition. Because the impinging jet onto the fore-end face of the uppermost block is not high-temperature combustion gas but virtually pure oxygen, a similarity about chemical reaction is necessary in addition to those about convective heat transfer to realize a similarity condition. - Analytical Method for Prediction of Suction Performance of Ejector-Jet
Kouichiro Tani, Susumu Hasegawa, Shuichi Ueda, Takeshi Kanda, Harunori Nagata
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 58, 4, 228, 236, JAPAN SOC AERONAUT SPACE SCI, 2015年, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), To reduce the cost of space transportation, air-breathing engines are considered to be candidates for propulsion. However, to cover a wide range of flight speeds, the propulsion system has to operate in various modes to be efficient under incoming atmospheric-air conditions. The Japan Aerospace Exploration Agency is proposing a rocket-based combined cycle engine for operation under various condition, an ejector-jet mode being adopted for the low-speed regime. The suction performance ejector-jets has long been studied experimentally and numerically at JAXA, and little success has been achieved in explaining the deterioration of suction performance with high-temperature gas or light gas such as helium. In the present study, based on former models, a simple one-dimensional model was introduced incorporating the mixing effects of the primary flow (rocket flow) and secondary flow (induced air flow). The results were compared using several experimental and numerical data to check the plausibility of the model. It was found that if greater mixing occurs, suction performance is degraded, explaining the actual phenomena of the experiments. - Accuracy and applicable range of a reconstruction technique for hybrid rockets
Harunori Nagata, Hisahiro Nakayama, Mikio Watanabe, Masashi Wakita, Tsuyoshi Totani
Advances in Aircraft and Spacecraft Science, 1, 3, 273, 289, Techno Press, 2014年07月01日, [査読有り], [筆頭著者]
英語, 研究論文(学術雑誌), Accuracy of a reconstruction technique assuming a constant characteristic exhaust velocity (c*) efficiency for reducing hybrid rocket firing test data was examined experimentally. To avoid the difficulty arising from a number of complex chemical equilibrium calculations, a simple approximate expression of theoretical c* as a function of the oxidizer to fuel ratio (ξ ) and the chamber pressure was developed. A series of static firing tests with the same test conditions except burning duration revealed that the error in the calculated fuel consumption decreases with increasing firing duration, showing that the error mainly comes from the ignition and shutdown transients. The present reconstruction technique obtains ξ by solving an equation between theoretical and experimental c* values. A difficulty arises when multiple solutions of ξ exists. In the PMMA-LOX combination, a ξ range of 0.6 to 1.0 corresponds to this case. The definition of c* efficiency necessary to be used in this reconstruction technique is different from a c* efficiency obtained by a general method. Because the c* efficiency obtained by average chamber pressure and ξ includes the c* loss due to the ξ shift, it can be below unity even when the combustion gas keeps complete mixing and chemical equilibrium during the entire period of a firing. Therefore, the c* efficiency obtained in the present reconstruction technique is superior to the c* efficiency obtained by the general method to evaluate the degree of completion of the mixing and chemical reaction in the combustion chamber. - Thermal Design Procedure for Micro- and Nanosatellites Pointing to Earth
Tsuyoshi Totani, Hiroto Ogawa, Ryota Inoue, Tilok K. Das, Masashi Wakita, Harunori Nagata
JOURNAL OF THERMOPHYSICS AND HEAT TRANSFER, 28, 3, 524, 533, AMER INST AERONAUTICS ASTRONAUTICS, 2014年07月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), This paper proposes a thermal design procedure for micro- and nanosatellites that can be completed in one year. Two thermal design concepts keep components within their design temperature range, reducing the temperature change by using the whole structure for heat storage and reducing the temperature change of the inner structure where the most temperature-sensitive components are mounted. One- and two-nodal analysis methods are used for the former and latter concepts, respectively, to clarify the combinations of optical properties for the structures and components to keep within the design temperature range of the components. Finally, multinodal analysis is performed for detail design based on the optical properties clarified from the one- and two-nodal analyses. This thermal design procedure was applied to the Hodoyoshi-1 satellite, which is a cube about 50cm on a side, has two inner plates and has solar cells on the body, is on a sun-synchronous orbit at an altitude of about 500km, and is pointing to Earth. The thermal design of the Hodoyoshi-1 satellite was completed in about 10 months. - Influence of detonation cell size on propagation of cylindrical detonation wave
Masashi WAKITA, Kazuya SAJIKI, Tsunetaro HIMONO, Tsuyoshi TOTANI, Harunori NAGATA
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 12, ists29, Pa_1, Pa_7, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 2014年04月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), To achieve reliable transmission of detonation wave to a pulse detonation engine (PDE) combustor, authors examined a combination method of "predetonator", "reflector" and "overfilling of the driver gas" experimentally. A detonation wave propagates around our reflector changing its shape through three transition processes; from planer to cylindrical, toroidal, and planar again. Here, successful transmission to self-sustainable expanding cylindrical detonation wave is key issue. The authors used high sensitivity driver gas mixture (stoichiometric H2-O2 mixture) for the center of the cylindrical part to make the cylindrical detonation wave transmit in target gas mixture easily. To generalize the influence of the target gas composition on the necessary overfilling radius of the driver gas mixture, we employ stoichiometric H2-O2 mixture diluted by nitrogen or argon as target gas mixture. In this study, we showed that the ration of width of the cylindrical path on cell size of propagation limits of both dilution cases are about 1 when the driver gas is supplied enough to create a stable cylindrical detonation wave over 50 mm. Accordingly, when the cell size of the target gas mixture becomes over comparable size to the width of the cylindrical path, the stable expanding cylindrical detonation wave does not sustain. - Heat Storage Material without Phase-change for Micro and Nano Satellite
TOTANI Tsuyoshi, SATOH Toshifumi, WAKITA Masashi, NAGATA Harunori
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 12, 29, Po_4_1, Po_4_5, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 2014年, [査読有り], [最終著者]
英語, The thermal analysis of a micro cubic satellite pointing to the Earth on a sun-synchronous and circular orbit has been carried out using one-nodal analysis. The altitude of the orbit is 500 km. The local time of descending node of the orbit is 11 AM. The combination of the solar absorptivity and the infrared emissivity on the surface of the satellite under which the temperature of the satellite is kept within the allowable temperature range, from 0 to 40 degree Celsius, has been clarified. As the heat capacity is larger, the number of the combinations of the solar absorptivity and the infrared emissivity increases. In order to increase the heat capacity of nano and micro satellites, the development of a heat storage material has been performed. It is desirable that the heat storage materials for micro and nano satellites have the characteristic of not phase- change but crystal transformation at heat storage because a container for heat storage material is not required. Trans-1,4- polybutadiene transforms crystal structure at the temperature of heat storage. Trans-1,4-polybutadiene is produced and the heat storage performance is measured. The produced trans-1,4-polybutadiene has the amount of heat storage of about 80 J/g at the heat storage temperature of 74 deg. C. This amount corresponds to about 70% amount of heat storage of a literature data (112 kJ/kg).The density of the produced trans-1,4-polybutadiene is 706 kg/m3. - New Procedure for Thermal Design of Micro- and Nano-satellites Pointing to Earth
TOTANI Tsuyoshi, INOUE Ryota, OGAWA Hiroto, Kumar DAS Tilok, WAKITA Masashi, NAGATA Harunori
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 12, 29, Pf_11, Pf_20, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 2014年, [査読有り], [最終著者]
A new procedure for the thermal design of micro- and nano-satellites is proposed for completing the thermal design of micro- and nano-satellites within about one year. First, two concepts of thermal design are considered for maintaining the temperature change of units within an allowable range. One concept involves decreasing the temperature change of units by using the whole thermal capacity of the micro- and nano-satellite. The other concept involves decreasing the temperature change of the inner structure on which units with a narrow allowable temperature range are mounted and which is insulated conductively from the outer structure. Then, the temperatures of micro- and nano-satellites designed with the former concept are calculated using a one-node analysis method. The temperatures of micro- and nano-satellites designed with the latter concept are calculated using a two-node analysis method. The combinations of optical properties of the structures and units to maintain the temperature of units within the allowable range are obtained by using one- or two-node analysis. Finally, the multinode analyses are carried out to obtain a detailed design based on the optical properties obtained from the one-node analysis or two-node analysis. This thermal design procedure is applied to the Hodoyoshi-1 satellite, which is about 50 cm wide, 50 cm deep, 50 cm high, has a mass of about 50 kg, two inner plates, and solar cells on the body, flies on the Sun-synchronous orbit at the altitude of 500 km, and is pointing to the Earth. The thermal design of this micro-satellite was completed within about ten months. Possible problems with the procedure are tested, and the procedure is verified. - Effect of combustion pressure on regression rate of solid fuel under an impinging oxidizer jet counterflow diffusion flame
Ken Terakawa, Tatsuya Saito, Yuji Nakamura, Tsuneyoshi Matsuoka, Harunori Nagata, Tsuyoshi Totani, Masashi Wakita
JOURNAL OF THERMAL SCIENCE AND TECHNOLOGY, 9, 2, JTST0010, JTST0010, JAPAN SOC MECHANICAL ENGINEERS, 2014年, [査読有り]
英語, 研究論文(学術雑誌), Flame spread and counterflow diffusion flame experiments are widely conducted to investigate the combustibility of solid fuels. Although the use of the gas phase Damkohler number to organize the flame spread rate or regression rate of a solid fuel is effective under constant pressure, some research point out the possibility that the combustion pressure may be an independent factor in determining the regression rate. This research employs a counterflow diffusion flame to investigate the effects of combustion pressure on regression rate, and clarifies the deviation of results using the classical Damkohler number under varying pressures. First, a numerical flow analysis was conducted to determine the oxidizer velocity gradient near the fuel surface, which is an essential factor in evaluating the non-dimensional regression rate. Next, using an enclosed combustion chamber with independently variable oxidizer flux and pressure, experiments with a quasi two-dimensional flame were conducted with polyethylene solid fuel and nitrogen diluted oxygen oxidizer, and the regression rate was measured for two experiment series, constant pressure, and constant oxidizer flux. By comparing the two series, the effect of pressure on non-dimensionalized regression rate is clarified. The results suggest that contrary to the theoretical reaction rate of the gas phase, the non-dimensional regression rate increases when the combustion pressure is decreased, even in the thermal regime. This suggests that the classic method of organizing the regression rate with Damkohler number in thermal regime could not be implemented with varying pressure conditions, possibly due to the change in diffusion rates involved with varying pressures. - One Nodal Thermal Analysis for Nano and Micro Satellites on Sun-Synchronous and Circular Orbits
Tsuyoshi TOTANI, Hiroto OGAWA, Ryota INOUE, Masashi WAKITA, Harunori NAGATA
Transactions of the Japan Society for Aeronautical and Space Sciences, Aerospace Technology Japan, 11, 71, 78, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 2013年08月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), One nodal thermal analysis of nano and micro cubic satellites pointing to the Earth on sun-synchronous and circular orbits is carried out. The altitudes of the orbits are from 300 to 1,000 km. The local time of descending node is from 6 to 12. The combinations of the solar absorptivity and the infrared emissivity on the surface of the satellite in which the satellite satisfies the allowable temperature range, from 0 to 40 deg. C., are clarified for each of the above orbits. As the parameter of heat capacity of the satellite over one surface area of the satellite increases, the choice of combinations of the solar absorptivity and the infrared emissivity increases. The number of combinations in the case of the orbits without the shadow region is much larger than that with the shadow region. The number of combinations in the orbits without the shadow region increases with higher altitude and larger projected area with respect to the sun. The number of combinations in the orbits with the shadow region increases with higher altitude, larger projected area and smaller angle of the shadow region. - 固体燃料管内を燃え拡がる火炎の幾何学的相似条件
松岡 常吉, 永田 晴紀, 中村 祐二
実験力学, 13, 2, 178, 184, The Japanese Society for Experimental Mechanics, 2013年06月, [査読有り]
日本語, A number of studies about flame spread over solid combustibles have been reported over the past few decades. However, there are only a few studies about flame spread in a cylindrical enclosure. Especially, the static characteristics, such as flame height and width, have never been clarified. To reveal a geometrical similarity condition of flame spread in tube, we performed experiments in various conditions. The flame height increased as inner diameter, ambient pressure and oxidizer velocity increased. The non-dimensional flame height, determined as the flame height divided by the inner diameter, was proportional to Reynolds number. The flame spreading in different directions were compared and the shapes were almost similar to each other. It suggested that gravity effect on the flame shape was small. Meanwhile, the flame width did not depend on the flow field, and thus a quenching distance between the flame and the fuel surface was constant. From these results, it was found that the geometrical similarity condition was given by Reynolds number. In addition, considering the heat transfer from the flame to unburned fuel surface, we showed the similarity condition of the dynamic characteristic, i.e., non-dimensional flame spread rate. Also, the condition which can simultaneously satisfy both the static and the dynamic similarity was discussed in this paper. - CAMUI型ハイブリッドロケットの10年(<小特集>身近に感じられる宇宙開発)
永田 晴紀
日本機械学會誌, 116, 1134, 323, 326, 一般社団法人日本機械学会, 2013年05月, [招待有り], [筆頭著者]
日本語 - 事例から学ぶ燃焼物理 IV ハイブリッドロケット起動時に発生する異常燃焼の機構解明
永田晴紀, 飯島直純, 金井竜一朗
日本燃焼学会誌, 54, 170, 251-258, 258, 日本燃焼学会, 2012年11月, [招待有り], [筆頭著者]
日本語, During large scale CAMUI-type motor development, the authors frequently encountered anomalous combustion, a sudden pressure increase leading to destroy of the motor. Repeated static firing tests finally revealed that the cause of the anomalous combustion is the low initial fuel temperature. However, the mechanism responsible for the anomalous combustion is still unclear. Although a series of firing tests with a small combustor could not reproduce the anomalous combustion successfully, results showed a clear correlation between the initial fuel temperature and chamber pressure overshoot; chamber pressure overshoot does not occur when the fuel temperature is above the Leidenfrost point. From this result, the authors offer a hypothesis that the low fuel temperature below the Leidenfrost point enhanced heat transfer from the fuel to liquid oxygen and caused local blowoff. Accumulation of combustible mixture follows the blowoff and may cause the anomalous combustion. A possible reason why the firing tests could not reproduce the anomalous combustion is that Damkohler number in the combustion chamber was larger than those in the large-scale CAMUI-type motors. A preliminary experiment showed that the small combustor could reproduce the anomalous combustion by decreasing the Damkohler number in the combustion chamber. Detailed experimental study will follow the preliminary experiment to clarify the mechanism of the anomalous combustion. - Transition characteristics of combustion modes for flame spread in solid fuel tube
Tsuneyoshi Matsuoka, Shota Murakami, Harunori Nagata
COMBUSTION AND FLAME, 159, 7, 2466, 2473, ELSEVIER SCIENCE INC, 2012年07月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), This paper provides a new concept based on the Damkohler number (Da) to describe the complete transition behavior found in a flame spread in a solid combustible tube. Through a series of experiments performed with various diameters of the tube, ambient pressure, and oxidizer velocity within a wide range, three combustion modes are observed for the flame spread in a solid fuel tube namely combustion dominated by heat transfer (mode 1), by chemical kinetics (mode 2), and slow combustion sustained under very high blowing conditions (so-called "stabilized combustion": mode 3). Previous studies on the flame spread in tubes have shown that each transition, from model to mode 2 (transition 1-2) and from mode 2 to mode 3 (transition 2-3), is characterized by an equivalent velocity and by a friction velocity respectively. Meanwhile, for a flame spread on a fuel plate, it is widely known that both transitions are summarized by the Da. To achieve a comprehensive understanding of the transition characteristics of the combustion modes for the flame spread in the tube, the flame-spread rates under various conditions are experimentally investigated to elucidate the parameters that determine both transitions. First, the authors introduce a laminar friction velocity for the laminar flow region and revealed that transition 2-3 is determined by the laminar and turbulent friction velocity for laminar flow and turbulent flow regime respectively. The correlation between the Da and the friction velocity was experimentally obtained to show that transition 2-3 is consequently determined by the Da. This finding suggests that transition 2-3 corresponds to a blow-off limit that is observed for flame spread on a fuel plate. Second, the same correlation between the non-dimensional flame-spread rate and the Da is obtained, and it clearly showed that the transition 1-2 was determined by the Da. In conclusion, both transition phenomena are physically identical to those observed for on-plate flame spread, except the transition 2-3 occurs instead of the blow-off. (C) 2012 The Combustion Institute. Published by Elsevier Inc. All rights reserved. - Development of Pulse Detonation Engine Initiator Using Reflector for Large Bore Combustor
WAKITA Masashi, TAMURA Masayoshi, TERASAKA Akihiro, SAJIKI Kazuya, TOTANI Tsuyoshi, NAGATA Harunori
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 10, ists28, Pa_31, Pa_36, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 2012年, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), To achieve reliable transmission of detonation wave to a pulse detonation engine (PDE) combustor, authors have proposed a PDE initiator, which consists of a predetonator and a reflector. A detonation wave propagates around the reflector changing its shape through three transition processes; from planer to cylindrical, toroidal, and planar again. Our previous study revealed that the transition to the cylindrical detonation wave upstream of the board plays a significant role in detonating hydrogen-air mixture in a 100-mm-diam-combustor. A self-sustainable condition of the cylindrical detonation wave is severe when the radius of the wave front is small. In cases using hydrogen-oxygen mixture as driver gas for the 100-mm-diam-combustor, we had to fulfill with driver gas entire upstream of the board at the critical condition for the transition to the cylindrical wave. On the other hand, curvature of the cylindrical detonation wave front becomes smaller with increasing radius of the front, so the self-sustainable condition of the cylindrical wave must be mitigated for a large bore combustor. In this study, we investigated the necessary filling diameter of the driver gas to detonate hydrogen-air cylindrical detonation by using a 500-mm-diam-cylindrical-combustor. - CAMUI Type Hybrid Rocket as Small Scale Ballistic Flight Testbed
NAGATA Harunori, UEMATSU Tsutomu, ITO Kenichi
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 10, ists28, To_1_1, To_1_5, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 2012年, [査読有り], [筆頭著者]
英語, 研究論文(学術雑誌), The authors have been developing CAMUI (Cascaded Multistage Impinging-jet) type hybrid rockets, explosive-flee small rocket motors. This is to downsize the scale of suborbital flight experiments on space related technology development. A key idea is a new fuel grain design to increase gasification rates of a solid fuels. By the new fuel grain design, the combustion gas repeatedly impinges on fuel surfaces to hasten the heat transfer to the fuel. Suborbital flight experiments by sounding rockets provide variety of test beds to accumulate basic technologies common to the next step of space development in Japan. By using hybrid rockets one can take the cost advantage of small-scale rocket experiments. This cost advantage improves robustness of space technology development projects by dispersion of risk. - Automatic Circulation Control of Working Fluid in Liquid Droplet Radiator
TOTANI Tsuyoshi, TAKEKOSHI Takuhiro, WAKITA Masashi, NAGATA Harunori
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 10, ists28, Pf_1, Pf_8, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 2012年, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), Liquid Droplet Radiator (LDR) is an important candidate for disposing large quantities of waste heat more than 1 MW which will be handled by large space structures such as Space Power Satellite. The working fluid is heated through a heat exchanger by the waste heat generated in a large structure in space. Then, the working fluid is emitted in space through nozzles of the droplet generator toward a droplet collector as multiple streams of droplets. During the flight in space, the droplets lose thermal energy via radiative heat transfer. After the cooled droplets are captured by the droplet collector, the working fluid is recycled to the heat exchanger by a circulating pump. The automatic control system on the circulation of working fluid in a liquid droplet radiator has been built using a programmable logic controller. The proportional control of flow rate with the term of the variation of the counter flow in the gear pump and the relaxation of the change of an target flow rate has succeeded within 5 percent at the automatic circulation control of the working fluid from 100 ml/min to 200 ml/min and from 200 ml/min to 100 ml/min. - Effects of Heat Transfer in Divergent Section of Laval Nozzle on Exhaust Velocity and Area Ratio
Yuki Iwaki, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 54, 185-86, 212, 220, JAPAN SOC AERONAUT SPACE SCI, 2011年11月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), The effects of heating or cooling of the supersonic flow in a Laval nozzle have been investigated numerically. We focus on the exhaust velocity and the area ratio at given expansion ratios, which are ranged from 30 to 16,000. This range is equivalent to the area ratio from 4 to 400 at the specific heat ratio of 1.3 under isentropic expansion. Two types of heat profile are considered: pulsed heat transfer (PHT) and distributed heat transfer (DHT). The relations of Rayleigh flow and isentropic expansion are used for PHT. The exhaust velocity is higher than the isentropic value for the case where heat is provided near the throat. In other cases, the exhaust velocity is less than the isentropic value. The equivalent point of heat transfer is introduced for DHT. Using this equivalent point, the results for DHT exhibit the same trend as the results for PHT. This indicates that the effects of DHT can be predicted directly from results for PHT without numerical analyses. - 遺伝的アルゴリズムを用いたCAMUI型燃料グレインの最適設計
野原正寛, 金子雄大, 萩原俊輔, 永田晴紀
日本機械学会論文集 B編(Web), 77, 777, WEB ONLY 1249-1258, 1258, 一般社団法人 日本機械学会, 2011年, [査読有り], [最終著者]
日本語, The authors have been developing Cascaded Multistage Impinging-jet (CAMUI) type hybrid rockets. A CAMUI type hybrid rocket uses a fuel grain consisting of several cylindrical fuel blocks with two ports. To minimize both of c * loss due to O/F shifting and residual fuel weight after burning, an appropriate design of initial fuel grain shape is necessary. However, obtaining an optimum design of initial grain shape is not easy because there are many design variables influencing one another. To solve this problem, the authors employed Genetic Algorithm (GA) combined with a numerical model forecasting performance history of a CAMUI type motor. GA can acquire an approximate optimum solution for problems with a vast search space in practical time. The numerical model gives c * loss and residual fuel weight to evaluate the degree of performance of each initial fuel grain shape. A fuel grain design proposed by this method showed residual fuel weight as small as 6.36% of the initial weight and c * loss less than 1%. © 2011 The Japan Society of Mechanical Engineers. - Driver Gas Reduction Effect of Pulse-Detonation-Engine Initiator Using Reflecting Board
Masashi Wakita, Ryusuke Numakura, Takatoshi Asada, Masayoshi Tamura, Tsuyoshi Totani, Harunori Nagata
JOURNAL OF PROPULSION AND POWER, 27, 1, 162, 170, AMER INST AERONAUT ASTRONAUT, 2011年01月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), To reduce driver gas usage of a pulse detonation engine operating in airbreathing mode, the authors experimentally examined a combination method of a reflecting board and overfilling of the driver gas. This method has the potential to reduce the predetonator diameter by half and shorten the overfilling distance It to the reflecting board position w. Experiments with stoichiometric hydrogen-oxygen and hydrogen-air mixtures as driver and target gases, respectively, showed that the overfilling distance necessary to have a planar detonation wave propagate in a detonation chamber is reduced to 30 mm when a reflecting board is used with a reflecting board clearance of w = 10 mm. With an overfilling distance of 30 mm, the transformation of the detonation wave from cylindrical to toroidal did not occur because of the mixing effect of the driver gas and the target gas around the reflecting board. A 100-mm-thick reflecting board prevents the mixing effect, and a successful transformation from cylindrical to toroidal becomes possible with an overfilling distance as small as 17.2 mm. - Combustion characteristics of the end burning hybrid rockets in laminar flow
Tsuneyoshi Matsuoka, Harunori Nagata
ACTA ASTRONAUTICA, 68, 1-2, 197, 203, PERGAMON-ELSEVIER SCIENCE LTD, 2011年01月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), In this study, we aim to clarify the blowoff mechanism for flame spreading in an opposed laminar flow in narrow solid fuel ducts. To clarify this mechanism we conducted two experiments. First, we observed the changes of the flame spread rate at various oxygen velocities, ambient pressures, and port diameters. For flame spreading in laminar flow, combustion modes could be classified into 3 distinct regimes based on the strength of the opposed flow, i.e., chemical regime, thermal regime, and stabilized regime. This result is consistent with the result in turbulent flow. In the stabilized regime, quenching distance is almost constant despite oxygen velocity. In order to investigate the effect of ambient pressure and port diameter of fuels on blowoff limit, transition oxygen velocity is observed. As a result, transition oxygen velocity is proportional to the logarithm of the ambient pressure and port diameter. This relation is applicable despite the flow condition. Furthermore, we calculated velocity gradient at the fuel surface to reveal the determining factor of the blowoff limit in laminar flow. Consequently, velocity gradient, which is considered to dominate flow separation in laminar flow, would not be constant. This is because the velocity gradient at the fuel surface could not be evaluated by only the assumption of Hagen-Poiseuille flow but other parameters, such as vaporized fuel gas and natural convection by buoyancy should be included. (C) 2010 Elsevier Ltd. All rights reserved. - 紫外線硬化樹脂に金属膜を設けた表面微細周期構造の光学特性
石川 直幸, 戸谷 剛, 脇田 督司, 永田 晴紀
Thermophysical properties, 31, 0, 118, 120, 2010年11月, [査読有り], [最終著者]
日本語 - The Effect of Fuel Grain Size on the Combustion Characteristics in the Primary Combustion Chamber of Staged Combustion Hybrid Rocket
Harunori NAGATA, Kenta HASHIBA, Hiroya SAKAI, Tsuyoshi TOTANI, Masashi WAKITA
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 8, ists27, Pa7, Pa11, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 2010年07月, [査読有り], [筆頭著者]
英語, 研究論文(学術雑誌), To clarify the fuel gasification characteristics in a primary combustion chamber of a staged combustion hybrid rocket, the effect of fuel grain size on the regression rate of a grain was investigated experimentally. The grain size distribution in the combustion region achieved a steady state in 30 seconds burning duration. Examining fuel size distributions and fuel consumption rate at steady states, we obtained a history of fuel size and the regression rate of a grain in the combustion region. Regression rate increases with decreasing grain size. With a constant oxidizer flow rate, the regression rate is a function of grain size and independent to the initial grain size. After an initial transient the grain size decreases following the classical d-square law in droplet combustion: The square of the grain size decreases linearly with time. Although why the regression history of a grain in the combustion region follows the d-square law is not clear, this result is useful to estimate the fuel gasification rate of a staged combustion hybrid rocket. - 3D測定を応用したハイブリッドロケット燃焼器の複雑形状に対する熱流動解析(<小特集>マルチフィジックスCFD/EFDの最前線)
岸田 耕一, 金子 雄大, 大島 伸行, 永田 晴紀
日本機械学會論文集. B編, 76, 765, 789, 794, 一般社団法人日本機械学会, 2010年, [査読有り], [最終著者]
日本語, This paper investigates a thermal-fluid dynamics of CAMUI (Cascaded Multistage Impinging-jet) type hybrid rocket developed in Hokkaido University by using a large eddy simulation of turbulence. The performance of the hybrid rocket is sensitive to the changing shape of its chamber. To clarify this effects, numerical simulations were conducted using measured shapes. The results show the flow structures such as impinging fountain flow depending on the shapes at different burning time. Thease structures generate the particular heat flux distributions on the surface. - Effect of Temporal Variations of Internal Ballistics on Fuel Regression Rate in the CAMUI Hybrid Rocket
KANEKO Yudai, KISHIDA Kouichi, OSHIMA Nobuyuki, NAKASHIMA Takuji, WAKITA Masashi, TOTANI Tsuyoshi, NAGATA Harunori
Journal of Space Engineering, 3, 1, 52, 65, The Japan Society of Mechanical Engineers, 2010年, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), In order to clarify the temporal variations of internal ballistics during firing in a cascaded multistage impinging-jet (CAMUI)-type fuel grain, observations of instantaneous flow fields by numerical analysis along with instantaneous grain geometries were conducted. Two static firing tests were conducted under the same conditions, with the exception of firing duration, to obtain the temporal shapes of fuel grains and the characteristics of the regression progress. Two numerical analyses were conducted using the initial and instantaneous geometries to observe internal flow fields. A pair of vortices that is formed near the circumference of the grain due to the change in direction of the flow from the wall jet to the port flow induces a fan-like regression distribution. Two wall jets collide with each other at the center of the grain, roll up and form a fountain-like flow and a pair of gap-scale vortices. These vortices cause an unequal regression distribution on the end faces. On the downstream-end face, the vortices enhance the local regression rate near the axis of the grain. On the forward-end face, in addition to the region near the axis of the grain, the local regression rate at the reattachment points of the vortices increases. These gap-scale vortices disappear as regression progresses because of the dilation of the clearance between the fuel blocks. As a result, the regression rate distributions on both end faces become nearly flat with the progress of fuel regression. - Preliminary Thermal Design of UNITEC-1
TOTANI Tsuyoshi, II Haruaki, WAKITA Masashi, NAGATA Harunori
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 8, ists27, Pf_1, Pf_6, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 2010年, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), UNITEC-1 is a nano-spacecraft that flies to Venus. The preliminary thermal design of UNITEC-1 has been carried out. It has been clear from 1 node analysis that slow tumbling of the rotation axis is effective to reduce the temperature variations. It is difficult that both temperatures of the worst-case cold condition and the worst-case hot condition are within the allowable temperature ranges. It is desirable to conduct the survival competition as soon as possible after UNITEC-1 separates from a rocket. It is recommended from this viewpoint to use the black Kapton on the external surface except solar cells. It is clarified from the multi nodes analyses using Thermal Desktop/ SINDA/FLUINT that the temperature of the transmitter changes from 14.8 to 21.0 degree Celsius in an operational sequence under the worst-case cold condition. The temperature change of the battery can be suppressed from 15.5 to 16.6 degree Celsius in an operational sequence under the worst-case cold condition using the insulator between the battery and the internal surface. The maximum difference of temperature occurs between UOBC3 and UOBC6, and is 2.8 K. This difference is enough small to conduct the survival competition under the equal condition. The minimum period for the mission is about 114 days after UNITEC-1 separates from a rocket. - ノズル加熱による推力および比推力の向上に関する評価方法の検討
長沼 哲史, 岩城 裕樹, 佐藤 峻哉, 戸谷 剛, 脇田 督司, 永田 晴紀
日本航空宇宙学会論文集 = Journal of the Japan Society for Aeronautical and Space Sciences, 58, 677, 171, 177, 社団法人 日本航空宇宙学会, 2010年, [査読有り], [最終著者]
日本語, A numerical analysis program is created to research effect of heat transfer for propellant flow in Laval nozzle and estimate improvements of thrust and specific impulse. Several types of gases are assumed as propellant. The energy ratio is defined as ratio of energy supplied to propellant by convective heat transfer to enthalpy of propellant at the inlet of nozzle. The energy ratio increases with elongating length of divergent nozzle, and finally becomes maximum value that depends on Prandtl number, propellant temperature and wall temperature at the inlet of nozzle. The conversion efficienc... - EFFECT OF JET VELOCITY ON SCALE EFFECT IN OXIDIZER IMPINGING REGION
Yudai Kaneko, Mitsunori Itoh, Massasi Wakita, Tsuyoshi Totani, Harunori Nagata
Advances in the Astronautical Sciences, 138, 629, +, UNIVELT INC, 2010年, [査読有り], [最終著者]
英語, 研究論文(国際会議プロシーディングス), Diffusion combustion in a stagnation point boundary layer of a gaseous oxygen jet over a solid fuel was investigated to clarify effects of jet velocity on a similarity condition of fuel regression rates. This combustion field simulates the upstream-end face of the uppermost fuel block of CAMUI type hybrid rocket fuel grain. Increasing the flow velocity from 5.5 m/s to 11.5 m/s caused an increase in the regression rate from 0.22 mm/s to 0.26 mm/s. This result shows that the chemical reaction effect is not negligible in oxidizer impinging region. - REGRESSION PROGRESS OF FUEL GRAIN IN CAMUI TYPE HYBRID ROCKET MOTOR
Harunori Nagata, Akihito Kakikura, Mitsunori Ito, Yudai Kaneko, Kazuhiro Mori, Kenta Ueshima, Tsutomu Uematsu, Tsuyoshi Totani
Advances in the Astronautical Sciences, 138, 611, +, UNIVELT INC, 2010年, [査読有り], [最終著者]
英語, 研究論文(国際会議プロシーディングス), Static firing tests clarified how the fuel flow rate varies with the progress of fuel regression in a 'cascaded multistage impinging-jet' (CAMUI) type hybrid rocket motor. The fuel gasification rate decreases with progressing fuel regression because of two causes. One is decreasing gas flow density in ports. The other is decreasing area of end faces. The fuel gasification rate decreases rapidly when end faces disappear. A simple model of the regression progress was proposed. Fuel grains collected after firing tests with various burning duration approved this model. The model serves as a foundation to develop regression formulas applicable to this unconventional type fuel grain. - Preliminary Design of Winged Experimental Rocket by University Consortium
WAKITA Masashi, YONEMOTO Koichi, AKIYAMA Tomoki, ASO Shigeru, KOHSETSU Yuji, NAGATA Harunori
Transactions of the Japan Society for Aeronautical and Space Sciences, Aerospace Technology Japan, 7, ists26, Tg_21-Tg_26, Tg_26, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 2009年, [査読有り], [最終著者]
英語, The project of Winged Experimental Rocket described here is a proposal by the alliance of universities (University Consortium) expanding and integrating the research activities of reusable space transportation system performed by individual universities, and is the proposal that aims at flight proof of the results of advanced research conducted by the universities and JAXA using the university-centered experimental launch systems. This paper verifies the validity of the winged experimental rocket by surveying the technical issues that should be demonstrated and by estimating the airframe scale, weight and finally the total cost. The development schedule of this project was set to five years, where two airframes of different scales will be developed to minimize the risks. A 1.5-meter-long airframe will be first manufactured and conduct flight tests in the third year to verify the design issues. Then, a 2.5-meter-long airframe will be finally developed and conduct a complete flight demonstration of various research issues in the fifth year. - Development of 90 kgf Class CAMUI Hybrid Rocket for a CanSat Experiment
NAGATA Harunori, UEMATSU Tsutomu, ITO Mitsunori, KAKIKURA Akihito, KANEKO Yudai, MORI Kazuhiro, MURAI Norikazu, SATO Tatsuhiro, MITSUHASHI Ryuichi, TOTANI Tsuyoshi
Transactions of the Japan Society for Aeronautical and Space Sciences, Aerospace Technology Japan, 7, ists26, Tu_1-Tu_5, Tu_5, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 2009年, [査読有り], [筆頭著者]
英語, A newly designed CAMUI hybrid rocket motor of 900 N (90 kgf) thrust class, CAMUI-90, was developed. It uses a combination of polyethylene and liquid oxygen as propellants. CAMUI hybrid rocket is an explosive-flee small rocket motor to realize a small launch system with low cost and flexibility. The motor produces a thrust of 900 N for four seconds, keeping the optimal characteristic exhaust velocity of the fuel-oxidizer combination (exceeding 1800 m/s). A main application of the CAMUI-90 motor is for a CanSat experiment. A launch vehicle employing CAMUI-90 motor, 120 mm in diameter and 3.05 m in length, accelerates a payload of 500 g to 140 m/s in four seconds and reaches to an altitude of about 1 km. The first launch of this vehicle was on December 2006. - Fuel Regression Rate Behavior of CAMUI Hybrid Rocket
KANEKO Yudai, ITOH Mitsunori, KAKIKURA Akihito, MORI Kazuhiro, UEJIMA Kenta, NAKASHIMA Takuji, WAKITA Masashi, TOTANI Tsuyoshi, OSHIMA Nobuyuki, NAGATA Harunori
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, SPACE TECHNOLOGY JAPAN, 7, ists26, Pa_77, Pa_80, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 2009年, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), A series of static firing tests was conducted to investigate the fuel regression characteristics of a Cascaded Multistage Impinging-jet (CAMUI) type hybrid rocket motor. A CAMUI type hybrid rocket uses the combination of liquid oxygen and a fuel grain made of polyethylene as a propellant. The collision distance divided by the port diameter, H/D, was varied to investigate the effect of the grain geometry on the fuel regression rate. As a result, the H/D geometry has little effect on the regression rate near the stagnation point, where the heat transfer coefficient is high. On the contrary, the fuel regression rate decreases near the circumference of the forward-end face and the backward-end face of fuel blocks. Besides the experimental approaches, a method of computational fluid dynamics clarified the heat transfer distribution on the grain surface with various H/D geometries. The calculation shows the decrease of the flow velocity due to the increase of H/D on the area where the fuel regression rate decreases with the increase of H/D. To estimate the exact fuel consumption, which is necessary to design a fuel grain, real-time measurement by an ultrasonic pulse-echo method was performed. - Thermal Design of a Solar Thermal Thruster for Piggyback Satellites
IWAKI Yuuki, TOTANI Tsuyoshi, NAGATA Harunori
TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, SPACE TECHNOLOGY JAPAN, 7, ists26, Pb_71, Pb_76, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 2009年, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), A method of thermal analysis for a solar thermal thruster was created to aid in the thermal design of the thruster. The method consists of two types of thermal analysis: an analysis program for propellant flow, and an analysis of the temperature distribution of the thruster wall using Pro/Engineer. The numerical results were compared with experimental results to confirm the validity of the method, and there was good agreement between them. A thermal design was created using this thermal analysis method to estimate the performance of a solar thermal thruster for the orbital transfer of piggyback satellites mounted on an H2A rocket. When the thruster is made from heat-resistant steel and the propellant is water, the analytical results showed that the Isp is 203 s, the thrust is 16.6 mN, and the maximum temperature of the thruster is 1088 K. The diameter of the concentrator also was calculated, and it was found to be small enough for the concentrator to be mounted on piggyback satellites. - 縦列多段衝突噴流(CAMUI)方式を用いたハイブリッドロケットの燃料後退特性
伊藤 光紀, 前田 剛典, 柿倉 彰仁, 金子 雄大, 森 一大, 中島 卓司, 脇田 督司, 植松 努, 戸谷 剛, 大島 伸行, 永田 晴紀
日本航空宇宙学会論文集 = Journal of the Japan Society for Aeronautical and Space Sciences, 55, 646, 516, 526, 社団法人 日本航空宇宙学会, 2007年11月, [査読有り], [最終著者]
日本語, A series of lab-scale firing tests was conducted to investigate the fuel regression characteristics of Cascaded Multistage Impinging-jet (CAMUI) type hybrid rocket. The alternative fuel grain used in this rocket consists of a number of cylindrical fuel blocks with two ports, which were aligned along the axis of the combustion chamber with a small gap. The ports are aligned staggered with respect to ones of neighboring blocks so that the combustion gas flow impinges on the forward-end surface of each block. In this fuel grain, forward-end surfaces, back-end surfaces and ports of fuel blocks ... - Detonation transition limit at an abrupt area change using a reflecting board
Masashi Wakita, Ryusuke Numakura, Yusuke Itoh, Shigetoshi Sugata, Tsuyoshi Totani, Harunori Nagata
JOURNAL OF PROPULSION AND POWER, 23, 2, 338, 344, AMER INST AERONAUTICS ASTRONAUTICS, 2007年03月, [査読有り], [最終著者]
英語, 研究論文(学術雑誌), To realize quick initiation of detonation in the combustion chamber of a pulse detonation engine operating in the, air-breathing mode, in which the combustible gas is a fuel-air mixture, the authors have proposed a new pulse detonation engine initiator using a "reflecting board" near the exit of a predetonator tube. In this study, we clarify the transition limit of this new initiator by examining the detonation cell size at the predetonator exit and the mechanism that gives this transition limit. The combustible mixtures are stoichiometric hydrogen-oxygen mixtures diluted with nitrogen or argon. The main results obtained in this study are as follows. When the incident detonation wave interacts with the reflecting board before it completely disappears due to the rarefaction wave from the predetonator exit, the number of cells between the exit and the board defines the transition limit from the planar to cylindrical detonation waves. Even when the cylindrical detonation does not occur, the reflecting board converts a planar detonation wave into a torus-shape pressure wave. This pressure wave encompasses the combustible gas in the detonation chamber and concentrate on the axis, causing a detonation bubble behind the board. The necessary minimum diameter of the predelonator with a reflecting board is expressed by D-c = 6.3 lambda. - Determining factor for the blowoff limit of a flame spreading in an opposed turbulent flow, in a narrow solid-fuel duct
Nozomu Hashimoto, Harunori Nagata, Tsuyoshi Totani, Isao Kudo
COMBUSTION AND FLAME, 147, 3, 222, 232, ELSEVIER SCIENCE INC, 2006年11月, [査読有り]
英語, 研究論文(学術雑誌), This study clarified the blowoff mechanism for a flame spreading in an opposed turbulent flow in narrow solid fuel ducts. To clarify this mechanism, two experiments were conducted. The first experiment was to investigate the influence of ambient pressure and fuel duct size on the blowoff limit. The results indicated that the flow velocity at the point when blowoff occurred, V-g,V-t, increased with ambient pressure. This tendency could not be confirmed by a well-known expression for the Damkohler number, which is defined as the ratio of the characteristic flow time to the characteristic chemical time. Subsequently, to clarify the determining factor for the blowoff, the second experiment, which observed the flow field near the flame leading edge, was conducted. The results show that the flow separation in front of the flame leading edge, which provided sufficient residence time of oxidizer and gaseous fuel, is necessary for the flame to spread in an opposed oxidizer flow. From the results, it is found that the oxidizer friction velocity, u(*), which is an indicator of the turbulent momentum transfer, is the determining factor for the flame blowoff limit. When the friction velocity is larger than a critical value, flame blowoff occurs in the fuel duct, due to the absence of flow separation. (c) 2006 The Combustion Institute. Published by Elsevier Inc. All rights reserved. - 異種混合気間を回折するデトネーション波に及ぼす反射板の伝播促進効果
沼倉 龍介, 脇田 督司, 伊藤 雄介, 菅田 成俊, 永田 晴紀, 戸谷 剛, 工藤 勲
日本燃焼学会誌 = Journal of the Combustion Society of Japan, 48, 145, 265, 272, 日本燃焼学会, 2006年08月, [査読有り]
日本語, To realize a quick initiation of detonation in insensitive fuel-air mixtures in the combustion chamber of a PDE operating in the air-breathing mode, the authors have proposed a new detonation initiator using a circular disk as a "reflecting board" near the exit of a predetonator tube. When a fuel-oxygen mixture fills the predetonator tube as a driver gas, the mixture change and the abrupt area change occur simultaneously at the exit. This paper describes the promoting effect of the reflecting board on the detonation transition through the mixture change. The combustible mixtures in the combustion chamber are stoichiometric hydrogen-oxygen mixtures diluted with nitrogen. Main results obtained in this study are in the followings: The detonation wave maintains the propagation velocity in the fuel-oxygen mixture right after the abrupt area change, this results in the increase of the distance at which the head of a rarefaction wave reaches the axis of the predetonator and the incident planar detonation wave disappears. Because the cell size right after the abrupt area change is near the size of the fuel-oxygen mixture, the cylindrical detonation wave survives the axial expansion even when the reflecting board separation is less than the propagation limit of the cylindrical detonation wave propagating in the diluted mixture. - Numerical and experimental studies on circulation of working fluid in liquid droplet radiator
T Totani, T Kodama, K Watanabe, K Nanbu, H Nagata, Kudo, I
ACTA ASTRONAUTICA, 59, 1-5, 192, 199, PERGAMON-ELSEVIER SCIENCE LTD, 2006年07月, [査読有り]
英語, 研究論文(学術雑誌), A model of the circulation of the working fluid in a liquid droplet radiator has been developed. The model is based on Bernoulli's law and the loss of the hydraulic head. The behavior of the circulation of the working fluid calculated from the model is compared with that obtained from experiments in the case that the flow rate of the circulating working fluid is changed. In radiators, the flow rate of the circulating working fluid is changed in order to match the change of the waste heat generated in large-space structures. The flow rates of the circulating working fluid calculated from the model correspond to those obtained from the experiments well. The circulation mechanism of the working fluid in the liquid droplet radiator has been clarified. The model developed in the present work will allow us to control the flow rate of the working fluid in the liquid droplet radiator automatically. (C) 2006 Elsevier Ltd. All rights reserved. - Development of CAMUI hybrid rocket to create a market for small rocket experiments
Harunori Nagata, Mitsunori Ito, Takenori Maeda, Mikio Watanabe, Tsutomu Uematsu, Tsuyoshi Totani, Isao Kudo
Acta Astronautica, 59, 1-5, 253, 258, PERGAMON-ELSEVIER SCIENCE LTD, 2006年07月, [査読有り], [筆頭著者]
英語, 研究論文(学術雑誌), By introducing various innovative ideas, the difficult-to-develop small hybrid-type rocket is successfully developed. The main purpose is to drastically reduce the cost of rocket experiments and thus, attract potential users such as metrological and microgravity researchers. A key idea is a new fuel grain design to accelerate the gasification rate of solid fuel. The new fuel grain design, designated as CAMUI as an abbreviation of "cascaded multistage impinging-jet", is that the gas flow repeatedly collides with the solid fuel surface to accelerate the heat transfer to the fuel. To install a regenerative cooling system using cryogenic liquid oxygen as coolant in a small launcher, the authors devised a valveless supply system (with no valves in the liquid oxygen flow line). Four serial successful launch verification tests by 10 kg vehicle equipped with a 50 kgf thrust CAMUI motor have shown the feasibility of the motor system. The meteorological observation model of 400 kgf class motor is under development and the development of microgravity experiment class of 1.5-2 tonf motor will follow subsequently. The authors plan to complete the development of the 400 kgf class motor for meteorological observation model by the end of FY2005. © 2006 Elsevier Ltd. All rights reserved. - 反射板を用いたPDEイニシエーターにおける爆轟波の再開始機構
脇田 督司, 沼倉 龍介, 伊藤 雄介, 永田 晴紀, 戸谷 剛, 工藤 勲
日本航空宇宙学会論文集 = Journal of the Japan Society for Aeronautical and Space Sciences, 53, 620, 414, 418, 社団法人 日本航空宇宙学会, 2005年09月05日
日本語, Quick initiation of a detonation wave in a combustion chamber is important to realize high-performance pulse detonation engine. A possible method is to generate a detonation wave in a pre-detonator and release the detonation wave into the chamber. In this paper, a reflecting board is installed in the combustion chamber near the pre-detonator exit where the tube diameter expands abruptly. It prevents the detonation wave from disappearing at the expanding region near the tube exit. The re-initiation mechanisms of a detonation wave near the reflecting board were observed by using the soot film... - CAMUIハイブリッドロケット内部の熱・流動特性の数値解析
渡辺 三樹生, 永田 晴紀, 戸谷 剛, 工藤 勲
日本航空宇宙学会論文集 = Journal of the Japan Society for Aeronautical and Space Sciences, 53, 618, 337, 342, 社団法人 日本航空宇宙学会, 2005年07月05日
日本語, The authors have proposed an advanced fuel configuration to overcome the defect of conventional hybrid rockets, i.e., the low thrust level. The key feature of this new type of hybrid rocket, named Cascaded Multi-staged Impinging jet (CAMUI), is that the cylindrical fuel blocks with two ports parallel to the axis are arranged in a row in the combustion chamber. This fuel configuration allows mixing and combustion to occur in and around the impinging jet regions. In the present paper, a CFD simulation clarifies the fundamental features of the flow field and the heat transfer distributions in ... - Thermal design of liquid droplet radiator for space solar-power system
T Totani, T Kodama, H Nagata, Kudo, I
JOURNAL OF SPACECRAFT AND ROCKETS, 42, 3, 493, 499, AMER INST AERONAUT ASTRONAUT, 2005年05月
英語, 研究論文(学術雑誌), The waste heat from the space solar-power system, which supplies 5 MW of electricity to a power transmission line on Earth, is estimated, and the liquid droplet radiator for handling the waste heat are examined on the basis of experimental results obtained under microgravity for droplet generation and droplet collection of the liquid droplet radiator. The following results have been obtained. First, an active heat removal system for the power generation unit in the photovoltaic power system is not necessary when the concentration ratio of solar energy is smaller than 1.34, whereas for the liquid droplet radiator, with silicon oil as working fluid, in the solar dynamic power system, the droplet sheet for radiating the waste heat must be 147 m long, 65.1 m wide, and 0.998 m thick. Second, the droplet sheet of the liquid droplet radiator, in which the working fluid is silicon oil, must be 107 m long, 43.2 m wide, and 0.998 m thick to manage the waste heat from the power distribution unit and the power transmission unit in the photovoltaic power system, whereas it must be 107 m long, 65.2 to wide, and 0.998 m thick in the solar dynamic power system. - 「反射板を用いたPDEイニシエーターにおける爆轟波の再開始機構」
脇田 督司, 沼倉 龍介, 伊藤 雄介, 永田 晴紀, 戸谷 剛, 工藤 勲
日本航空宇宙学会誌, 53, 620, 414, 418, 一般社団法人 日本航空宇宙学会, 2005年
日本語, Quick initiation of a detonation wave in a combustion chamber is important to realize high-performance pulse detonation engine. A possible method is to generate a detonation wave in a pre-detonator and release the detonation wave into the chamber. In this paper, a reflecting board is installed in the combustion chamber near the pre-detonator exit where the tube diameter expands abruptly. It prevents the detonation wave from disappearing at the expanding region near the tube exit. The re-initiation mechanisms of a detonation wave near the reflecting board were observed by using the soot film method. Main results obtained in this study are in the followings: Re-initiation of a detonation wave due to the Mach reflection of a shock wave is observed on the surface of the reflecting board and the propagation promoting effect is observed. The effectiveness of the reflecting board is a strong function of the clearance between the pre-detonator exit and the reflecting board, and the promotion effect sharply decreases with increasing the clearance beyond the distance, in which the incident planar detonation wave maintains. By equipping with a reflective board with a suitable clearance, the critical cell size increases by 2 or 3 times. - Experimental study on convergence of droplet streams under microgravity
T Totani, T Kodama, K Watanabe, H Nagata, Kudo, I
MICROGRAVITY SCIENCE AND TECHNOLOGY, 17, 3, 31, 38, Z A R M TECHNIK PUBLISHING DIV, 2005年
英語, 研究論文(学術雑誌), Experiments on the convergence of two droplet streams have been carried out under microgravity in order to develop a technique for converging droplet streams under microgravity and to examine the behavior of droplets in a vacuum and under microgravity after the binary droplets collide with each other The working fluid is silicone oil with a low vapor pressure. In this study, a method of orienting the droplet generators toward a con vergence point has been tested. In all of the 68 experiments conducted under microgravity, it is confirmed that droplet streams are converged. It has been concluded that the method of orienting multiple droplet generators to a converging point is effective for converging droplet streams under microgravity. The behaviors of the colliding droplets under microgravity and in a vacuum have been classified into five types. The five types of behavior are mapped on a We (Weber number) - B (impact parameter) diagram. The range of Weber numbers in the experiments is from 200 to more than 3000. - CAMUI 型(縦列多段衝突噴流型)ハイブリッドロケットの 開発と微小重カロケット実験への応用
永田 晴紀
日本マイクログラビティ応用学会誌, 22, 1, 47, 日本マイクログラビティ応用学会, 2005年
日本語, Small-scale reusable sounding rocket system is under development to provide three-minutes microgravity condition to a 10-kg payload. The propulsion system is a hybrid type that uses solid fuel (plastics) and liquid oxygen as propellants and free from explosives, resulting in the dramatically reduced launch cost. To enhance the burning rate of the solid fuel and to augment the thrust, the rocket has employed a new fuel grain design. This new design, named CAMUI as an abbreviation of "Cascaded Multistage Impinging-jet" , allows mixing and combustion to occur around stagnation points on fuel surfaces. Successful launch experiments using a 50-kgf GAMUI engine have proved the feasibility of the basic idea of the system. Finally, a possible configuration of the microgravity test vehicle is presented. - Current Status of Rocket Developments in Universities – Development of CAMUI Hybrid Rocket
Nagata, H, Itoh, M, Maeda, T, Kato, R, Totani, T, Kudo, I, Uematsu, T
Journal of Space Technology and Science, 21, 1, 31, 38, 2005年, [査読有り], [筆頭著者]
英語, 研究論文(学術雑誌) - 微小重力下における液滴流の収束および衝突挙動
戸谷 剛, 児玉 拓也, 渡辺 健介, 永田 晴紀, 工藤 勲
JASMA : Journal of the Japan Society of Microgravity Application = 日本マイクログラビティ応用学会誌, 21, 0, 14, 14, 2004年11月04日, [査読有り]
英語 - Totani, T., Nagata, H., Kudo, I. and Iwasaki, A.: "Measurement Technique for Pumping Performance of a Centrifugal Collector under Microgravity", Review of Scientific Instruments, 75(2): 515-523 (2004)*
2004年, [査読有り] - Relationship between platinum wire temperature and catalytic heat release rate on platinum in unsteady-state hydrogen-air mixture
Daisuke Nakamura, Harunori Nagata, Tsuyoshi Totani, Isao Kudo
Heat Transfer - Asian Research, 33, 1, 1, 11, 2004年01月, [査読有り]
英語, The authors have proposed a hydrogen concentration probe using a catalytic reaction on the surface of a platinum wire. To use this probe for detecting the concentration change in a supersonic mixing layer, the response of the catalytic heat release rate must depend only on the change of concentration around the probe. The catalytic heat release rate on the surface of the platinum wire in an unsteady state was measured by a constant-temperature hot-wire anemometer and a shock tube to investigate the relationship between the response of the catalytic heat release rate and the temperature of the platinum wire. The catalytic heat release rate began increasing upon the introduction of the shock wave. The rate of increase of catalytic heat release depended on the temperature of the platinum wire when the temperature of the hot wire was low. However, the dependence was very weak when the temperature of the hot wire was above 400 °C. This shows that it is not the catalytic reaction but rather molecular transfer from the flow to the surface of the platinum wire is the controlling step when the temperature of the platinum wire is high. In conclusion, the temperature of the platinum wire must be above 400 °C to use the hydrogen concentration probe in a supersonic mixing layer. © 2003 Wiley Periodicals, Inc. - 二段燃焼ハイブリッドロケットの先導研究
秋葉 鐐二郎, 青木 嘉範, 加勇田 清勇, 藤井 篤之, 永田 晴紀, 佐鳥 新
日本航空宇宙学会論文集 = Journal of the Japan Society for Aeronautical and Space Sciences, 51, 591, 141, 150, 社団法人 日本航空宇宙学会, 2003年04月05日
日本語, The staged combustion hybrid rocket is under development by our research group since 1999. This hybrid rocket engine consists of two combustion chambers. The primary combustion chamber is the fuel tank itself filled with granular solid fuels. The fuel rich gas generated by the first stage combustion flows into the secondary combustion chamber, which is located in the bottom of the primary combustion chamber. The additional oxidizer is injected to the secondary combustion chamber in order to attain an optimal specific impulse by completing combustion. There are two types of the primary combu... - 短時間微小重力実験による液滴ラジエータ要素の機能試験
戸谷 剛, 永田 晴紀, 工藤 勲
JASMA : Journal of the Japan Society of Microgravity Application, 20, 1, 22, 29, 日本マイクログラビティ応用学会, 2003年01月31日
日本語, The liquid droplet radiator (LDR) is an important candidate to resolve an technical issue that is how to reject of large quantities of waste heat from large structures in space, which handle high power (from megawatts to gigawatts). The performance of LDR elements under microgravity condition has been investigated. It has been clarified that (1) the diameters of droplets and spacing between droplets generated under microgravity can be formulated by the equations based on the law of conservation of mass in the process of generating droplets, (2) the uniform droplet stream is captured under microgravity without splashes, (3) the pumping performance of the working fluid of the centrifugal droplet collector under microgravity can be predicted by the sum of velocity head and pressure head generated in the gyrostatic flow, (4) the gear pump can also function normally under microgravity. - Hashimoto, N., Watanabe, S., Nagata, H., Totani, T., Kudo, I.:"Opposed-Flow Flame Spread in a Circular Duct of a Solid Fuel: Influence of Cannel Height on Spread Rate", Proceedings of the Combustion Institute, Vol.29, pp.245-250, 2003.*
2003年 - 「非定常場における水素-空気混合気の白金触媒反応発熱量と熱線温度との関係」
中村 大輔, 永田 晴紀, 戸谷 剛, 工藤 勲
日本機械学会論文集(B編), 69, 677, 126, 131, 一般社団法人 日本機械学会, 2003年
日本語, The authors have proposed a hydrogen concentration probe using catalytic reaction on Pt wire surface. To use this probe to detect a concentration change in a supersonic mixing layer, the response of the catalytic heat release rate must depend only on concentration change around the probe. Catalytic heat release rate on the Pt wire surface in unsteady state is measured using a constant temperature type hotwire anemometer technique and a shock tube to investigate the relation of the response of the catalytic heat release rate and Pt wire temperature. Catalytic heat release rate begins increasing at the arrival of the shock wave. The increasing rate of the catalytic heat release depends on the Pt wire temperature when the wire temperature is low. However, the dependence is very weak when the wire temperature is over about 680 K. This shows that not the catalytic reaction but molecular transfer from the flow to the Pt wire surface is the controlling step when Pt wire temperature is high enough. As a conclusion, the Pt wire temperature over about 680 K is necessary to use the hydrogen concentration probe in a supersonic mixing layer. - 「二段燃焼ハイブリッドロケットの先導研究(Leading Studies of the Staged Combustion Hybrid Rocket)」
秋葉りょう二郎, 青木嘉範, 加勇田清勇, 藤井篤之, 永田晴紀, 佐鳥新
日本航空宇宙学会論文集, 51, 591, 141, 150, 一般社団法人 日本航空宇宙学会, 2003年
日本語, The staged combustion hybrid rocket is under development by our research group since 1999. This hybrid rocket engine consists of two combustion chambers. The primary combustion chamber is the fuel tank itself filled with granular solid fuels. The fuel rich gas generated by the first stage combustion flows into the secondary combustion chamber, which is located in the bottom of the primary combustion chamber. The additional oxidizer is injected to the secondary combustion chamber in order to attain an optimal specific impulse by completing combustion. There are two types of the primary combustion. One is nicknamed as the incinerator type; the other is called the multi-grain type. This new type engine is featured with a wide range throttling capability and an extensive freedom in selecting the fuel material. This paper deals with the incinerator type. Presented are preliminary experiments as well as the systems description. - 非定常場における水素-空気混合気の白金触媒反応発熱量と熱線温度との関係(熱工学,内燃機関,動力など)
中村 大輔, 永田 晴紀, 戸谷 剛, 工藤 勲
日本機械学會論文集. B編, 69, 677, 126, 131, 一般社団法人日本機械学会, 2003年
日本語, 研究論文(学術雑誌), The authors have proposed a hydrogen concentration probe using catalytic reaction on Pt wire surface. To use this probe to detect a concentration change in a supersonic mixing layer, the response of the catalytic heat release rate must depend only on concentration change around the probe. Catalytic heat release rate on the Pt wire surface in unsteady state is measured using a constant temperature type hotwire anemometer technique and a shock tube to investigate the relation of the response of the catalytic heat release rate and Pt wire temperature. Catalytic heat release rate begins increas... - Opposed-flow flame spread in a circular duct of a solid fuel: Influence of channel height on spread rate
N Hashimoto, S Watanabe, H Nagata, T Totani, Kudo, I
PROCEEDINGS OF THE COMBUSTION INSTITUTE, 29, 245, 250, COMBUSTION INST, 2003年, [査読有り]
英語, 研究論文(学術雑誌), The influence of channel height on flame spread in a circular duct of the solid fuel in an opposed-flow configuration was examined. Polymethylmethacrylate cylinders with a circular duct (diameter of 1, 2, or 3 mm) were used as fuel specimens, and both flame-spreading and stabilized combustion were observed. In the case of stabilized combustion, the flame cannot spread into the duct because of the high oxygen velocity The flame-traveling velocity is the velocity at which the flame widens the duct by fuel consumption. Therefore, the flame-traveling velocity in stabilized combustion is significantly low compared with flame-spreading combustion. In the case of flame-spreading combustion, the equivalence velocity, which contains channel height information, defines whether the regime is the thermal or the chemical regime. When the equivalent velocity is higher than a certain value, the flame-spread rate is controlled by chemical effects. On the whole, the flame-spread rate decreases with the decrease of channel height in the case of flame-spreading combustion because of the curvature effect. Owing to the curvature effect, the area ratio of the flame to that of the solid surface decreases with decreasing channel height, and this is conspicuous when the channel height is low. The curvature effect is negligible when the channel height is sufficiently large compared with the flame stand-off distance. - 小型衛星のためのハイブリッドロケットの打ち上げ機の開発
渡辺 三樹生, 中山 久広, 永田 晴紀, 戸谷 剛, 工藤 勲, 伊藤 献一, 大和田 陽一
JASMA : Journal of the Japan Society of Microgravity Application, 19, 2, 112, 116, 日本マイクログラビティ応用学会, 2002年04月30日
日本語, For the use of ballistic launches of a small satellite, the development study of a high thrust hybrid rocket motor has been made. To enhance the regression rate of the solid fuel and augment the thrust, the authors employed a new fuel configuration. This new configuration allows mixing and combustion to occur around the stagnation points on solid fuel surfaces. The static firing tests using a LOX/PMMA hybrid motor with a LOX cooling system have proved sufficiently prompt ignition, high thrust level, and stable combustion. Based on the obtained data, a flight performance of a small hybrid rocket for a ballistic test launch was estimated. - インターネットを利用したコラボレーション支援環境の構築とその運用
吉川 茂雄, 戸谷 剛, 永田 晴紀
日本ディスタンスラーニング学会会誌, 3, 0, 19, 25, 日本ディスタンスラーニング学会, 2002年03月
日本語 - Flame shapes of fuel droplet cloud in high temperature gaseous environment under micro-gravity
H Enomoto, H Nagata, D Segawa, T Kadota
JSME INTERNATIONAL JOURNAL SERIES B-FLUIDS AND THERMAL ENGINEERING, 45, 1, 102, 107, JAPAN SOC MECHANICAL ENGINEERS, 2002年02月, [査読有り]
英語, 研究論文(学術雑誌), In order to investigate the spray combustion mechanism, a new methodology (Fine Wire Sustaining method) was established. Fine wires of 14mum in diameter were used to sustain the droplets. Any arrangement of the droplets could be performed with this method. In this study, 33 fuel droplets arranged in symmetrically were subjected to the quiescent high temperature air in an electric furnace. The temperature of the environment air was about 1000K. Fuel was n-eicosane and the mean droplet diameter was 0.58mm. The standard deviation of the droplet diameter was 0.02mm. A high-speed video camera of 250ftp was provided to observe the auto-ignition and flames of fuel droplet clouds. The experiments were done at atmospheric pressure using the JAMIC drop shaft that provides 10 seconds of effective period of time for the micro-gravity As the results, the time histories of the diameter of the particle flames had maximum and that of the diameter of the group flame had the minimum. - Nagata, H., Kudo, I., Ito, K., Nakamura, S., Takeshita, Y.:"Interactive Combustion of Two-dimensionally Arranged Quasi-droplet Clusters under Microgravity", Combustion and Flame, 129:392-400(2002)*
2002年 - 「液滴ラジエータ用遠心式液滴回収器の微小重力下での性能」
戸谷 剛, 伊丹 雅洋, 藪田 茂, 永田 晴紀, 工藤 勲, 岩崎 晃, 細川 俊介
『日本機械学会論文集(B編)』, 68, 674, 2780, 2787, 一般社団法人 日本機械学会, 2002年
日本語, The Liquid Droplet Radiator (LDR) has an advantage over conventional radiators in terms of the rejected heat power-weight ratio. LDR has been taken notice as an advanced radiator for high-power generation systems which will be prerequisite for large space structures. In this study, the performance of a centrifugal droplet collector under microgravity condition has been investigated from the viewpoint of operational space use of LDR in the future. It has been concluded that (1) a centrifugal collector is able to transport working fluid to a recirculating pump under microgravity condition
(2) the ability to pump working fluid is formulated as the sum of pressure head and velocity head generated in the centrifugal collector where the velocity is c (0 <
c ≤ 1) times as fast as in the rigid rotational flow
(3) splashing of the working fluid occurs at that position, when working fluid strikes against part of the entrance of the pitot tube on the centrifugal collector. - Totani, T., Itami, M., Nagata, H., Kudo, I., Iwasaki, A., Hosokawa, S.:"Performance of Droplet Collector in Liquid Droplet Radiator under Microgravity", Microgravity Science and Technology, 13(2):42-45(2002)*
2002年 - 「液滴ラジエータ用液滴生成器の微小重力下での性能」
戸谷 剛, 伊丹 雅洋, 藪田 茂, 永田 晴紀, 工藤 勲, 岩崎 晃, 細川 俊介
『日本機械学会論文集(B編)』, 68, 668, 1166, 1173, 一般社団法人 日本機械学会, 2002年
日本語, The Liquid Droplet Radiator (LDR) has an advantage over comparable conventional radiators in terms of the rejected heat power-weight ratio. Therefore, the LDR has attracted as an advanced radiator for high-power space systems that will be prerequisite for large space structures. In this study, the performance of a droplet emittor under microgravity condition has been investigated from the viewpoint of operational space use of the LDR in the future. From experiments, it is considered that the droplet emittor can produce uniform droplet streams under microgravity condition in the non-dimensional wave number range from 0.215 to 0.490. In this range, the droplet diameter and the spacing range are from 204 to 285 [μm] and from 445 to 1160 [μm] respectively. And it is concluded that these diameter and spacing can be estimated by the equations based on the law of conservation of mass in the process of generating droplets. - 「小型衛星のためのハイブリッドロケットの打ち上げ機の開発」
『日本マイクログラビティ応用学会誌』, 19, 2, 112, 116, 2002年 - 液滴ラジエータ用液滴生成器の微小重力下での性能(流体工学,流体機械)
戸谷 剛, 伊丹 雅洋, 藪田 茂, 永田 晴紀, 工藤 勲, 岩崎 晃, 細川 俊介
日本機械学會論文集. B編, 68, 668, 1166, 1173, 一般社団法人日本機械学会, 2002年
日本語, 研究論文(学術雑誌), The Liquid Droplet Radiator (LDR) has an advantage over comparable conventional radiators in terms of the rejected heat power-weight ratio. Therefore, the LDR has attracted as an advanced radiator for high-power space systems that will be prerequisite for large space structures. In this study, the performance of a droplet emittor under microgravity condition has been investigated from the viewpoint of operational space use of the LDR in the future. From experiments, it is considered that the droplet emittor can produce uniform droplet streams under microgravity condition in the non-dimensio... - 液滴ラジエータ用遠心式液滴回収器の微小重力下での性能(流体工学,流体機械)
戸谷 剛, 伊丹 雅洋, 藪田 茂, 永田 晴紀, 工藤 勲, 岩崎 晃, 細川 俊介
日本機械学會論文集. B編, 68, 674, 2780, 2787, 一般社団法人日本機械学会, 2002年
日本語, 研究論文(学術雑誌), The Liquid Droplet Radiator (LDR) has an advantage over conventional radiators in terms of the rejected heat power-weight ratio. LDR has been taken notice as an advanced radiator for highpower generation systems which will be prerequisite for large space structures. In this study, the performance of a centrifugal droplet collector under microgravity condition has been investigated from the viewpoint of operational space use of LDR in the future. It has been concluded that (1) a centrifugal collector is able to transport working fluid to a recirculating pump under microgravity condition; (2)... - 液滴ラジエータの作動流体循環に関する微小重力実験
戸谷 剛, 薮田 茂, 宮本 拓哉, 永田 晴紀, 工藤 勲
JASMA : Journal of the Japan Society of Microgravity Application, 18, 0, 87, 87, 2001年10月01日
英語 - 端面燃焼式ハイブリッドロケットの基礎研究:その2燃焼特性
橋本 望, 加藤 隆博, 永田 晴紀, 工藤 勲
日本航空宇宙学会論文集 = Journal of the Japan Society for Aeronautical and Space Sciences, 49, 565, 40, 47, 一般社団法人 日本航空宇宙学会, 2001年02月05日
日本語, To overcome defects of conventional hybrid rockets such as the loss of specific impulse, which is caused by the O/F shift during the combustion, and the low combustion efficiency, the authors have proposed a new idea of design. The point of this idea, named “End-Burning Hybrid Rocket, ” is that oxidizer gas flows in the gap space of a porous solid fuel bed. Diffusion flame is formed at the end of the solid fuel bed. Experimental studies were made to clarify the basic combustion characteristics of the propellant. Results show that pressure exponent of the burning rate with the same equivalence ratio is approximately 0.85 and virtually independent with the equivalence ratio. Using this result, a designing method of End-Burning Hybrid Rocket Motor is shown. Finally, thrust and specific impulse is estimated as functions of oxidizer gas flow rates to investigate the throttling characteristics of the motor. - 端面燃焼式ハイブリッドロケットの基礎研究:その1燃焼安定性
加藤 隆博, 橋本 望, 永田 晴紀, 工藤 勲
日本航空宇宙学会論文集 = Journal of the Japan Society for Aeronautical and Space Sciences, 49, 565, 33, 39, 一般社団法人 日本航空宇宙学会, 2001年02月05日
日本語, To overcome defects of conventional hybrid rockets such as low combustion efficiency and the O/F shift during the combustion, the authors have proposed a new form of hybrid rocket fuel. The fuel is a fibrous bed in which oxidizer gas flows. Stable diffusion flame appears at the exit surface. Previous researches show that sudden increase of the fuel regression rate occurs with the increase of ambient pressure. This sudden increase is attributed to the flame spreading between fuel fibers. To clarify the limit of fuel gap space the diffusion flame can spread into, experimental study was made. Critical gap space, which means the minimum gap space the diffusion flame can spread into, was obtained experimentally as a function of oxygen gas flow velocity and ambient pressure. Using this result, necessary conditions to realize a stable combustion with this new fuel form are shown. - 「端面燃焼式ハイブリッドロケットに関する基礎研究 その1:燃焼安定性(A Preliminary study of End-Burning Hybrid Rocket: Part 1 Combustion Stability)」
加藤 隆博, 橋本 望, 永田 晴紀, 工藤 勲
『日本航空宇宙学会論文集』, 49, 565, 33, 39, 一般社団法人 日本航空宇宙学会, 2001年
日本語, To overcome defects of conventional hybrid rockets such as low combustion efficiency and the O/F shift during the combustion, the authors have proposed a new form of hybrid rocket fuel. The fuel is a fibrous bed in which oxidizer gas flows. Stable diffusion flame appears at the exit surface. Previous researches show that sudden increase of the fuel regression rate occurs with the increase of ambient pressure. This sudden increase is attributed to the flame spreading between fuel fibers. To clarify the limit of fuel gap space the diffusion flame can spread into, experimental study was made. Critical gap space, which means the minimum gap space the diffusion flame can spread into, was obtained experimentally as a function of oxygen gas flow velocity and ambient pressure. Using this result, necessary conditions to realize a stable combustion with this new fuel form are shown. - Harunori NAGATA, Masahiro SASAKI, Takakage ARAI, Tsuyoshi TOTANI, Isao KUDO, "Evaluation of Mass Transfer Coefficient and Hydrogen Concentration in Supersonic Flow by Using Catalytic Reaction," Proceedings of the Combustion Institute, 28: 713-719, 2001.*
2001年 - 「折り目が残るインフレータブルチューブの微小重力下での展開挙動(Deployments of Inflatable Tubes with a Plastic Fold under Microgravity)」
戸谷 剛, 潮 敬之, 永田 晴紀, 工藤 勲
『日本機械学会論文集(C)』, 67, 655, 633, 640, 2001年
英語, 研究論文(学術雑誌), Inflatable structures have attracted considerable attention as space structures. In many space experiments, folds occurred during deployments of inflatable structures. They had undesirable influences on the body of space structures. This paper is intended as an investigation of deploying behaviors under microgravity of inflatable tubes that have a plastic fold on the center and a mass block on the tip. During their deployments, spring-back phenomena were happened in some experimental conditions. In order to examine these spring-back phenomena, a numerical simulation was conducted using the equation of motion under a consideration of restorable moments with the plastic folds. As a result, simulations corresponded well with experiments. Consequently, a following conclusion was obtained: Spring-back phenomena were caused by restorable moments of the plastic folds. © 2001, The Japan Society of Mechanical Engineers. All rights reserved. - 「空気-水素2次元超音速混合層の白金触媒反応を用いた水素濃度分布測定(Hydrogen Concentration of 2-D Supersonic Air-Hydrogen Mixing Layer Using Platinum Catalytic Reaction)」
Takakage Arai, Jiro Kasahara, Junji Miura, Fuminori Sakima, Harunori Nagata
『日本機械学会論文集(C)』, 67, 656, 934, 939, Japan Society of Mechanical Engineers, 2001年
日本語, 研究論文(学術雑誌), Abstrcat To investigate development of an air-hydrogen supersonic shear layer and distribution of hydrogen concentration, a hydrogen jet was injected into a cold air supersonic free-stream in a paralell direction. The free stream Mach number was 1.81. Using a catalytic reaction on a platinum wire, heat release due to catalytic reaction, a heat transfer coefficient and hydrogen concentration were measured. It was shown that paralell injection was found to affect on mixing condition. The effect of paralell injection on hydrogen concentration profile was clarified. It seemed that there was the stoichiometric condition at the outer edge of shear layer. It was confirmed that the diffusion of Hydrogen, including turbulent mixing, had an effect of flow configuration. - 「端面燃焼式ハイブリッドロケットに関する基礎研究 その2:燃焼特性(A Preliminary study of End-Burning Hybrid Rocket: Part 1 Combustion Characteristics)」
橋本 望, 加藤 隆博, 永田 晴紀, 工藤 勲
『日本航空宇宙学会論文集』, 49, 565, 40, 47, 一般社団法人 日本航空宇宙学会, 2001年
日本語, To overcome defects of conventional hybrid rockets such as the loss of specific impulse, which is caused by the O/F shift during the combustion, and the low combustion efficiency, the authors have proposed a new idea of design. The point of this idea, named “End-Burning Hybrid Rocket, ” is that oxidizer gas flows in the gap space of a porous solid fuel bed. Diffusion flame is formed at the end of the solid fuel bed. Experimental studies were made to clarify the basic combustion characteristics of the propellant. Results show that pressure exponent of the burning rate with the same equivalence ratio is approximately 0.85 and virtually independent with the equivalence ratio. Using this result, a designing method of End-Burning Hybrid Rocket Motor is shown. Finally, thrust and specific impulse is estimated as functions of oxidizer gas flow rates to investigate the throttling characteristics of the motor. - 折り目が残るインフレータブルチューブの微小重力下での展開挙動
戸谷 剛, 潮 敬之, 永田 晴紀, 工藤 勲
日本機械学會論文集. C編, 67, 655, 633, 640, 一般社団法人日本機械学会, 2001年
英語, Inflatable structures have attracted considerable attention as space structures. In many space experiments, folds occurred during deployments of inflatable structures. They had undesirable influences on the body of space structures. This paper is intended as an investigation of deploying behaviors under microgravity of inflatable tubes that have a plastic fold on the center and a mass block on the tip. During their deployments, spring-back phenomena were happened in some experimental conditions. In order to examine these spring-back phenomena, a numerical simulation was conducted using the ... - 空気-水素2次元超音速混合層の白金触媒反応を用いた水素濃度分布測定 : 流体工学,流体機械
新井 隆景, 笠原 次郎, 三浦 淳二, 咲間 文順, 永田 晴紀
日本機械学會論文集. B編, 67, 656, 934, 939, 一般社団法人日本機械学会, 2001年
日本語, To investigate development of an air-hydrogen supersonic shear layer and distribution of hydrogen concentration, a hydrogen jet was injected into a cold air supersonic free-stream in a parallel direction. The free stream Mach number was 1.81, Using a catalytic reaction on a platinum wire, heat release due to catalytic reaction, a heat transfer coefficient and hydrogen concentration were measured. It was shown that parallel injection was found to affect on mixing condition. The effect of parallel injection on hydrogen concentration profile was clarified. It seemed that there was the stoichio... - 液滴ラジエータ要素 (液滴回収器,ギアポンプ)の微小重力下での性能
戸谷 剛, 藪田 茂, 永田 晴紀, 工藤 勲, 岩崎 晃, 細川 俊介
JASMA : Journal of the Japan Society of Microgravity Application, 17, 0, 18, 19, 2000年10月01日
英語 - 端面燃焼式ハイブリッドロケット用プロペラントの燃焼機構の研究
永田 晴紀, 橋本 望, 加藤 隆博, 藤田 修, 伊藤 献一, 工藤 勲, 秋葉 鐐二郎
JASMA : Journal of the Japan Society of Microgravity Application, 17, 3, 172, 177, 日本マイクログラビティ応用学会, 2000年07月31日
日本語, Experimental investigations are carried out about the combustion of new form of hybrid rocket propellants, in which oxidizer gas flows gap space in a fibrous fuel bed, under normal and micro gravity conditions. The new form of propellant has a potential to improve combustion efficiency and thrust level of hybrid rocket motors. Regression rates of fibrous strand fuels are measured with lower oxygen gas flow rates of near the combustion limit. Main results are in the followings: Gravity effects decrease with decreasing the oxygen gas flow rate until the effect almost diminishes near the combustion limit. In a certain range of oxygen gas flow rates near the combustion limit, the equivalence ratio keeps a constant value. The range of oxygen gas flow rates, in which equivalence ratio is constant, does not depend on the ambient pressure. - Kentaro TAKAHASHI, Harunori NAGATA, Isao KUDO, "Behavior Monitoring of the Deployment of an Inflatable Disk under Microgravity for a Cold Welding Test Satellite," Space Forum, Vol. 6, pp. 397-402, 2000.
2000年 - Sosuke NAKAMURA, Harunori NAGATA, Isao KUDO, Kenichi ITO, Yasuhiro TAKESHITA, "Research on Flame Shape of Spherical Quasi-Liquid Samples under Microgavity Conditions," Space Forum, Vol. 6, pp. 329-334, 2000.
2000年 - 微小重力下における液滴ラジエータ要素の性能試験
伊丹 雅洋, 戸谷 剛, 永田 晴紀, 工藤 勲, 岩崎 晃, 細川 俊介
JASMA : Journal of the Japan Society of Microgravity Application, 16, 0, 114, 115, 1999年10月01日
英語 - 面状に配置された擬似液滴燃料群の火災形状に関する研究
中村 聡介, 永田 晴紀, 工藤 勲, 伊藤 献一, 北野 邦尋, 竹下 保弘
JASMA : Journal of the Japan Society of Microgravity Application, 16, 3, 191, 197, 日本マイクログラビティ応用学会, 1999年07月31日
日本語, To investigate the interaction between the droplets in the spray combustion, microgravity experiments were performed. The experimental apparatus mainly consists of a combustion chamber, an 8 mm video camera and a 35 mm single-lens reflex camera. Fuel droplets of butanol or hexanol are arranged 2-dimensionally, representing the fuel droplet cloud in the splay. Electrically heated nichrome wires ignite the fuel samples simultaneously. The 35 mm camera observes the flame shape history. The effects of sample space on the flame shape are investigated . Main results obtained are in the followings: Because of the interaction between droplets, the unsteady combustion period. in which fuel vapor is accumulated around the droplets, becomes long . As the result, the steady state combustion is not observed in the combustion process and the flame size keeps increasing until it dis appears. - インフレータブルチューブの無重力環境における展開シミュレーション
高野 昌宏, 永田 晴紀, 工藤 動
日本機械学會論文集. C編, 65, 633, 1978, 1984, 一般社団法人日本機械学会, 1999年05月25日
日本語, An inflatable tube is used for separation of two satellites which artificially generate variable gravity environment by rotating each other. This tube is deployed by nitrogen gas and it gets sufficient rigidity finally. Behavior at deployment of the inflatable tube which had been stowed in a test apparatus at first was monitored under microgravity using the world longest dropshaft. The satellite model was successfully deployed, even though it encountered a catastrophic break Up. It verified that inflatable tube had excellent characteristics of recovering from a break up by increasing inner ... - Evaluation of supersonic turbulent mixing using catalytic combustion of constant temperature Pt wire
T Arai, H Nagata, A Endo, H Sugiyama, S Morita, H Hosokawa
JSME INTERNATIONAL JOURNAL SERIES B-FLUIDS AND THERMAL ENGINEERING, 42, 1, 65, 70, JAPAN SOC MECHANICAL ENGINEERS, 1999年02月
英語, 研究論文(学術雑誌), Supersonic combustion using catalytic wire at constant temperature in a cold supersonic flow field was investigated in a square duct with a backward-facing Step. The free stream Mach number was M(m) = 1.81. Hydrogen was injected transversely behind a backward-facing step into a cold air free-stream. The heat released from the catalytic combustion had no effect on the temperature of the catalyst. This indicates that the reaction rate of the catalytic combustion observed in this study was determined by the concentration of H(2) and/or O(2) on the surface of the catalyst. The spatial distribution of heat released from the catalytic combustion in supersonic turbulent mixing layer, corresponds to the spatial distribution of concentration of H(2) and/or O(2) in local, was obtained. It was found that the most suitable position for supersonic combustion was at the outer edge of the mixing layer. - Combustion Characteristics of Fibrous Fuels for Dry Towel Hybrid Rocket Motor
Harunori NAGATA, Keiji OKADA, Takashi SAN'DA, Takahiro KATO, Ryojiro AKIBA, Shin SATORI, Isao KUDO
Journal of Space Technology and Science, 13, 1, 1, 1999年, [査読有り]
英語, 研究論文(学術雑誌) - Takakage ARAI, Harunori NAGATA, Akira Endo, Hiromu SUGIYAMA, Shuji MORITA, and Hiroshi HOSOKAWA, "Evaluation of Supersonic Turbulent Mixing Using Catalytic Combustion of Constant Temperature Pt Wire", JSME International Journal, Series B, Vol.42, No.1,・・・
1999年
Takakage ARAI, Harunori NAGATA, Akira Endo, Hiromu SUGIYAMA, Shuji MORITA, and Hiroshi HOSOKAWA, "Evaluation of Supersonic Turbulent Mixing Using Catalytic Combustion of Constant Temperature Pt Wire", JSME International Journal, Series B, Vol.42, No.1, pp.65-70 (1999). - 「面状に配置された疑似燃料液滴群の火炎形状に関する研究」
中村 聡介, 永田 晴紀, 工藤 勲, 伊藤 献一, 北野 邦尋, 竹下 保弘
日本マイクログラビティ応用学会誌, Vol.16, No.3, 191, 197, 日本マイクログラビティ応用学会, 1999年
日本語 - 「水素-空気超音速混合層における触媒反応を利用した水素濃度の評価」
永田 晴紀, 細川 博, 新井 隆景, 森田 修至, 工藤 勲
日本機械学会論文集(B編), 65, 636, 2666, 2671, Japan Society of Mechanical Engineers, 1999年
日本語, 研究論文(学術雑誌), The authors propose a new simple method which can be used to evaluate hydrogen concentration in hydrogen-air supersonic mixing layers without the need for costly apparatus. The catalytic reaction occurs on an electrically heated platinum wire in hydrogen-air supersoic mixing layers. By- adapting the technique of constant temperature type hotwire anemometers, a catalytic heat release rate is measured. A series of experiments with different Pt wire temperatures shows that Pt wire temperature has little effect on the catalytic heat release rate, implying that the rate of transfer of molecules to the Pt wire surface is the controlling factor. Accordingly, the heat release rate is related to the hydrogen concentration in the flow. The profile of hydrogen concentration is obtained by assuming the equivalent spatial distribution of heat and mass transfer. Stoichiometric conditions are found to be realized in the mixing layer. - 「インフレータブルチューブの無重力環境における展開シミュレーション」
高野 昌宏, 永田 晴紀, 工藤 動
日本機械学会論文集(C編), 65, 633, 1978, 1984, 一般社団法人日本機械学会, 1999年
日本語, An inflatable tube is used for separation of two satellites which artificially generate variable gravity environment by rotating each other. This tube is deployed by nitrogen gas and it gets sufficient rigidity finally. Behavior at deployment of the inflatable tube which had been stowed in a test apparatus at first was monitored under microgravity using the world longest dropshaft. The satellite model was successfully deployed, even though it encountered a catastrophic break Up. It verified that inflatable tube had excellent characteristics of recovering from a break up by increasing inner gas pressure. Typical break up condition of tube was analyzed by numerical simulation. Experimental data was successfully explained by the analysis. - 水素-空気超音速混合層における触媒反応を利用した水素濃度の評価
永田 晴紀, 細川 博, 新井 隆景, 森田 修至, 工藤 勲
日本機械学會論文集. B編, 65, 636, 2666, 2671, 一般社団法人日本機械学会, 1999年
日本語, The authors propose a new simple method which can be used to evaluate hydrogen concentration in hydrogen-air supersonic mixing layers without the need for costly apparatus. The catalytic reaction occurs on an electrically heated platinum wire in hydrogen-air supersoic mixing layers. By adapting the technique of constant temperature type hotwire anemometers, a catalytic heat release rate is measured. A series of experiments with different Pt wire temperatures shows that Pt wire temperature has little effect on the catalytic heat release rate, implying that the rate of transfer of molecules t... - 酸化剤噴出面で燃焼するフィルター状固体燃料の燃焼特性
加藤 隆博, 永田 晴紀, 秋葉 鐐二郎, 工藤 勲
JASMA : Journal of the Japan Society of Microgravity Application, 15, 0, 63, 64, 1998年10月01日
英語 - 微小重力における液滴ラジエータの評価
伊丹 雅洋, 戸谷 剛, 永田 晴紀, 工藤 勲, 岩崎 晃, 細川 俊介
JASMA : Journal of the Japan Society of Microgravity Application, 15, 0, 38, 39, 1998年10月01日
英語 - 水素-空気混合気熱面点火における点火遅れに及ぼす直流電界の影響
瀬川 大資, 永田 晴紀, 岸 武行, 角田 敏一, 津江 光洋, 河野 通方
日本機械学會論文集. B編, 64, 623, 2319, 2324, 一般社団法人日本機械学会, 1998年07月25日
日本語, The present study was carried out to reveal the possibility of controlling the ignition delay of hydrogen-air mixtures by applying electric fields. A quiescent stoichiometric hydrogen-air mixture was ignited by a suddenly heated thin wire of nickel or tungsten. DC electric fields were applied between the wire and outer electrode plates parallel with the wire. The mean ignition delay was calculated stochastically from the measured ignition delays which scattered considerably. Both with the nickel wire and with the tungsten wire, positive voltages applied to the outer electrode plates resulte... - 定温触媒線を用いた触媒燃焼による超音速乱流混合の評価方法
新井 隆景, 永田 晴紀, 遠藤 彰, 杉山 弘, 森田 修至, 細川 博
日本機械学會論文集. B編, 64, 619, 793, 799, 一般社団法人日本機械学会, 1998年03月25日
日本語, Supersonic combustion using catalytic wire at constant temperature in a cold supersonic flow field was investigated in a square duct with a backward-facing step. The free stream Mach number was of M_m=1.81. Hydrogen was injected transversely behind a backward-facing step into a cold air free stream. The heat release due to the catalytic combustion has no effect of the temperature of catalyst. It indicates that the reaction rate of the catalytic combustion observed in this study was determined by the consentration of H_2 and/or O_2 on the surface of the catalyst. The spatial distribution of ... - 可燃性混合気熱面点火の点火遅れに及ぼす電界の影響
瀬川 大資, 永田 晴紀, 岸 武行, 角田 敏一, 津江 光洋, 河野 通方
日本機械学會論文集. B編, 64, 617, 298, 304, 一般社団法人日本機械学会, 1998年01月25日
日本語, The present study was carried out to reveal the possibility of controlling the ignition delay of combustible mixtures by applying electric fields. A thin nickel wire was used as a hot surface to ignite the mixtures. It was suddenly heated up and then its temperature was kept constant. Quiescent propane-air mixtures were used as combustible mixtures. DC electric fields were applied between the nickel wire and the outer electrode plates parallel with the nickel wire. As the applied voltage to the electrode plates increased, both the mean values and the fluctuations of the ignition delay decre... - 「可燃性混合気熱面点火の点火遅れに及ぼす電界の影響」
瀬川 大資, 永田 晴紀, 岸 武行, 角田 敏一, 津江 光洋, 河野 通方
『日本機械学会論文集(B編)』, 64, 617, 298, 304, 一般社団法人日本機械学会, 1998年
日本語, The present study was carried out to reveal the possibility of controlling the ignition delay of combustible mixtures by applying electric fields. A thin nickel wire was used as a hot surface to ignite the mixtures. It was suddenly heated up and then its temperature was kept constant. Quiescent propane-air mixtures were used as combustible mixtures. DC electric fields were applied between the nickel wire and the outer electrode plates parallel with the nickel wire. As the applied voltage to the electrode plates increased, both the mean values and the fluctuations of the ignition delay decreased regardless of the electric polarities. Without electric fields, both the mean ignition delay and the fluctuation varied with the equivalence ratio of the mixture. With increasing the applied voltage up to =2kV, the variations became to disappear. When the surface temperature was low, the stoichiometric mixture could be ignited only by applying the electric fields. - 「水素-空気混合気熱面点火における点火遅れに及ぼす直流電界の影響」
瀬川 大資, 永田 晴紀, 岸 武行, 角田 敏一, 津江 光洋, 河野 通方
『日本機械学会論文集(B編)』, 64, 623, 2319, 2324, 一般社団法人日本機械学会, 1998年
日本語, The present study was carried out to reveal the possibility of controlling the ignition delay of hydrogen-air mixtures by applying electric fields. A quiescent stoichiometric hydrogen-air mixture was ignited by a suddenly heated thin wire of nickel or tungsten. DC electric fields were applied between the wire and outer electrode plates parallel with the wire. The mean ignition delay was calculated stochastically from the measured ignition delays which scattered considerably. Both with the nickel wire and with the tungsten wire, positive voltages applied to the outer electrode plates resulted in almost the same change of the mean ignition delay ; The mean ignition delay showed a slight increase at the lower applied voltages, while it became very short at the highest applied voltages. When the nickel wire was used and negative voltages were applied, the mean ignition delay decreased even at the lower voltages. However, it showed little change with the tungsten wire and the negative applied voltages. - 「定温触媒線を用いた触媒燃焼による超音速乱流混合の評価方法」
新井隆景, 永田晴紀, 遠藤彰, 杉山弘, 森田修至, 細川博
『日本機械学会論文集(B編)』, 64, 619, 793, 799, 一般社団法人日本機械学会, 1998年
日本語 - 後向きステップを過ぎる低温の超音速流れ中に垂直に噴射される水素の触媒燃焼〔流体工学, 流体機械〕
新井 隆景, 遠藤 彰, 永田 晴紀, 杉山 弘, 森田 修至
日本機械学會論文集. B編, 63, 614, 3318, 3324, 一般社団法人日本機械学会, 1997年10月25日
日本語, Supersonic combustion using a catalytic combustion in a cold supersonic flow field was investigated in a square duct with a backward-facing step. The free stream Mech number was M_m=1.81. Hydrogen was injected transversely behind a backward-facing step into a cold air free stream. Using a catalyst in a cold supersonic turbulent mixing layer, it was found that hydrogen reacted stably to oxygen in the air flow. The relationship between the heat release due to catalytic combustion and supersonic flow properties, which influence the supersonic combustion, was clarified experimentally. The spati... - 微小重力場における擬似液滴燃料列の燃え広がり
永田 晴紀, 工藤 勲, 北野 邦尋, 中村 聡介, 伊藤 献一, 竹下 保弘
JASMA : Journal of the Japan Society of Microgravity Application, 14, 0, 41, 42, 1997年10月01日
英語 - インフレータブルチューブの微小重力下における展開
高野 昌宏, 永田 晴紀, 工藤 勲
JASMA : Journal of the Japan Society of Microgravity Application, 14, 0, 27, 28, 1997年10月01日
英語 - ウェットタオル式ハイブリッドロケットに関する基礎研究
永田 晴紀, 秋葉 鐐二郎, 棚次 亘弘, 高野 雅弘, 横田 力男, 加勇田 清勇
日本航空宇宙学会誌 = Journal of the Japan Society for Aeronautical and Space Sciences, 45, 522, 365, 370, 日本航空宇宙学会, 1997年07月
日本語 - 「後向きステップを過ぎる低温の超音速流中に垂直に噴射される水素の触媒燃焼」
新井隆景, 遠藤彰, 永田晴紀, 杉山弘, 森田修至
『日本機械学会論文集 (B編)』, 63, 614, 3318, 3324, 一般社団法人 日本機械学会, 1997年, [国内誌]
日本語, Supersonic combustion using a catalytic combustion in a cold supersonic flow field was investigated in a square duct with a backward-facing step. The free stream Mech number was M m = 1.81. Hydrogen was injected transversely behind a backward-facing step into a cold air free stream. Using a catalyst in a cold supersonic turbulent mixing layer, it was found that hydrogen reacted stably to oxygen in the air flow. The relationship between the heat release due to catalytic combustion and supersonic flow properties, which influence the supersonic combustion, was clarified experimentally. The spatial distribution of heat release generated by catalytic combustion in the supersonic turbulent mixing layer is discussed. It was found that the heat release due to the catalytic combustion had a maximum at the outer edge of the mixing layer. - 「ウェットタオル式ハイブリッドロケットに関する基礎研究」
永田 晴紀, 秋葉 鐐二郎, 棚次 亘弘, 高野 雅弘, 横田 力男, 加勇田 清勇
『日本航空宇宙学会誌』, 45, 522, 365, 370, 日本航空宇宙学会, 1997年
日本語, To overcome defects of conventional hybrid rocket motors such as poor mass ratio and low combustion efficiency, the authors propose a new idea of design. The point of this idea is that the motor uses ‘wet towel propellant’, in which liquid oxidizer fills gap space in the fuel bed consisting of plastic fibers or films. Specific impulses of the propellants are estimated theoretically with HAN and LOX as liquid oxidizer. To investigate pressure sensitivity of the burning rates of the propellants, burning rates of LOX-polyaramid propellants are measured in a pressure range of 0.1 to 2MPa with a specially designed strand burner. A striking feature in the experimental results is that the enhancement of the burning rate with ambient pressure is as big as the pressure exponent being larger than one. Further innovation is necessary for realization of the proposed idea.
その他活動・業績
- 地球磁気圏X線撮像衛星GEO-Xの現状
江副祐一郎, 船瀬龍, 船瀬龍, 永田晴紀, 三好由純, 中嶋大, 三石郁之, 布施綾太, 川端洋輔, BODEN Ralf C., 中島晋太郎, KAMPS Landon, 信原佑樹, 平井翔太, 石川久美, 沼澤正樹, 佐藤佑樹, 萩野浩一, 松本洋介, 細川敬祐, 伊師大貴, 米山友景, 上野宗孝, 山崎敦, 長谷川洋, 三田信, 三谷烈史, 藤本正樹, 川勝康弘, 岩田隆浩, 満田和久, 平賀純子, 笠原慧, 小泉宏之, 佐原宏典, 金森義明, 森下浩平, 日本天文学会年会講演予稿集, 2024, 2024年 - Regenerative Cooling of Graphite Nozzles for Throat Erosion Suppression
Hiroki Kojima, Landon T. Kamps, Yuki Nobuhara, Giuseppe Gallo, Harunori Nagata, AIAA SCITECH 2023 Forum, 2023年01月19日
American Institute of Aeronautics and Astronautics - 地球磁気圏X線撮像計画GEO-X(GEOspace X-ray imager)の現状 V
中嶋大, 江副祐一郎, 船瀬龍, 永田晴紀, 三好由純, 沼澤正樹, 石川久美, 萩野浩一, 三石郁之, KAMPS Landon, 川端洋輔, 布施綾太, BODEN Ralf, 三谷烈史, 米山友景, 中島晋太郎, 上野宗孝, 山崎敦, 長谷川洋, 三田信, 藤本正樹, 川勝康弘, 岩田隆浩, 松本洋介, 細川敬祐, 平賀純子, 満田和久, 小泉宏之, 笠原慧, 佐原宏典, 金森義明, 森下浩平, 日本天文学会年会講演予稿集, 2023, 2023年 - UNISEC二十年の歴史と今後の展望
川島レイ, 桑原聡文, 坂本啓, 永田晴紀, 船瀬龍, 宮崎康行, 山崎政彦, 中須賀真一, 宇宙科学技術連合講演会講演集(CD-ROM), 67th, 2023年 - 高推力推進系を有する超小型衛星GEO-Xのバスシステム検討状況
布施綾太, BODEN Ralf, 中島晋太郎, 川端洋輔, 松下将典, 秋山茉莉子, 船瀬龍, 船瀬龍, 江副祐一郎, KAMPS Landon, 永田晴紀, 鈴木聡宏, NERY Vinicius, 伊藤湧太郎, 筒井真輝, 望月友貴, 小川巧海, 荻野浩佑, 草野湧貴, 下村俊介, 瀬戸翔一, 中村陸希, 宇宙科学技術連合講演会講演集(CD-ROM), 67th, 2023年 - 超小型衛星による惑星磁気圏のX線撮像観測
中嶋大, 江副祐一郎, 三好由純, 永田晴紀, 船瀬龍, 宇宙科学技術連合講演会講演集(CD-ROM), 67th, 2023年 - ライドシェア小型宇宙機用ハイブリッド化学推進系の開発
KAMPS Landon, 平井翔大, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 67th, 2023年 - 小型衛星用推進系の低圧環境下推力取得に向けたディフューザ形状の最適化の検討
糸魚川 大和, 小野寺 遼, ガロ ジュセッペ, ケンプス ランドン, 永田 晴紀, 年次大会, 2023, S191p-06, 2023年
In order to increase the opportunities for deep space exploration by small satellites, our laboratory has been working on developing a safe, high-thrust hybrid rocket motor. To obtain the thrust characteristics of the hybrid rocket under low-pressure conditions, the authors developed a HATS (High Altitude Test Stand) for small satellite thrusters that can maintain a low-pressure environment during combustion tests and found that the current combustion conditions and diffuser shape could not maintain a low-pressure environment. Therefore, this study designed and verified a new diffuser with a second throat in the middle of the diffuser to maintain a low-pressure environment. Using the two theoretical models presented in a previous study, we confirmed that there exists a design solution for a diffuser shape that can maintain a low-pressure environment under the current combustion conditions. Although the design solution is close to the lower limit of the second throat diameter, the design solution obtained from the theoretical model is not highly reliable because few previous studies experimentally investigated the lower limit. Therefore, the design solution was verified using a sub-scale diffuser and cold gas., 一般社団法人 日本機械学会, 日本語 - Nytrox/HDPEを用いたハイブリッドロケットの燃焼特性
池田 拓矢, 信原 佑樹, ケンプス ランドン, 永田 晴紀, 年次大会, 2023, S191p-04, 2023年
A hybrid rocket kick motor is being developed for low-cost deep space exploration. Nytrox, a mixture of nitrous oxide (N2O) and oxygen, will be investigated as an oxidizer for the kick motor. Nytrox has the advantage for use in kick motors of being safer than N2O because it is less likely to cause decomposition reactions. The objective of this study is to obtain the fuel regression rate equation by conducting combustion tests to design a kick motor using Nytrox and high-density polyethylene (HDPE). The Nytrox used in this study was generated under conditions where the probability of N2O decomposition reactions occurring was as low as possible. The fuel regression rate equation of Nytrox/HDPE was obtained from five combustion tests. Comparison with the fuel regression rate of N2O /HDPE obtained from two combustion tests indicated that the fuel regression rate of Nytrox may be larger than that of N2O, depending on the fuel. However, the number of experiments is small, and more data is needed to obtain accurate data., 一般社団法人 日本機械学会, 日本語 - 自己加圧供給される亜酸化窒素の流量特性に関する研究
小野寺 遼, 糸魚川 大和, Kamps Landon, 永田 晴紀, 年次大会, 2023, S191p-09, 2023年
Nitrous oxide is an attractive option as an oxidizer for rockets because of its self-pressurized supply and safety. On the other hand, it is very close to the critical point at room temperature, so the prediction of its state quantity requires very complicated calculations. Prediction of the oxidizer flow rate is important for hybrid rockets, and various prediction equations have been proposed. In particular, the Homogeneous Non-Equilibrium Flow Model, which assumes that the mass flow rate of nitrous oxide takes a value between liquid single-phase flow and gas-liquid two-phase flow, is considered to be accurate. However, this HNE flow model does not consider the effect of injector geometry, and the authors proposed a new model, HNEIS, which considers injector geometry. From flow rate tests using several injectors with different geometries, it was found that the flow rate varies depending on the injector geometry even when the supply pressure is equal. It was also clear that the HNEIS could express this trend, whereas the HNE could not. On the other hand, there is a dissociation between the actual flow rate and the predicted flow rate by HNEIS, and the model needs to be improved., 一般社団法人 日本機械学会, 日本語 - カーボン粒子により導電性を付与した燃料樹脂の燃焼特性
江澤 悠太, 永田 晴紀, Kamps Landon, 信原 佑樹, 年次大会, 2023, S191p-07, 2023年
The objective of this study is to design and develop a hybrid rocket that is re-ignitable and does not require an igniter, focusing on the conductivity obtained by the addition of a large amount of carbon black. By conducting five and eight combustion tests using pure PLA fuel and conductive fuel with a large amount of carbon black added to PLA as solid fuel and gaseous oxygen as oxidizer, respectively, the fuel regression rate equations necessary for the proper design of hybrid rockets were established. The respective fuel regression rate equations were as follows(R2 = 0.8154), Conductive fuel:
(R2 = 0.3975). Since the oxidizer index in the fuel regression rate equation exceeded 1 in the conductive fuel results, more data should be collected in the future to improve the equation., 一般社団法人 日本機械学会, 日本語
- Application of Low Concentration Hydrogen Peroxide for Hybrid Rocket Propulsion System for Small Spacecraft
Masashi Wakita, Yusuke Takada, Kodai Iwanaga, Toshiaki Iizuka, Hironori Sahara, Landon T. Kamps, Harunori Nagata, AIAA AVIATION 2022 Forum, 2022年06月27日
American Institute of Aeronautics and Astronautics - 観測ロケットS-520を用いた宇宙環境下でのハイブリッドキックモータ実験の提案
ケンプス, ランドン, 平井, 翔大, 永田, 晴紀, KAMPS, Landon, HIRAI, Shota, NAGATA, Harunori, 観測ロケットシンポジウム2021 講演集 = Proceedings of Sounding Rocket Symposium 2021, 2022年03月
第4回観測ロケットシンポジウム(2022年3月14-15日. ハイブリッド開催(JAXA相模原キャンパス& オンライン))
4th Sounding Rocket Symposium(March 14-15, 2022. Hybrid(in-person & online) Conference (Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan
資料番号: SA6000175012
レポート番号: Ⅱ-7, 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS), 英語 - 導電性プラスチックを用いた点火装置の点火条件
信原佑樹, 平井翔大, LEUNG Yownin Albert, KAMPS Landon, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 66th, 2022年 - GEO-X計画の現状と将来展望
江副祐一郎, 船瀬龍, 船瀬龍, 川端洋輔, 中島晋太郎, 永田晴紀, KAMPS Landon, 中嶋大, 三石郁之, 石川久美, 沼澤正樹, 三好由純, 上野宗孝, 宇宙科学技術連合講演会講演集(CD-ROM), 66th, 2022年 - 端面燃焼式ハイブリッドロケットにおける吹き飛び発生機構の解明
早坂宏己, 深田真衣, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 66th, 2022年 - 小型相乗り宇宙機用ハイブリッドキックモータの開発状況
平井翔大, KAMPS Landon, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 66th, 2022年 - 液体酸素を酸化剤とする端面燃焼式ハイブリッドロケットの開発
永田晴紀, 鈴木翔, 李介維, 添田健太郎, KAMPS Landon, 脇田督司, 宇宙科学技術連合講演会講演集(CD-ROM), 66th, 2022年 - エタノール/LOXを用いた液体ロケットにおけるグラファイト製ノズルスロートの浸食解析
糸魚川 大和, 平井 翔太, 津地 歩, 金井 竜一朗, ケンプス ランドン, 永田 晴紀, 年次大会, 2022, J191p-03, 2022年
Interstellar Technologies, Inc. has been developing a series of liquid rockets using ethanol/LOX called MOMO. These rockets use graphite nozzle throat inserts for which nozzle erosion, referring to the increase in diameter of the nozzle throat due to the chemical reactions between the combustion gas and the graphite, has become a problem. This study aims to develop a predictive equation for the nozzle erosion rate of liquid rockets using ethanol/LOX. The authors analyzed data from seven long-duration combustion experiments and calculated the nozzle erosion rate. Using the heat transfer coefficient of the nozzle throat, the heat transfer equation at the nozzle throat was solved and the nozzle wall temperature was also estimated. As a result, it was found that the nozzle erosion phenomenon of liquid rockets using ethanol/LOX has a similar trend to that of nozzle erosion in hybrid rockets using HDPE/LOX. A good correlation was obtained by using the nozzle erosion rate prediction equation of Kamps et al., which has chamber pressure, equivalence ratio and nozzle wall temperature as parameters., 一般社団法人 日本機械学会, 日本語 - 60wt%予熱過酸化水素水を用いたハイブリッドロケットの保炎特性
小野 玄太, 岩永 昂大, 高田 裕亮, 脇田 督司, ケンプス ランドン, 永田 晴紀, 年次大会, 2022, J191-07, 2022年
As the number of micro-deep space probes launched by large rockets using the piggyback increases, the demand for safe, low-cost, re-ignitable, and high-thrust kick motors for efficient inter-orbit satellite transfer is expected to increase as well. Low-concentration hydrogen peroxide is the most suitable oxidizer for hybrid kick motors for micro-spacecraft in terms of safety, low cost, and ease of storage. However, it has difficulty in ignitability and flame holding due to its high water content. In this study, we considered a method to decompose low-concentration hydrogen peroxide by external heating using a heat exchanger before supplying it to a CAMUI-type fuel. A ground combustion experiment showed that flame holding was successfully achieved, however, there were vibrations in the combustion chamber pressure. In order to obtain the lower limit of the heating required for stable combustion, we plan to evaluate the flame holding limit and calculate the heat input to the hydrogen peroxide., 一般社団法人 日本機械学会, 日本語 - 地球磁気圏X線撮像計画GEO-X(GEOspace X-ray imager)の現状 IV
中嶋大, 江副祐一郎, 船瀬龍, 永田晴紀, 三好由純, 萩野浩一, 沼澤正樹, 石川久美, 三石郁之, KAMPS L., 川端洋輔, 布施綾太, 米山友景, 中島晋太郎, 上野宗孝, 山崎敦, 長谷川洋, 三田信, 藤本正樹, 川勝康弘, 岩田隆浩, 平賀純子, 満田和久, 小泉宏之, 笠原慧, 佐原宏典, 金森義明, 森下浩平, 日本天文学会年会講演予稿集, 2022, 2022年 - 液体酸素が流れる固体燃料管列の燃え拡がり特性—Flame Spreading Characteristics of Solid Fuel Tube Row with Liquid Oxygen
野中, 響己, 津地, 歩, 李, 介維, 永田, 晴紀, 脇田, 督司, NONAKA, Hibiki, TSUJI, Ayumu, LI, Kaii, NAGATA, Harunori, WAKITA, Masashi, 令和3年度宇宙輸送シンポジウム: 講演集録 = Proceedings of Space Transportation Symposium FY2021, 2022年01月
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: SA6000173012
STCP-2021-012, 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS), 日本語 - 端面燃焼式ハイブリッドロケットにおける推力制御応答遅れ機構の解明
千葉健太郎, 鈴木翔, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 66th, 2022年, [査読有り]
© 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. This study is an investigation of the response mechanisms in Axial-Injection End-Burning Hybrid Rockets. A numerical model was developed based on the conservation of mass in the chamber. Since it was not known whether the fuel shape is affected by oxidizer port velocity, several single port firing tests were conducted to confirm the effect of fuel regression shape and ensure the precision of the model. The calculation results show that the slow first-order lag can be explained by the time in which ports merge with one another, and the fast first-order lag can be explained by the unsteady residence time in the chamber. However, there is no hysteresis characteristics in calculation results. Therefore, the hysteresis characteristics are not caused by unsteady residence time in the chamber and port merging combustion., American Institute of Aeronautics and Astronautics - Development of safe, low-cost, re-ignitable rocket ignition system
Shota Hirai, Landon T. Kamps, Harunori Nagata, AIAA Propulsion and Energy 2021 Forum, 2021年08月09日
American Institute of Aeronautics and Astronautics - Fuel Regression Characteristics of CAMUI type Hybrid Rocket Using Nitrous Oxide
Yuki Nobuhara, Landon T. Kamps, Harunori Nagata, AIAA Propulsion and Energy 2021 Forum, 2021年08月09日
American Institute of Aeronautics and Astronautics - Numerical Analysis of Nozzle Transient Heating and Erosion in Hybrid Rockets burning HDPE
Marco Rotondi, Mario Tindaro Migliorino, Daniele Bianchi, Landon T. Kamps, Harunori Nagata, AIAA Propulsion and Energy 2021 Forum, 2021年08月09日
American Institute of Aeronautics and Astronautics - Influence of Port Manufacturing Accuracy on Backfiring in Axial-Injection End-Burning Hybrid Rocket
Sho Suzuki, Ayumu Tsuji, Kentaro Soeda, Landon T. Kamps, Harunori Nagata, AIAA Propulsion and Energy 2021 Forum, 2021年08月09日
American Institute of Aeronautics and Astronautics - Ignition and flame-holding characteristics of 60wt% hydrogen peroxide in a CAMUI-type hybrid rocket fuel
Yusuke Takada, Kodai Iwanaga, Hajime Inoue, Shota Inoue, Masashi Wakita, Landon T. Kamps, Hironori Sahara, Toshiaki Iizuka, Harunori Nagata, AIAA Propulsion and Energy 2021 Forum, 2021年08月09日
American Institute of Aeronautics and Astronautics - Reconstruction techniques for determining O/F in hybrid rockets
Yuji Saito, Landon T. Kamps, Ayumu Tsuji, Harunori Nagata, AIAA Propulsion and Energy 2021 Forum, 2021年08月09日
American Institute of Aeronautics and Astronautics - 小型相乗り宇宙機用ハイブリッドキックモーターの開発状況
平井翔大, KAMPS Landon, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 65th, 2021年 - ハイブリッドキックモータを搭載する超小型深宇宙探査機の熱設計
友永優太, 戸谷剛, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 65th, 2021年 - GEO-X計画とその将来ビジョン-太陽系X線天文学
江副祐一郎, 船瀬龍, 船瀬龍, 永田晴紀, 三好由純, 中嶋大, 三石郁之, 川端洋輔, KAMPS Landon, 石川久美, 上野宗孝, 宇宙科学技術連合講演会講演集(CD-ROM), 65th, 2021年 - 地球磁気圏X線撮像計画GEO-X(GEOspace X-ray imager)の現状
江副祐一郎, 船瀬龍, 船瀬龍, 永田晴紀, 三好由純, 中嶋大, 三石郁之, 石川久美, 上野宗孝, 山崎敦, 長谷川洋, 三田信, 藤本正樹, 川勝康弘, 岩田隆浩, 満田和久, 平賀純子, 笠原慧, 佐原宏典, 金森義明, 森下浩平, 三谷烈史, 沼澤正樹, 日本地球惑星科学連合大会予稿集(Web), 2021, 2021年 - 3Dプリント燃料を利用したロケット用再点火装置の実験的研究
平井 翔大, KAMPS Landon, 永田 晴紀, 年次大会, 2021, J191-16, 2021年
The authors are developing a safe, low-cost, compact, and lightweight reignition device to provide reignition capability to a hybrid kick motor as a propulsion system for small satellites. So far, only HDPE has been used as a solid fuel for the reignition device, with which the desired results have not been obtained. In this study, we measured the amount of gasification of fuel due to the radiant heat of glow plugs using multiple filaments for 3D printing, with the aim of identifying suitable replacement materials for HDPE. A total of more than 100 experiments, including preliminary experiments, revealed that HDPE was the least suitable fuel for the reignition system. On the other hand, it was found that Carbon Fibered PLA is the most effective as it has a large amount of fuel gasification and can maintain the fuel shape., 一般社団法人 日本機械学会, 日本語 - 北海道大学・室蘭工業大学f3センター 革新航空機ユニット
内海 政春, 今井 良二, 溝端 一秀, 中田 大将, 永田 晴紀, 北海道支部講演会講演概要集, 2021.58, 1311, 2021年
一般社団法人 日本機械学会, 日本語 - 北海道大学・室蘭工業大学f3センターグリーン船舶ユニット
村井 祐一, 堀本 康文, 朴 炫珍, 田坂 裕司, 大石 義彦, 永田 晴紀, 北海道支部講演会講演概要集, 2021.58, 1312, 2021年
一般社団法人 日本機械学会, 日本語 - 北海道大学・室蘭工業大学f3工学センターと小型宇宙推進ユニット
永田 晴紀, ケンプス ランドン, 平井 翔大, 戸谷 剛, 村井 祐一, 内海 正春, 北海道支部講演会講演概要集, 2021.58, 1314, 2021年
一般社団法人 日本機械学会, 日本語 - 北海道大学・室蘭工業大学f3工学教育研究センターマイクロサットユニット
戸谷 剛, 坂本 祐二, 永田 晴紀, 北海道支部講演会講演概要集, 2021.58, 1313, 2021年
一般社団法人 日本機械学会, 日本語 - 地球磁気圏X線撮像計画GEO-X(GEOspace X-ray imager)の現状 III
江副祐一郎, 船瀬龍, 船瀬龍, 永田晴紀, 三好由純, 中嶋大, 三石郁之, 石川久美, 川端洋輔, 中島晋太郎, KAMPS Landon, 上野宗孝, 山崎敦, 長谷川洋, 三田信, 三谷烈史, 藤本正樹, 川勝康弘, 岩田隆浩, 満田和久, 平賀純子, 笠原慧, 小泉宏之, 佐原宏典, 金森義明, 森下浩平, 沼澤正樹, 日本天文学会年会講演予稿集, 2021, 2021年 - ハイブリッドロケット技術を応用したロケット用再点火装置内部の可視化および燃焼形態の解明
井上翔太, KAMPS Landon, 平井翔大, 高田裕亮, 深田真衣, LEUNG Yownin Albert, 永田晴紀, 燃焼シンポジウム講演論文集(CD-ROM), 59th, 2021年 - 白金触媒を用いた60wt%過酸化水素のハイブリッドロケットにおける保炎特性
岩永昂大, 高田裕亮, 小野玄太, 脇田督司, KAMPS Landon, 佐原宏典, 飯塚俊明, 永田晴紀, 燃焼シンポジウム講演論文集(CD-ROM), 59th, 2021年 - 60wt%過酸化水素水を用いたCAMUI型ハイブリッドロケットの点火特性
高田 裕亮, 高梨 知広, 脇田 督司, 永田 晴紀, 年次大会, 2020, J19103, 2020年
Micro-space-probes are ideal platforms for future deep space exploration missions because multiple micro-space-probes can piggyback together on most existing Earth launch vehicles. If the piggyback destination is GTO, the microsatellite can escape from the earth's gravity to go into deep space with only 700 m/s of additional acceleration. Low-concentration (60wt%) hydrogen peroxide has favorable storage characteristics as an oxidizer for in-space hybrid rocket propulsion but has unfavorable ignition and flame holding characteristics due to its high-water content. In this study, separate tests for ignition and flame holding were attempted using 60wt% hydrogen peroxide as an oxidizer and high-density polyethylene as fuel in a CAMUI-type configuration. A Pt catalyst, pressurized chamber, and low flow-rate oxidizer bypass line were used to promote ignition. The results of visualization experiments show that ignition was achieved, but flame holding was not achieved, probably due to an excessive oxidizer flow rate., 一般社団法人 日本機械学会, 日本語 - 超小型宇宙機用ハイブリッドロケット推進機の開発
ケンプス ランドン, 影山 理沙, 脇田 督田, 永田 晴紀, 年次大会, 2020, J19121, 2020年
Hybrid rockets are suitable candidate propulsion systems for space exploration missions because they can produce similar levels of thrust to solid and liquid bi-propellant rockets without the fire/explosion safety hazard. Space probes equipped with a hybrid rocket thrust can more readily be integrated in piggy-back launches without the need for lengthy and expensive risk mitigation activities. This study reports the results of four static firing tests of an engineering model hybrid rocket thruster using high-density polyethylene as fuel and liquid nitrous oxide as oxidizer for space-probe propulsion. Measurements of thrust, chamber pressure and propellant mass flow rates reveal that performance in atmospheric conditions was nominal: Isp ≦ 220s; CF ≦ 1.4., 一般社団法人 日本機械学会, 日本語 - 端面燃焼式ハイブリッドロケットの推力制御応答に関するシミュレーション
津地 歩, 山田 藍, 深田 真衣, 脇田 督司, 永田 晴紀, 年次大会, 2020, J19109, 2020年
Axial-injection end-burning hybrid rocket (EBHR) is excellent in fuel regression rate and throttling characteristics than conventional hybrid rockets. However, in the firing tests of throttling operation using multiple port fuel, a very long response time was observed. In this research, the authors postulated that the fuel regression shape of each port might be affected by chamber pressure to explain the reason for long response time and verified this hypothesis by using numerical simulation. The calculation result showed the long response time, and the response time coincides with the shape-changing time, which indicates that the shape-changing causes a long response time. Moreover, the history of characteristic exhaust velocity suggests that the fuel-rich condition may make response time short., 一般社団法人 日本機械学会, 日本語 - 亜酸化窒素-PMMA燃料における燃え拡がり-安定燃焼の遷移条件
深田 真衣, 佐藤 元紀, 津地 歩, 永田 晴紀, 脇田 督司, 年次大会, 2020, J19110, 2020年
When designing axial- injection end-burning hybrid rocket (EBHR), it is essential to obtain the boundary between stabilized combustion mode and flame spreading mode. Oxidizer friction velocity has been used as an indicator of them. The threshold value has obtained in the case with PMMA fuel and O2 as a constant value and used when designing EBHR. Recently, however, N2O is said to be useful for oxidizer, so we have to research if we can use N2O for EBHR. In this study, we aim to obtain the threshold value of friction velocity with PMMA fuel and N2O as the oxidizer. We conducted some combustion experiments and explored the relationship between friction velocity, chamber pressure, and flame behavior. As a result, we observed pressure dependencies of the threshold value of friction velocity, which is different from the result of the O2 experiment. There are two possibilities why such a result arose, 1) difference of oxidizer chemical species, 2) development of boundary layer., 一般社団法人 日本機械学会, 日本語 - ハイブリッドロケットにおける燃焼データ解析手法の開発
永田晴紀, KAMPS Landon, 齋藤勇士, 燃焼シンポジウム講演論文集(CD-ROM), 58th, 2020年 - ハイブリッドロケットの点火器に関する研究
池田華子, KAMPS Landon, 永田晴紀, 燃焼シンポジウム講演論文集(CD-ROM), 58th, 2020年 - 亜酸化窒素を用いた端面燃焼式ハイブリッドロケットの燃料後退機構
奥田椋太, 小水弘大, 津地歩, 三輪拓実, 横堀秀一, 添田建太郎, 永田晴紀, 燃焼シンポジウム講演論文集(CD-ROM), 58th, 2020年 - ハイブリッドロケットの軸方向後燃料退速度に関する3Dスキャナを用いた研究
奥田晃崇, KAMPS Landon, 伊藤聖司, 影山理沙, VISCOR Tor, 井上翔太, 池田華子, 高田祐亮, 脇田督司, 永田晴紀, 燃焼シンポジウム講演論文集(CD-ROM), 58th, 2020年 - Visualization of fuel regression rate in axial-injection end-burning hybrid rocket
Takumi Miwa, Ayumu Tsuji, Ryota Okuda, Shuichi Yokobori, Kentaro Soeda, Landon Kamps, Harunori Nagata, AIAA Propulsion and Energy 2020 Forum, 1, 11, 2020年
© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Reconstruction techniques determined the fuel regression rate in previous studies relevant to axial-injection end-burning hybrid rocket. However, the applicability of these techniques remains unclear. This study aimed to investigate whether the fuel regression rate determined using reconstruction techniques matches the actual flame traveling velocity. A visualization chamber, which had a window made of PMMA as an optical path, was used as a method for measuring the flame traveling velocity directly. Several firing tests revealed that if the ratio of pressure-unsteady-state duration to overall firing duration was small enough, the regression rate calculated by reconstruction technique matches well with the flame traveling velocity determined by visualization. However, when the pressure-unsteady-state duration was relatively long, the reconstruction technique overestimates the fuel regression rate. This is because the pressure-unsteady-state regime is characterized by an increasing characteristic exhaust velocity efficiency. This result suggests that the pressure exponent for the fuel regression rate under high chamber pressure is smaller than the previously reported value, n = 1.20 [1]. - Evaluation of the thermal onset of graphite nozzle erosion
Seiji Ito, Landon Kamps, Satoshi Yoshimaru, Harunori Nagata, AIAA Propulsion and Energy 2020 Forum, 1, 14, 2020年
© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. In this research, ten static firing tests were conducted using a 100 N-thrust class hybrid rocket motor with a water-cooled nozzle. Chamber pressure, equivalence ratio, and nozzle inner wall temperature were varied to evaluate the thermal onset of graphite nozzle erosion. The oxidizer was liquid nitrous oxide or gaseous oxygen, and fuel was high-density polyethylene. Results show that cooling the nozzle was effective in reducing the nozzle erosion rate and delaying the erosion onset time. In three of the ten tests, water-cooling completely prevented nozzle erosion. Furthermore, experimental results with various nozzle wall temperature supported the hypothesis of a chemical-kinetic-limited nozzle erosion onset threshold. Although chamber pressure affects the threshold of the nozzle erosion onset factor, a rough generalization that the onset of nozzle erosion takes place at temperatures around 1500 K holds. Most remarkably, in one test, nozzle erosion is shown to have been prevented even when the nozzle temperature remained above 1500 K. - 地球磁気圏X線撮像計画GEO-X(GEOspace X-ray imager)の現状 II
江副祐一郎, 船瀬龍, 永田晴紀, 三好由純, 笠原慧, 中嶋大, 三石郁之, 石川久美, 上野宗孝, 山崎敦, 長谷川洋, 三田信, 満田和久, 藤本正樹, 川勝康弘, 岩田隆浩, 平賀純子, 小泉宏之, 佐原宏典, 金森義明, 森下浩平, 日本天文学会年会講演予稿集, 2020, 2020年 - 地球磁気圏X線撮像計画GEO-X(GEOspace X-ray imager)の現状 III
江副祐一郎, 船瀬龍, 永田晴紀, 三好由純, 笠原慧, 中嶋大, 三石郁之, 石川久美, 上野宗孝, 山崎敦, 長谷川洋, 三田信, 藤本正樹, 川勝康弘, 岩田隆浩, 満田和久, 平賀純子, 小泉宏之, 佐原宏典, 金森義明, 森下浩平, 沼澤正樹, 日本天文学会年会講演予稿集, 2020, 2020年 - 地球磁気圏X線撮像計画GEO-Xの現状
江副祐一郎, 船瀬龍, 永田晴紀, 三好由純, 中嶋大, 三石郁之, 石川久美, 山崎敦, 長谷川洋, 藤本正樹, 上野宗孝, 地球電磁気・地球惑星圏学会総会及び講演会(Web), 148th, 2020年 - 亜酸化窒素-光硬化性樹脂単ポート燃料の安定燃焼特性
深田真衣, 津地歩, 奥田椋太, 山田藍, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 64th, 2020年 - グラファイトノズル浸食の熱的開始条件に関する研究
伊藤聖司, KAMPS Landon, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 64th, 2020年 - ハイブリッドキックモータを搭載した超小型宇宙機の熱設計
田端健一, 友永優太, 戸谷剛, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 64th, 2020年 - Initial Firing Tests of Aluminum Rod/Water Hybrid Rockets
Yuji Saito, Landon Kamps, Ayumu Tsuji, Masashi Wakita, Hiroyuki Koizumi, Keisuke Asai, Harunori Nagata, AIAA Propulsion and Energy 2020 Forum, 1, 11, 2020年, [査読有り]
© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. The combustion of aluminum and water is of relevance to many propulsion and energy conversion applications. The authors propose a concept for realizing aluminum/water hybrid rockets using PMMA (polymethyl methacrylate)/oxygen hybrid rocket combustion for heating. Three experiments were conducted to examine aluminum/water combustion, and the feasibility of heating by PMMA/oxygen combustion. The results show that the melting and combustion of rod-aluminum were achieved, however the aluminum/water reaction could not be confirmed before combustion extinction., American Institute of Aeronautics and Astronautics - アブレータ材料によるハイブリッドロケットノズル浸食抑制に関する研究
奥田 晃崇,Landon T. Kamps,櫻井 和人, 井上 卓,Tor Viscor,内山 絵里香,池田 華子, 吉丸 利,脇田 督司,永田 晴紀, 第57回燃焼シンポジウム講演論文集, 57th, 2019年11月, [最終著者] - 冷却によるグラファイトノズルの浸食抑制効果
吉丸 利,永田 晴紀,伊藤 聖司,Landon T. Kamps, 第57回燃焼シンポジウム講演論文集, 2019年11月, [最終著者] - 端面燃焼式ハイブリッドロケットにおける燃焼室特性長さが及ぼす c*効率への影響
押見 灯里,津地 歩,小水 弘大,添田 建太郎,横堀 秀一, 永田 晴紀, 第57回燃焼シンポジウム講演論文集, 2019年11月 - 亜酸化窒素を用いた端面燃焼式ハイブリッドロケットの推力制御特性
小水 弘大,奥田 椋太,津地 歩,押見 灯里, 添田 建太郎,横堀 秀一,永田 晴紀, 第57回燃焼シンポジウム講演論文集, 2019年11月, [最終著者] - デトネーション波のセル規則性が拡大環状流路における伝播特性に与える影響
本谷 誠正,桧物 恒太郎,脇田 督司,永田 晴紀, 第57回燃焼シンポジウム講演論文集, 57th, 2019年11月, [最終著者] - 地球磁気圏X線撮像計画GEO-X(GEOspace X-ray imager)の現状
江副祐一郎, 三好由純, 笠原慧, 船瀬龍, 永田晴紀, 中嶋大, 三石郁之, 石川久美, 山崎敦, 長谷川洋, 三田信, 満田和久, 藤本正樹, 川勝康弘, 岩田隆浩, 上野宗孝, 平賀純子, 小泉宏之, 佐原宏典, 日本天文学会年会講演予稿集, 2019, 2019年 - Fuel regression characteristics in hybrid rockets using n2o/hdpe
Seiji Ito, Landon Kamps, Kazuhito Sakurai, Lisa Kageyama, Terutaka Okuda, Harunori Nagata, AIAA Propulsion and Energy Forum and Exposition, 2019, 2019年
© 2019 by American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Twenty nine static firing tests were conducted under varying experimental conditions to construct fuel regression rate formulas, invesgate the effect of radiant heat on fuel regression rate and confirm the validity of the integral method for fuel regression rate. Results show that the conventional fuel regression formula does not predict regression rate accurately. Furthermore, when treating the regression rate coefficient as a function of equivalence ratio, there is no local maxium near the stoichiometric mixture ratio as expected of adiabatic flame temperature. Five empirical models are tested, of which models 3 – 5 improve the accuracy of the fuel regression rate formula, model 4 being the most accurate of all models. Empirical model 4 predicts the fuel regression rate in multiple tests to within ± 10 %. - Development of N2O/HDPE hybrid rocket for microsatellite propulsion
Landon Kamps, Pau Molas-Roca, Erica Uchiyama, Tomohiro Takanashi, Harunori Nagata, Proceedings of the International Astronautical Congress, IAC, 2019-October, 2019年
Copyright © 2019 by the International Astronautical Federation (IAF). All rights reserved. This research compares the performance of three alternative hybrid rocket propulsion systems using liquid nitrous oxide and/or gas oxygen as the oxidizer for use as microsatellite thrusters. Internal ballistic performance predictions are based on recent experimental research and codified using the MATLAB App designer for open release. A self-pressurizing liquid nitrous oxide hybrid rocket thruster is shown to outperform a high-pressure gas oxygen hybrid rocket thruster in ?V (1412 m/s versus 1326 m/s) and initial (wet) mass (95 kg versus 107 kg), even though time-averaged Isp is smaller (300 s versus 317 s). This is because the gas oxygen system requires very high-pressure storage vessels (> 60 MPa) and a larger fuel grain which results in a heavier hybrid rocket motor. Nozzle erosion is included in this analysis, and shown to result in a positive feedback of oxidizer flow rate relative to fuel flow rate that prevents Isp from decreasing in the liquid nitrous oxide thruster. The high-performing liquid nitrous oxide hybrid rocket thruster is produced in CAD, and shown to have adequate space for mounting all necessary sub-systems and payload modules, with room for future adjustments as necessary. - Effect of aft chamber volume on hybrid rocket combustion efficiency
L. Kageyama, L. Kamps, H. Nagata, AIAA Propulsion and Energy Forum and Exposition, 2019, 2019年
© 2019 by American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Hybrid rocket motors are attractive propellant systems due to their advantages over their solid and liquid counterparts, but their low combustion efficiency has been a large hinderance to its application to actual space projects. In this study, the effect of characteristic chamber length L*, a parameter related to residence time, and its effects on c* efficiency is experimentally investigated. A total of 7 static firing tests using 100N-scale hybrid rocket motors using high-density polyethylene as fuel and liquid nitrous oxide as oxidizer were performed, with changing inert aft chamber volumes and nozzle throat diameters. A data reconstruction technique was used to obtain the time histories of c* efficiency and other parameters. The results show a slightly positive correlation between L* and c* efficiency; however, the effects of mixing and improved fuel regression due to the addition of aft chambers appeared to be more influential to the improvement of efficiency. - 超小型深宇宙探査機用ハイブリッドロケットキックモータの開発
伊藤 聖司, 内山 絵里香, 櫻井 和人, ケンプス ランドン, 影山 理沙, 永田 晴紀, 年次大会, 2019, 0, J19105P, 2019年There are few opportunities to launch a satellite for deep space exploration, because of the low budget in Japan. A piggy-back microsatellite takes lower cost than a satellite launched alone. A microsatellite can go GTO with the main satellite, and then the microsatellite goes to deep space using a kick motor. To achieve the thrust and specific impulse that the kick motor requires, we need chemical rockets. Hybrid rockets are the best choice from the viewpoint of safety. Nitrous oxide is selected as the oxidizer because the kick motor also requires storable oxidizer. The relationship between L* and c* efficiency, the fuel regression rate formula and the characteristic of nozzle erosion are obtained from static firing tests. Furthermore, the change in velocity is calculated for each size of the kick motor.
, 一般社団法人 日本機械学会, 日本語 - EXPERIMENTAL INVESTIGATION OF C* EFFICIENCY IN NITROUS OXIDE HYBRID ROCKETS
Erika Uchiyama, Yurika Kiyotani, Landon Kamps, Harunori Nagata, PROMOTE THE PROGRESS OF THE PACIFIC-BASIN REGION THROUGH SPACE INNOVATION, 166, 109, 115, 2019年, [査読有り]
Hybrid Rockets have advantages of low cost and high safety but there are few practical uses at the current state of the art. The combustion characteristics of N2O, which is very useful oxidizer, have not been researched in particular. This study is the investigation to clarify the dependency of the c* (characteristic exhaust velocity) efficiency eta(c)* in nitrous oxide (N2O) hybrid rockets on operating conditions through experimentation. Several firing tests were conducted using a 200N thrust class conventional hybrid rocket motor employing high density polyethylene (HDPE) as the fuel and liquid nitrous oxidizer as the oxidizer. The results reveal that there is no clear dependency of eta(c)* on mixture ratio, pressure or characteristic length, suggesting that efficiency must be improved through other design parameters., UNIVELT INC, 英語 - Investigation of graphite nozzle erosion in hybrid rockets using N2O/HDPE
Landon Kamps, Kazuhito Sakurai, Kohei Ozawa, Harunori Nagata, AIAA Propulsion and Energy Forum and Exposition, 2019, 2019年, [査読有り]
© 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Seventeen static firing tests are carried out on a small-scale hybrid rocket motor using liquid nitrous oxide as the oxidizer to investigate the chemical erosion characteristics of graphite nozzles. Over 200 data are obtained by employing an innovative data reduction method to determine time-resolved values for nozzle throat diameter, nozzle throat pressure, equivalence ratio, nozzle throat wall temperature and more. An analytical model is formed based on previous research, and used to develop an informed empirical formula for experimental correlations. Empirical correlations are shown to predict nozzle erosion rates with a coefficient of determination upwards of 0.81, whereas the analytical model results in a coefficient of determination of only 0.26. Nozzle erosion rates reached upwards of 0.25 mm/s for equivalence ratios around unity, and chamber pressures around 5 MPa. Lastly, the conditions at the onset of erosion are examined, and used to demonstrate, quantitatively, that chemical erosion can be more easily mitigated when using nitrous oxide as the oxidizer than oxygen., 英語 - 希釈混合気へ伝播する半球状デトネーション波の伝播限界
中西勇作, 桧物恒太郎, 本谷誠正, 脇田督司, 永田晴紀, 燃焼シンポジウム講演論文集, 56th, ROMBUNNO.B214, 2018年11月14日
日本語 - 液滴ラジエータにおける液滴流の実効放射率の測定
両門健人, 高梨知広, 戸谷剛, 脇田督司, 永田晴紀, Thermophysical Properties, 39th, 166‐168, 2018年11月13日
日本語 - 液中レーザー溶融法におけるナノスケールパルス加熱による有機溶媒の熱分解挙動
末原 健太朗, 石川 善恵, 越崎 直人, 尾村 和信, 永田 晴紀, 応用物理学会学術講演会講演予稿集, 2018.2, 775, 775, 2018年09月05日
公益社団法人 応用物理学会, 日本語 - ハイブリッドロケットの燃焼データ解析法の開発とその応用
永田, 晴紀, ケンプス, ランドン, 齋藤, 勇士, Nagata, Harunori, Kamps, Landon, Saito, Yuji, 第1回ハイブリッドロケットシンポジウム 講演集 = Proceedings of the 1st Hybrid Rocket Symposium, 2018年06月
第1回ハイブリッドロケットシンポジウム(2018年6月28日-29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
1st Hybrid Rocket Symposium (June 28-29, 2018. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan
資料番号: SA6000126019
レポート番号: HR-2018-Keynote2, 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS), 日本語 - 亜酸化窒素を用いたハイブリッドロケットの特性排気速度効率—c* efficiency in nitrous oxide hybrid rockets
内山, 絵里香, 清谷, 優理香, Kamps, Landon, 永田, 晴紀, Uchiyama, Erika, Kiyotani, Yurika, Nagata, Harunori, 第1回ハイブリッドロケットシンポジウム 講演集 = Proceedings of the 1st Hybrid Rocket Symposium, 2018年06月
第1回ハイブリッドロケットシンポジウム(2018年6月28日-29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
1st Hybrid Rocket Symposium (June 28-29, 2018. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan
資料番号: SA6000126013
レポート番号: HR-2018-012, 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS), 日本語 - 地球磁気圏X線撮像計画GEO-Xの現状
江副祐一郎, 三好由純, 笠原慧, 船瀬龍, 永田晴紀, 上野宗孝, 中嶋大, 木村智樹, 石川久美, 三田信, 満田和久, 藤本正樹, 大橋隆哉, 日本天文学会年会講演予稿集, 2018, 2018年 - ミニカムイプロジェクトの変遷に見るプロジェクト教育の効果と課題
永田晴紀, 高梨知広, 脇田督司, 宇宙科学技術連合講演会講演集(CD-ROM), 62nd, ROMBUNNO.1S05, 2018年
日本語 - アクリル燃料を用いたCAMUI型ハイブリッドロケット内における燃焼場の可視化に関する研究
井上卓, KAMPS Landon, 櫻井和人, 内山絵里香, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 62nd, ROMBUNNO.1N05, 2018年
日本語 - 衛星全体の熱容量を用いた熱設計方針の適用範囲の拡大
神谷朋兆, 戸谷剛, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 62nd, ROMBUNNO.2F19, 2018年
日本語 - 超小型深宇宙探査機用ハイブリッドキックモータの開発
永田晴紀, 清谷優理香, 櫻井和人, 脇田督司, 戸谷剛, 衝撃波シンポジウム講演論文集(CD-ROM), 2017, ROMBUNNO.2A1‐2, 2018年
To realize deep space missions on a limited budget, a hybrid rocket kick motor for a laboratory-scale space probe being kicked from GTO to a deep space trajectory is being developed. Nitrous Oxide (N2O) is selected as an oxidizer to allow for long-term non-cryogenic storage in space, and High Density Polyethylene (HDPE) is selected as a fuel for safety and affordability.Two fuel configurations are being considered for this mission: a conventional tubular fuel grain, and a Cascaded Multistage Impinging-jet (CAMUI) type fuel grain. This study presents results and analysis from the first set of CAMUI-type static firing tests employing N2O as an oxidizer, with the aim of clarifying performance related issues such as achieving an optimum oxidizer to fuel mass ratio O/F of around 6~7, quantifying fuel regression rate etc. for design purposes. The results of two gaseous N2O firing tests and two liquid N2O firing test lead to the following two conclusions. The use of liquid N2O will be necessary to realize an appropriate flowrate for space applications, and CAMUI-type motors may not be suitable for achieving optimum O/F because they tend to burn extremely fuel rich. Additional firing tests are necessary to quantify the fuel regression rate in CAMUI-type fuel and conventional tubular fuel grains using N2O., 一般社団法人 日本機械学会, 日本語 - Investigation of graphite nozzle erosion in hybrid rockets using o2/c2h4
Landon Kamps, Shota Hirai, Kazuhito Sakurai, Tor Viscor, Yuji Saito, Raymond Guan, Hikaru Isochi, Naoto Adachi, Mitsunori Itoh, Harunori Nagata, 2018 Joint Propulsion Conference, 2018年01月01日
© 2018 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. A recently developed reconstruction technique titled nozzle-throat reconstruction technique is used to investigate graphite nozzle-throat-erosion in two scales of hybrid rocket motors, 30N-thrust class and 2000N-thrust class, using oxygen as the oxidizer and high density polyethylene as the fuel. Thirty seven static firing tests were conducted under varying experimental conditions to confirm the validity of the reconstruction technique results, investigate the conditions at the onset of erosion and to formulate an empirical predictive model of nozzle erosion rate. Results show that nozzle erosion increases the convective heat transfer coefficient to upwards of 2-4 times the value predicted by Bartz correlation. Furthermore, an empirical model is introduced that treats the combustion gas as a single oxidizing agent with heterogeneous rate constants that are distributions of equivalence ratio of the bulk fluid flow. This empirical model predicts the nozzle throat erosion histories in multiple tests to within ± 5%. - Investigation of regression rate under high-pressure in axial-injection end-burning hybrid rockets
Yushi Okutani, Yuji Saito, Masaya Kimino, Ayumu Tsuji, Kentaro Soeda, Harunori Nagata, 2018 Joint Propulsion Conference, 2018年01月01日
© 2018 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. This study is an investigation of preeminent fuel regression rate in Axial-injection End-burning hybrid rockets under high pressure conditions. The authors overcame a back firing problem at high pressure conditions by redesigning and optimizing the fuel port diameter. The fuel grain outer diameter was 38 mm and the port diameter was 0.4 mm. Firing tests were conducted using gaseous oxygen as oxidizer at chamber pressures range from 0.98 MPa to 1.44 MPa. The results of four static firing tests show that fuel regression rate increases as the chamber pressure increases, which is consistent with previous research results. Fuel regression rate reached approximately 12.6 mm/s at 1.44 MPa. The authors reformulated the relationship between fuel regression rate and chamber pressure based on the results of high chamber pressure region, and showed that the pressure exponent increased from 1.09 to 1.20. - The accuracy of reconstruction techniques for determining hybrid rocket fuel regression rate
Yuji Saito, Landon Kamps, Kodai Komizu, Kentaro Soeda, Daniele Bianchi, Francesco Nasuti, Harunori Nagata, 2018 Joint Propulsion Conference, 2018年01月01日
© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. This study is an investigation of the accuracy of reconstruction techniques for determining instantaneous fuel regression rate. Results of reconstruction techniques are compared with results obtained through the measurement of the pressure drop across the fuel in an Axial-Injection End-Burning hybrid rocket (EBHR). The results of numerous firing tests show that this method allows for the evaluation of the accuracy of the instantaneous fuel regression rates obtained by reconstruction techniques. The error bias of O/F values calculated by the reconstruction techniques were around ±10%, and were mainly caused by uncertainties in the measured values of oxidizer mass flowrate and the definition of firing duration. The instantaneous length of an EBHR-type fuel can be calculated from the measurement of the pressure drop across the fuel. However, the calculated fuel length history obtained by the pressure drop in a port does not coincide with that obtained by the reconstruction technique because of an underestimation in pressure drop. - Control of heat release energy during heating from solid-solid phase change material
Ryohei Gotoh, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata, International Heat Transfer Conference, 2018-August, 4425, 4432, 2018年
© 2018 International Heat Transfer Conference. All rights reserved. The use of phase change materials (PCMs) as a heat source and storage has become an important consideration in energy management. 2-amino-2-methyl-1,3-propanediol (AMP), which is a solid-solid PCM, stores approximately 264 J/g of heat energy at approximately 78 °C. AMP has the attractive characteristic of storing heat energy in its non-equilibrium solid state. This study clarified this characteristic to control the stored heat energy of AMP so that it can be used at when heat energy is needed. The thermal property was measured by differential scanning calorimetry (DSC) in the temperature range of -50 °C to 100 °C, and the structural property was measured through X-ray powder diffraction in the temperature range of -70 °C to 90 °C. The results revealed that the material's initial crystal structure changed from brittle crystalline (phase II) to plastic crystalline (phase I) at 80 °C during heating from room temperature to 90 °C. Phase I remained almost unchanged during cooling from 90 °C to -50 °C, but changed to the non-equilibrium state (phase Ig'). Then, phase Ig' changed to phase II with exoergic heat of 140 J/g in the temperature range between -14 °C and 36 °C during heating. This study clarified that the phase change temperature can be controlled by controlling the sample mass or by applying an external stimulus. In other words, AMP is a rare solid-solid PCM that does not release heat energy during cooling, but rather releases heat energy during heating. Moreover, the heat energy is controllable. This positive attribute makes AMP a good candidate for use in a heating assist system. - Experimental investigation of c* efficiency in nitrous oxide hybrid rockets
Erika Uchiyama, Yurika Kiyotani, Landon Kamps, Harunori Nagata, Advances in the Astronautical Sciences, 166, 109, 115, 2018年01月01日
© 2018 Univelt Inc. All rights reserved. Hybrid Rockets have advantages of low cost and high safety but there are few practical uses at the current state of the art. The combustion characteristics of N2O, which is very useful oxidizer, have not been researched in particular. This study is the investigation to clarify the dependency of the c* (characteristic exhaust velocity) efficiency ηc* in nitrous oxide (N2O) hybrid rockets on operating conditions through experimentation. Several firing tests were conducted using a 200N thrust class conventional hybrid rocket motor employing high density polyethylene (HDPE) as the fuel and liquid nitrous oxidizer as the oxidizer. The results reveal that there is no clear dependency of ηc* on mixture ratio, pressure or characteristic length, suggesting that efficiency must be improved through other design parameters. - デトネーション波の反射・再開始による起爆エネルギー評価法に関する研究
松岡将司, 大関敦, 桧物恒太郎, 脇田督司, 戸谷剛, 永田晴紀, 燃焼シンポジウム講演論文集, 55th, 386‐387, 2017年11月13日
日本語 - 共振器のQ値がマイクロキャビティによる放射波長制御に与える影響
佐藤潤弥, 戸谷剛, 脇田督司, 永田晴紀, Thermophysical Properties, 38th, 111‐113, 2017年11月01日
日本語 - 端面燃焼式ハイブリッドロケットの推力制御時における時間応答性に関する研究
齋藤 勇士, 君野 正弥, 津地 歩, 尾村 和信, 安河内 裕之, 添田 建太郎, 戸谷 剛, 脇田 督司, 永田 晴紀, 年会講演会講演集, 48, 8p, 2017年04月13日
日本航空宇宙学会, 日本語 - ハイブリッドロケットのc*効率について
片野田 洋, 永田 晴紀, 鹿児島大学工学部研究報告, 58, 58, 1, 6, 2017年03月
A c* efficiency of a hybrid rocket calculated by reconstruction technique, ηrec, is compared with the traditional c* efficiency, ηtrad, obtained by the time-averaged O/F ratio. Simulation data of combustion pressure and mass flow rate of oxidizer are provided for different three cases to calculate two types of c* efficiency. The calculated results show that 1) ηrec and ηtrad are almost equal when the theoretical characteristic exhaust velocity, c*th, varies almost linearly against the variation of O/F ratio during the firing, 2) ηrec is greater than ηtrad when c*th varies non-linearly against the variation of O/F ratio, 3) ηrec is appropriate as a c* efficiency of a hybrid rocket., 鹿児島大学, 日本語 - ハイブリッドロケットに関する燃焼技術
永田 晴紀, 日本燃焼学会誌, 59, 190, 243, 252, 2017年
This paper summarizes researches about combustion technologies related to hybrid rockets, focusing on the enhancement of regression rates of solid fuels and the development of techniques for measuring fuel gasification rate. Methods for enhancing the regression rate were divided into two groups: improving the physical or chemical properties of a solid fuel, and improving the combustion flow field in a combustion chamber. In the former group, the use of liquefying fuels like paraffin-wax has become the mainstream of recent hybrid rocket development activities because these fuels have been shown to provide regression rates that are two to three times higher than values encountered with conventional solid fuels. Measuring the history of the fuel gasification rate in a hybrid rocket is not easy, and it has been an important subject in this field. Measurement techniques fall into two categories: the direct measurement of fuel thickness or fuel weight, and reconstruction techniques that estimate fuel flow rate through easily measurable histories such as liquid oxidizer flow rate, chamber pressure, thrust, etc. Recently, a new method to estimate the nozzle throat area history during firing was developed by improving on a reconstruction technique., 一般社団法人 日本燃焼学会, 日本語 - スペースプレーンに必要な技術のためのFTBのシステムと飛行軌道の検討
丸祐介, 澤井秀次郎, 永田晴紀, 小林弘明, 坂東信尚, 吉光徹雄, 江口光, 宇宙科学技術連合講演会講演集(CD-ROM), 61st, 2017年 - 超小型探査機による惑星探査技術の開発
杉田精司, 鈴木宏二郎, 中須賀真一, 吉川一朗, 今村剛, 船瀬龍, 小泉宏之, 笠原慧, 吉岡和夫, 永田晴紀, 藤本正樹, 日本惑星科学会秋季講演会予稿集(Web), 2017, 2017年 - 超小型深宇宙探査機用ハイブリッドキックモータの開発
清谷 優理香, 平井 翔大, ケンプス ランドン, 山口 亮, 櫻井 和人, 脇田 督司, 戸谷 剛, 永田 晴紀, 年次大会, 2017, 0, S1920105, 2017年, [最終著者]To realize deep space missions on a limited budget, a hybrid rocket kick motor for a laboratory-scale space probe being kicked from GTO to a deep space trajectory is being developed. Nitrous Oxide (N2O) is selected as an oxidizer to allow for long-term non-cryogenic storage in space, and High Density Polyethylene (HDPE) is selected as a fuel for safety and affordability.Two fuel configurations are being considered for this mission: a conventional tubular fuel grain, and a Cascaded Multistage Impinging-jet (CAMUI) type fuel grain. This study presents results and analysis from the first set of CAMUI-type static firing tests employing N2O as an oxidizer, with the aim of clarifying performance related issues such as achieving an optimum oxidizer to fuel mass ratio O/F of around 6~7, quantifying fuel regression rate etc. for design purposes. The results of two gaseous N2O firing tests and two liquid N2O firing test lead to the following two conclusions. The use of liquid N2O will be necessary to realize an appropriate flowrate for space applications, and CAMUI-type motors may not be suitable for achieving optimum O/F because they tend to burn extremely fuel rich. Additional firing tests are necessary to quantify the fuel regression rate in CAMUI-type fuel and conventional tubular fuel grains using N2O.
, 一般社団法人 日本機械学会, 日本語 - 端面燃焼式ハイブリッドロケットの推力制御時におけるヒステリシス特性に関する研究
君野正弥, 齋藤勇士, 津地歩, 尾村和信, 安河内裕之, 添田建太郎, 戸谷剛, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 61st, ROMBUNNO.3H01, 2017年
日本語 - 気体亜酸化窒素を酸化剤とした端面燃焼式ハイブリッドロケットの燃料後退特性
尾村和信, 津地歩, 齋藤勇士, 君野正弥, 奥谷勇士, 小水弘大, 戸谷剛, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 61st, ROMBUNNO.3H04, 2017年
日本語 - 液体酸素を用いた端面燃焼式ハイブリッドロケットの実現可能性
津地歩, 齋藤勇士, 尾村和信, 君野正弥, 戸谷剛, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 61st, ROMBUNNO.3H03, 2017年
日本語 - 酸化剤にN2Oを用いたハイブリッドロケットにおける燃焼室圧力と特性排気速度効率の関係について
清谷優理香, 山口亮, 櫻井和人, KAMPS Landon, 井上卓, 脇田督司, 戸谷剛, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 61st, ROMBUNNO.3H02, 2017年
日本語 - ふく射センサの低温環境下でのふく射率および単一液滴流からの排熱量
嶋田泰三, 高梨知広, 両門健人, 戸谷剛, 永田晴紀, 脇田督司, 日本伝熱シンポジウム講演論文集(CD-ROM), 54th, ROMBUNNO.C332, 2017年
日本語 - 2‐amino‐2‐methyl‐1,3‐propanediolの固相‐固相結晶転移による潜熱を利用した蓄熱材の開発 過冷却状態の結晶化による発熱とガラス転移点の関係
後藤凌平, 永田晴紀, 戸谷剛, 脇田督司, 日本伝熱シンポジウム講演論文集(CD-ROM), 54th, ROMBUNNO.D124, 2017年
日本語 - 小型衛星放出機構から放出される50kg級衛星の熱設計
PANES Mitchao Delburg, 戸谷剛, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 61st, ROMBUNNO.2F21, 2017年
日本語 - Thermal Design and On-orbit Validation of the First Philippine Micro-satellite: DIWATA-1
Delburg Mitchao, Tsuyoshi Totani, Yuji Sakamoto, Masashi Wakita, Harunori Nagata, Proceedings of 47th International Conference on Environmental Systems, ICES-2017-130, 2017年, [査読有り]
英語 - Investigation of graphite nozzle-throat-erosion in a laboratory-scale hybrid rocket using GOX and HDPE
Landon Kamps, Shota Hirai, Yassine Ahmimache, Raymond Guan, Harunori Nagata, 53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017, 2017年01月01日
The authors of this paper employ a recently developed reconstruction technique titled nozzle-throat reconstruction technique to investigate graphite nozzle-throat-erosion in a laboratory-scale hybrid rocket motor using gaseous oxygen as an oxidizer and high density polyethylene as a fuel. Fifteen static firing tests were conducted under varying experimental conditions to confirm the validity of the reconstruction technique results, and to collect a wide range of nozzle-throat-erosion data. Furthermore, a technique for carrying out classical finite difference calculations for 1D convective and conductive heating based on the time histories of gas properties as determined by the reconstruction technique is introduced and used to estimate nozzle throat wall temperature history and convective heat transfer coefficient history. Results show a distinct trend where nozzle erosion rates increase in the beginning of a firing test, and subsequently decrease for the remainder of the firing test, even though nozzle throat temperatures continue to increase. It is shown that the decreasing nozzle-throat-erosion rates coincide with decreasing mass fluxes and that the erosion rates in this regime may be sensitive to oxidizer to fuel mass ratio. - Investigation of regression characteristics under relatively high-pressure in Axial-injection End-Burning Hybrid Rockets
Yuji Saito, Masaya Kimino, Tsuji Ayumu, Yushi Okutani, Kazunobu Omura, Hiroyuki Yasukochi, Kentaro Soeda, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata, 53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017, 2017年01月01日
This study is an investigation of Axial-Inection End-Burning Hybrid Rockets aimed at revealing fuel regression characteristics under relatively high pressure conditions. Firing tests were conducted using gaseous oxygen as oxidizer at chamber pressures and oxidizer port velocities ranging from 0.22 MPa to 1.05 MPa and 31 m/s to 103 m/s, respectively. The results of fifteen static firing tests show that fuel regression rate increases as the chamber pressure increases, and regression rates ranged from approximately 1.1 mm/s at 0.25 MPa to 5.4 mm/s at 0.71 MPa. Furthermore, it is observed that the fuel regression rate is not influenced by oxidizer port velocity. The athours encountered a problem refered to as backfiring in this paper, and developed a calculation model to investigate this problemanalytically. The calculation model explains why the back-firing problem tends to occur in relatively high-pressure conditions, and leads to the conclusion that increasing nozzle throat diameter is an effective means of preventing back-firing from occuring. - Numerical and experimental investigation of nozzle thermochemical erosion in hybrid rockets
Daniele Bianchi, Landon Kamps, Francesco Nasuti, Harunori Nagata, 53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017, 2017年01月01日
Despite some inherent disadvantages, hybrid rockets are today considered as having a great potential to become one of the future generation of propulsion systems, partly due to recent emphasis on propulsion safety, reliability, low development cost, reduced environmental pollution, and greater operability. Nevertheless, the hybrid rocket development has not achieved the same level of maturity as solid and liquid traditional systems. An aspect that has not been much dealt with in the open literature is that of nozzle erosion, whose minimization or reduction is one of the challenges in hybrid rocket propulsion. To this goal, a joint numerical and experimental investigation of nozzle throat erosion has been performed using a computational fluid dynamics approach compared to static firing tests carried out on a 2kN-class lab-scale hybrid rocket burning liquid oxygen and highdensity polyethylene. The numerical approach is able to capture the main features of the nozzle throat erosion behavior, fairly reproducing the throat erosion rate values and its dependence upon the oxidizer to fuel mixture ratio and motor chamber pressure. - Using equivalent burn time to improve regression characterization of the CAMUI type hybrid rocket engine
T. Viscor, H. Nagata, 53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017, 2017年01月01日
Burn time errors caused by various start-up transient effects have a large effect on the regression simulation model of the Cascaded Multi Impinging Jet hybrid rocket engine. This paper analyses these burn time errors and their effect on the regression simulations for short burn time engines. To address these the equivalent burn time is then defined as the time the engine would burn if it was burning at steady state level throughout the burn time to achieve the measured total impulse. The accuracy of the regression simulation with and without the use of equivalent burn time are then finally compared. Equivalent burn time alone without at the same time addressing other errors is found to be inadequate to clearly address the burn time errors. - 円筒デトネーション波の伝播限界へ流路幅と円筒波面の曲率が与える影響—Influence of channel width and curvature of cylindrical detonation wave on propagation limit
桧物, 恒太郎, 亀山, 頌太, 大関, 敦, 榎並, 聖也, 脇田, 督司, 戸谷, 剛, 永田, 晴紀, Himono, Tsunetaro, Kameyama, Shota, Ozeki, Atsushi, Enami, Takaya, Wakita, Masashi, Totani, Tsuyoshi, Nagata, Harunori, 宇宙航空研究開発機構特別資料: 第48回流体力学講演会/第34回航空宇宙数値シミュレーション技術シンポジウム論文集 = JAXA Special Publication: Proceedings of the 48th Fluid Dynamics Conference / the 34th Aerospace Numerical Simulation Symposium, JAXA-SP-16-007, 211, 216, 2016年12月27日
第48回流体力学講演会/第34回航空宇宙数値シミュレーション技術シンポジウム (2016年7月6日-8日. 金沢歌劇座), 金沢市, 石川
48th Fluid Dynamics Conference /the 34th Aerospace Numerical Simulation Symposium (July 6-8, 2016. The Kanazawa Theatre), Kanazawa, Ishikawa, Japan
To realize practical use of a Pulse Detonation Engine, reliable initiation of detonation waves is important and the amount of oxidizer for initiation is needed to decrease to get high specific impulse. So the authors propose a detonation initiator to solve these problems. In this initiator, detonation waves propagate around a reflector through some transition. First, planar detonation waves transit to cylindrical detonation waves. Second, cylindrical detonation waves transit to toroidal detonation waves. Finally, toroidal detonation waves propagate to planar detonation waves propagating large bore chamber. In the propagation and transition process, the propagation limit of cylindrical detonation waves is unclear even though there is some hypothesis. That is to say, cylindrical detonation waves are quenched by its curvature or the formation of their cell structure is inhibited by narrow channel. In this paper, cell sizes are correlated to threshold condition for quenching. The proportion of cell sizes to the channel width of planer channel and the radius of cylindrical detonation waves are researched. It became obvious that cylindrical detonation waves can propagate stably if the proportion of the cell size to the radius of cylindrical detonation waves is larger than 25 and have the potential to survive and propagate if the proportion of the cell size to the radius of cylindrical detonation waves is larger than 17.
形態: カラー図版あり
Physical characteristics: Original contains color illustrations
資料番号: AA1630031030
レポート番号: JAXA-SP-16-007, 宇宙航空研究開発機構(JAXA), 日本語 - 液滴ラジエータにおける単一液滴流からの排熱量
嶋田泰三, 高梨知広, 戸谷剛, 永田晴紀, 脇田督司, Thermophysical Properties, 37th, 96‐98, 2016年11月28日
日本語 - 液体酸素を用いた固体燃料管内燃え拡がりに関する研究
津地歩, 齋藤勇士, 横井俊希, 尾村和信, 嶋田泰三, 戸谷剛, 脇田督司, 永田晴紀, 燃焼シンポジウム講演論文集, 54th, C323, 2016年11月26日
日本語 - 端面燃焼式ハイブリッドロケットの燃料後退モデルに関する考察
齋藤勇士, 横井俊希, 津地歩, 尾村和信, 安河内裕之, 添田建太郎, 戸谷剛, 脇田督司, 永田晴紀, 燃焼シンポジウム講演論文集, 54th, P222, 2016年11月26日
日本語 - 燃料微小管内を燃え拡がる火炎への雰囲気圧力の影響
横井俊希, 齋藤勇士, 尾村和信, 津地歩, 安河内裕之, 添田建太郎, 脇田督司, 戸谷剛, 永田晴紀, 燃焼シンポジウム講演論文集, 54th, C324, 2016年11月26日
日本語 - デトネーション波の衝突・反射による遷移促進効果に関する研究
大関敦, 松岡将司, 桧物恒太郎, 脇田督司, 戸谷剛, 永田晴紀, 燃焼シンポジウム講演論文集, 54th, E341, 2016年11月26日
日本語 - 気球を利用したスペースプレーン技術実証機の研究
丸 祐介, 澤井 秀次郎, 永田 晴紀, 坂東 信尚, 坂井 真一郎, 吉光 徹雄, 江口 光, Maru Yusuke, Sawai Shujiro, Nagata Harunori, Bando Nobutaka, Sakai Shinichiro, Yoshimitsu Tetsuo, Eguchi Hikaru, 大気球シンポジウム: 平成28年度 = Balloon Symposium: 2016, 2016年11月
大気球シンポジウム 平成28年度(2016年11月1-2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: SA6000057016レポート番号: isas16-sbs-016, 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS), 日本語 - Measurement and prediction of local regression rate of solid burning with impinging oxidizer jet
Tsuneyoshi Matsuoka, Kyohei Kamei, Takafumi Yamazaki, Harunori Nagata, Yuji Nakamura, Thirteenth International Conference on Flow Dynamics (ICFD 2016), 2016年10月, [査読有り]
英語 - 高高度気球を利用した高速飛行FTBシステム
丸 祐介, 澤井 秀次郎, 坂東 信尚, 永田 晴紀, 吉光 徹雄, 坂井 真一郎, 後藤 健, 江口 光, 宇宙科学技術連合講演会講演集, 60, 6p, 2016年09月06日
日本航空宇宙学会, 日本語 - 二段式CAMUI型ハイブリッドロケットの飛行実験による空中発射運用のための液体酸素セトリング方法の実証
五十地 輝, 植松 努, 川端 良輔, 高梨 知広, 永田 晴紀, 宇宙科学技術連合講演会講演集, 60, 4p, 2016年09月06日
日本航空宇宙学会, 日本語 - 連続して月にスイングバイする双曲線軌道を用いた地球脱出エネルギーの増加
須田 俊太郎, 川勝 康弘, 澤井 秀次郎, 永田 晴紀, 戸谷 剛, 宇宙科学技術連合講演会講演集, 60, 5p, 2016年09月06日
日本航空宇宙学会, 日本語 - ハイブリッドロケットにおけるノズルスロート浸食取得方法の精度と適用性に関する研究
ケンプス ランドン, 齋藤 勇士, 平井 翔大, 永田 晴紀, 宇宙科学技術連合講演会講演集, 60, 6p, 2016年09月06日
日本航空宇宙学会, 英語 - 超小型衛星の短期開発を目指す熱設計法の妥当性の検証
戸谷 剛, 毛利 正宏, 脇田 督司, 永田 晴紀, 宇宙科学技術連合講演会講演集, 60, 6p, 2016年09月06日
日本航空宇宙学会, 日本語 - 端面燃焼式ハイブリッドロケットの推力制御特性に関する研究
齋藤 勇士, 横井 俊希, 安河内 裕之, 添田 建太郎, 戸谷 剛, 脇田 督司, 永田 晴紀, 宇宙科学技術連合講演会講演集, 60, 1, 6p, 18, 2016年09月06日
日本航空宇宙学会, 日本語 - Verification of Rapid Thermal Design Approach using Design and Flight Data of Hodoyoshi-1 Microsatellite
Tsuyoshi Totani, Hiroto Ogawa, Masashi Wakita, Harunori Nagata, Yusuke Kuramoto, Naoki Miyashita, Proceeding of 46th International Conference on Environmental Systems, ICES-2016-116, 2016年07月, [査読有り]
英語 - 端面燃焼式ハイブリッドロケットにおける燃料後退特性の圧力依存性に関する実験的および解析的考察
齋藤 勇士, 横井 俊希, 嶋田 泰三, 安河内 裕之, 添田 建太郎, 戸谷 剛, 脇田 督司, 永田 晴紀, 年会講演会講演集, 47, 9p, 2016年04月14日
© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. The regression characteristics of axial-injection end-burning hybrid rocket (EBHR) fuels having numerous small ports were experimentally investigated for the first time using a laboratory scale motor. In this paper, three requirements for EBHR fuel grains are explained in detail. High accuracy 3D printing allows for the production of fuel that satisfies the requirements for EBHR as defined in this paper. A data reduction method that overcomes the problem of multiple solutions to the c* equation is used to determine fuel regression rate with less than 10% error. Results of fifteen static firings tests show that fuel regression rate increases as the chamber pressure increases, which agrees with the trend revealed in previous studies (pressure exponent n is close to unity). No difference in combustion characteristics was found by comparing results of multi-port and single port fuel firing tests conducted in this and previous studies. A fuel regression model based on the Granular Diffusion Flame (GDF) model and Matthew and Frederick’s method is developed to investigate regression characteristics. Results calculated with this model agree with experimentally observed values, as well as the results calculated by Matthew and Frederick. However, this does hold true in tests with varying oxidizer port velocity. A GDF model only takes into account solid propellant regression, neglecting the effects of oxidizer velocity, and is shown in this study to be inappropriate for evaluating EBHR regression characteristics., 日本航空宇宙学会, 日本語 - 端面燃焼式ハイブリッドロケットにおける燃料後退特性の圧力依存性に関する実験的および解析的考察
齋藤勇士, 横井俊希, 嶋田泰三, 安河内裕之, 添田建太郎, 戸谷剛, 脇田督司, 永田晴紀, 日本航空宇宙学会年会講演会講演集(CD-ROM), 47th, ROMBUNNO.2D8, 2016年
日本語 - 円筒デトネーション波の伝播限界へ流路幅と円筒波面の曲率が与える影響
桧物恒太郎, 亀山頌太, 大関敦, 榎並聖也, 脇田督司, 戸谷剛, 永田晴紀, 流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム講演集(CD-ROM), 48th-34th, ROMBUNNO.2E13, 2016年
日本語 - 端面燃焼式ハイブリッドロケットの推力制御特性
永田晴紀, 齋藤勇士, 横井俊希, 嶋田泰三, 安河内裕之, 添田建太郎, 戸谷剛, 脇田督司, 日本機械学会年次大会講演論文集(CD-ROM), 2016, 0, S1920202, 2016年The authors have previously proposed the concept of end burning type hybrid rockets which would use cylindrical fuel grains consisting of an array of many small ports running in the axial direction through which oxidizer gas would flow. Because of difficulty in manufacturing a fuel grain that satisfied requirements such as high volumetric filling rate (above 0.97) and micro-sized port intervals, the end burning hybrid rocket had yet to be achieved. The authors succeeded to verify the novel end burning type hybrid rocket for the first time owing recent progress in 3D printing technology. This paper reports throttling firing tests to verify the excellent throttling characteristic of end-burning type hybrid rockets, that is, virtually no O/F shift during throttling.
, 一般社団法人 日本機械学会, 日本語 - ハイブリッドロケットにおけるノズルスロートエロージョン抑制材料に関する研究
山口亮, 川端良輔, 川端良輔, 平井翔大, KAMPS Landon, 脇田督司, 戸谷剛, 永田晴紀, 日本機械学会年次大会講演論文集(CD-ROM), 2016, 0, S1920201, 2016年These days, CAMUI type hybrid rocket has been developed. The rocket is faced with the problem of erosion of graphite nozzle. Nozzle erosion, which is mainly caused by oxidation reaction, decreases specific impulse of rockets. The authors selected some materials which can prevent nozzle erosion and evaluated anti-erosion materials for hybrid rocket nozzle by combustion experiments. We used 2.5 kN class CAMUI type hybrid rocket motor for combustion experiment. The authors used two types of fiber reinforced ceramics, silicon carbide ceramics, zirconia, tungsten carbide, silicon nitride ceramics, graphite coated by silicon carbide, alumina and carbon fiber carbon composite for nozzles. As an experimental result, tungsten carbide and graphite coated by silicon carbide showed good anti-erosion characteristics for hybrid rocket nozzles. High melting point and high thermal conductivity seem favorable for anti-erosion materials for hybrid rocket nozzles.
, 一般社団法人 日本機械学会, 日本語 - Method for determining nozzle throat erosion history in hybrid rockets
Landon Kamps, Yuji Saito, Ryosuke Kawabata, Yusuke Takahashi, Harunori Nagata, 52nd AIAA/SAE/ASEE Joint Propulsion Conference, 2016, 2016年01月01日
© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. The authors introduce a new reconstruction technique titled Nozzle Throat Reconstruction Technique or NTRT to estimate nozzle throat erosion history and oxidizerto-fuel mass ratio history in hybrid rockets. Nine five-second static firing tests were carried out on a 2kN-class Cascaded Multistage Impinging-jet type hybrid rocket motor under varying oxidizer flowrates to evaluate the accuracy of NTRT results. Nozzle throat erosion histories calculated by NTRT agreed well with measured values for initial and final nozzle throat radius, and successfully reconstructed the case where no measureable amount of nozzle throat erosion occurred. For equivalence ratios 0.6-1.4, the relationship between nozzle throat erosion rate and equivalence ratio determined by NTRT displays a trend consistent with chemical kinetic limited heterogeneous combustion theory, as well as predictions made by previous researchers. A numerical simulation was carried out to investigate the applicability of an alternative computational technique which uses a nozzle exit pressure measurement as input data, and revealed that pressure in vicinity of the static pressure port is likely to deviate from centerline pressure by up to 0.1 [MPa] or 30%. - Verification of the throttling characteristics of axial-injection end-burning type hybrid rockets
Yuji Saito, Toshiki Yokoi, Hiroyuki Yasukochi, Kentaro Soeda, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata, Proceedings of the International Astronautical Congress, IAC, 2016, S1920202, 2016年01月01日
Copyright © 2016 by the International Astronautical Federation (IAF). All rights reserved. The axial-injection end-burning type hybrid rocket originally proposed twenty years ago by Nagata and Hashimoto et al. recently recaptured the attention of researchers for its virtues such as constant ξ (oxidizer to fuel mass ratio) during firing and throttling operations. Previous studies revealed that, for combustion in a single-port fuel grain, the end-face regression rate in the axial direction is proportional to pressure, with a pressure exponent of 0.95. Accordingly, these rockets were expected to display good throttling characteristics. Given that no ξ shift occurs, keeping the oxidizer mass flow rate within 1% of its initial design point ensures specific impulse will remain within 97% of its design point. There are several requirements for realizing this type of hybrid rocket: 1) high fuel filling rate for obtaining an optimal ξ 2) small port intervals for increasing port merging rate; 3) ports arrayed across the entire fuel section. Because common manufacturing methods were unable to produce a fuel that satisfied these requirements, no previous researchers have conducted experiments with this kind of hybrid rocket. Recent advances in high-accuracy 3D printing have enabled such fuels to be produced for the first time. The fuel grains used in this study were produced by a high-precision light polymerized 3D printer. Each grain consisted of an array of 0.3 mm diameter ports for a fuel filling rate of 98%. Last year, the authors reported the results of multiple firing tests of an axial-injection end-burning type hybrid rocket using 3D printed fuel grains and verified that solid fuel regression rate is linearly dependent on pressure. In this study, the authors conducted a unique set of experiments to verify the throttling characteristics of the axial-injection end-burning type hybrid rocket. Oxidizer mass flow rate and chamber pressure were throttled during firings by actuating valves in a fluid circuit consisting of four oxidizer supply lines. Chamber pressure and oxidizer mass flow rate were measured during each firing. These experimental data were analyzed by a reconstruction technique to obtain ξ history. The results show that ξ remains almost constant during firing, even during throttling operations. Therefore, this study verifies that the axial-injection end-burning type hybrid rocket has superb throttling characteristics. Additionally, the study supports findings in previous research that indicate the pressure exponent is close to unity., 一般社団法人 日本機械学会, 日本語 - ハイブリッドロケットのノズル侵食履歴およびノズルスロート温度履歴の取得
永田晴紀, 川端良輔, 遠藤瞳, 金井竜一郎, 平井翔大, 脇田督司, 戸谷剛, 五十地輝, 燃焼シンポジウム講演論文集, 53rd, 244, 245, 2015年11月04日
日本語 - 気球を利用したスペースプレーン技術実証機の研究
丸 祐介, 澤井 秀次郎, 永田 晴紀, 坂東 信尚, 坂井 真一郎, 吉光 徹雄, 江口 光, Maru Yusuke, Sawai Shujiro, Nagata Harunori, Bando Nobutaka, Sakai Shin-ichiro, Yoshimitsu Tetsuo, 大気球シンポジウム: 平成27年度 = Balloon Symposium: 2015, 2015年11月
大気球シンポジウム 平成27年度(2015年11月5-6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: SA6000044003レポート番号: isas15-sbs-003, 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS), 日本語 - 反射を利用した爆轟波生成装置の連続運転条件における円筒爆轟波の生成
亀山頌太, 菊地敬太, 桧物恒太郎, 脇田督司, 戸谷剛, 永田晴紀, 宇宙航空研究開発機構特別資料 JAXA-SP-, JAXA-SP-14-010, 14-010, 173, 178, 2015年03月25日
第46回流体力学講演会/第32回航空宇宙数値シミュレーション技術シンポジウム (2014年7月3日-4日. 弘前文化センター), 弘前市, 青森県
46th Fluid Dynamics Conference / 32nd Aerospace Numerical Simulation Symposium (July 3-4, 2014. Hirosaki Bunka Center), Hirosaki, Aomori, Japan
To achieve practical uses of large bore Pulse Detonation Engines, the authors have proposed the enhanced reflection type detonation initiator. Here, successful transmission from a planar detonation wave in a predetonator to a self-sustainable expanding cylindrical detonation wave at a cylindrical path is a key issue. In a previous single-cycle research, the authors revealed that the necessary overfilling radius of the driver gas (stoichiometric H2-O2 mixture) to initiate a cylindrical wave is 75 mm. The aim of this study is to investigate the effect of multi-cycle operation condition on the transmission to a cylindrical detonation wave. The result shows that a planar detonation wave with theoretical CJ speed is formed but a cylindrical detonation wave is not formed at any overfilling radius from 50 to 175 mm. Soot track records at the cylindrical path surfaces and a mixing analysis of driver and target mixtures using ANSYS FLUENT strongly indicate that residual nitrogen gases at the exit of the predetonator have a negative effect on the transmission.
形態: カラー図版あり
Physical characteristics: Original contains color illustrations
資料番号: AA1530023031
レポート番号: JAXA-SP-14-010, 宇宙航空研究開発機構(JAXA), 日本語 - 金属膜を持つ表面微細構造の放射率
戸谷剛, 戸谷剛, 櫻井篤, 近藤良夫, 脇田督司, 永田晴紀, 日本機械学会熱工学コンファレンス講演論文集(CD-ROM), 2015, ROMBUNNO.C121, _C121-2_, 2015年
Technologies for drying a material with a flammable solvent at a low temperature are required. It is reasonable to evaporate a flammable solvent using radiation at the absorption band of a flammable solvent and to eject the flammable vapor with a cold air. The meta material with Au/Al2O3/Au was fabricated by using an sputtering equipment, an atomic layer deposition equipment, and an electron beam lithography equipment. The reflectances of the meta material structure were measured by using an FT-IR equipment with an integral sphere. It is clarified that the emissivities of the meta material structure are larger than 0.8 at the absorption band of toluene., 一般社団法人 日本機械学会, 日本語 - 超小型衛星に対する新しい熱設計手順の妥当性検証
毛利正宏, 須田俊太郎, 尾形明仁, 戸谷剛, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 59th, ROMBUNNO.3M01, 2015年
日本語 - 表面微細構造による放射波長制御に及ぼす形状の効果
戸谷剛, 戸谷剛, 色川俊雄, 脇田督司, 永田晴紀, 日本伝熱シンポジウム講演論文集(CD-ROM), 52nd, ROMBUNNO.G122, 2015年
日本語 - 反射を利用した大口径PDEイニシエータにおける円筒デトネーション波から円環デトネーション波への遷移に関する研究
大関敦, 菊地敬太, 桧物恒太郎, 亀山頌太, 脇田督司, 戸谷剛, 永田晴紀, 流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム講演集(CD-ROM), 47th-33rd, ROMBUNNO.2E12, 2015年
日本語 - ハイブリッドロケットにおけるノズルスロートエロージョンに関する研究
川端良輔, 齋藤勇士, 平井翔大, 脇田督司, 戸谷剛, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 59th, ROMBUNNO.1A06, 2015年
日本語 - 展開型太陽電池パネルを有する超小型衛星の新しい熱設計手順
尾形明仁, 戸谷剛, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 59th, ROMBUNNO.3M02, 2015年
日本語 - 結晶構造変化により蓄熱する蓄熱材の地上および軌道上試験
戸谷剛, 國拓也, 磯野拓也, 佐藤敏文, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 59th, ROMBUNNO.3M06, 2015年
日本語 - Verification firings of end-burning type hybrid rockets
Harunori Nagata, Hayato Teraki, Yuji Saito, Ryuichiro Kanai, Hiroyuki Yasukochi, Masashi Wakita, Tsuyoshi Totani, 51st AIAA/SAE/ASEE Joint Propulsion Conference, 2015年01月01日
© 2015 American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. The authors have proposed end burning type hybrid rockets. A key point of this idea is that a motor uses cylindrical fuel having many ports in axial direction, in which oxidizer gas flows. Because of difficult requirements to make an end burning type solid fuel grain, i.e., high volumetric filling rate above 0.95 and small port intervals, the end burning hybrid rocket was yet to be achieved. This paper reports results of verification firing tests of an end burning type hybrid rocket being realized for the first time. Recent progress of 3D printers made the production of the fuel grain possible. The results clearly show the initial transient and the steady periods of the end burning mode. It also proves a virtue of no oxidizer to fuel ratio shift during firing. Since the initial transient is a period for the exit end face to attain a steady state shape, an initial end face shape being close to the steady state shape can shorten this period. A firing test with a fuel having tapered ports showed that it attains a steady state in less than 1 second, which is much shorter than a non-tapered case of about 6 seconds. - 結晶転移により蓄熱する蓄熱材の蓄熱量に及ぼす劣化の影響
國拓也, 戸谷剛, 佐藤敏文, 脇田督司, 永田晴紀, Thermophys Prop, 35th, 226, 228, 2014年11月22日
日本語 - 端面燃焼式ハイブリッドロケット用固体燃料の後退特性
寺木勇人, 齋藤勇士, 金井竜一郎, 脇田督司, 戸谷剛, 永田晴紀, 燃焼シンポジウム講演論文集, 52nd, 556, 557, 2014年11月20日
日本語 - 低レイノルズ数域におけるCAMUI型固体燃料の燃料後退特性
遠藤瞳, 川端良輔, 脇田督司, 戸谷剛, 永田晴紀, 燃焼シンポジウム講演論文集, 52nd, 554, 555, 2014年11月20日
日本語 - 反射体を用いたPDEイニシエータの円錐部の短縮が爆轟波の伝播に与える影響
菊地敬太, 桧物恒太郎, 亀山頌太, 脇田督司, 戸谷剛, 永田晴紀, 燃焼シンポジウム講演論文集, 52nd, 332, 333, 2014年11月20日
日本語 - 一定レイノルズ数酸化剤衝突噴流のよどみ点近傍における固体燃料燃焼速度に雰囲気圧力が及ぼす影響
斉藤竜也, 寺川健, 脇田督司, 戸谷剛, 永田晴紀, 燃焼シンポジウム講演論文集, 52nd, 538, 539, 2014年11月20日
日本語 - 大気球を利用したスペースプレーン技術の高速飛行実験システムの開発研究
丸 祐介, 澤井 秀次郎, 小林 弘明, 坂東 信尚, 坂井 真一郎, 後藤 健, 永田 晴紀, スペースプレーン技術実証機WG, Maru Yusuke, Sawai Shujiro, Kobayashi Hiroaki, Bando Nobutaka, Sakai Shin-ichiro, Goto Ken, Nagata Harunori, 大気球シンポジウム: 平成26年度 = Balloon Symposium: 2014, 2014年11月
大気球シンポジウム 平成26年度(2014年11月6-7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: SA6000021011レポート番号: isas14-sbs-011, 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS), 日本語 - Radiation Enhancement by Metal Film on Micro Cavities in Resin
Tsuyoshi TOTANI, Toshio IROKAWA, Minoru IWATA, Masashi WAKITA, Harunori NAGATA, Proceedings of the 15th International Heat Transfer Conference, IHTC-15, IHTC15-8771, 2014年08月, [査読有り]
英語 - スペースプレーン技術の飛行実証のための気球を利用した飛行実験システムの開発
丸祐介, 澤井秀次郎, 小林弘明, 坂東信尚, 坂井真一郎, 後藤健, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 58th, 2014年 - 結晶構造変化により蓄熱する蓄熱材の軌道上試験
戸谷剛, 國拓也, 佐藤敏文, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 58th, ROMBUNNO.2E12, 2014年
日本語 - 酸化剤衝突噴流による固体燃料燃焼場の数値解析
縄田和也, 大島伸行, 斉藤達也, 脇田督司, 戸谷剛, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 58th, ROMBUNNO.2J08, 2014年
日本航空宇宙学会, 日本語 - 単一液滴流からの排熱量の測定と数値解析
高梨知広, 戸谷剛, 木村優斗, 永田晴紀, 脇田督司, 日本伝熱シンポジウム講演論文集(CD-ROM), 51st, ROMBUNNO.FSP406, 2014年
日本語 - ハイブリッドロケット燃焼データ解析法の精度と適用範囲に関する研究
齋藤勇士, LAMBOURG Aurelien, 川端良輔, 脇田督司, 戸谷剛, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 58th, ROMBUNNO.2J02, 2014年
日本語 - 円筒波の起爆と希釈燃料への伝播に流路幅が与える影響
桧物恒太郎, 菊地敬太, 亀山頌太, 脇田督司, 戸谷剛, 永田晴紀, 衝撃波シンポジウム講演論文集(CD-ROM), 2013, ROMBUNNO.3A2-5, 2014年
日本語 - 反射を利用した爆轟波生成装置の連続運転条件における円筒爆轟波の生成
亀山頌太, 菊地敬太, 桧物恒太郎, 脇田督司, 戸谷剛, 永田晴紀, 流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム講演集(CD-ROM), 46th-32nd, ROMBUNNO.2E03, 2014年
日本語 - 上空打ち上げ用ハイブリッドロケットのための断熱液体酸素タンクの検討
川端良輔, 稲場康彦, 石山達也, 脇田督司, 戸谷剛, 永田晴紀, 日本機械学会年次大会講演論文集(CD-ROM), 2014, 0, _S1920103, -_S1920103-, 2014年
A hypersonic flight experiment by a combination of a scientific observation balloon and a CAMUI-type hybrid rocket is proposed. This is one of the rockoon methods. In this experiment, liquid oxygen (LOX) is filled on the ground and thermally insulated until the launch in high altitudes. The gas phase volume in the LOX tank decrease with time due to heat input from ambient. Eventually, the gas phase disappears, followed by a sudden increase of the LOX tank pressure. The heat insulation of the LOX tank has to provide no less than five hours from the LOX filing to the sudden pressure increase. In this study, a numerical model to predict pressure histories in the LOX tank was developed. The results were compared and verified with experimental results in a small scale. Finally, pressure histories at a full-scale system were predicted to provide a guideline for the design of a thermal insulator., 一般社団法人 日本機械学会, 日本語 - 金属膜を持つ微小キャビティの構造と放射ピークの関係
戸谷剛, 戸谷剛, 色川俊雄, 脇田督司, 永田晴紀, 日本機械学会熱工学コンファレンス講演論文集(CD-ROM), 2014, ROMBUNNO.E215, _E215-2_, 2014年
Micro-cavities were made from Zr dispersing liquid on SiO_2 bases and were coated by Au for generating resonance phenomena. The theoretical wavelength of reflectance drop differed 4% from the experiment data. A numerical analysis taking into account the pitch was carried out to find causes of the difference. The wavelength of reflectance drop in the analysis were closer to experiment data than theoretical predictions. On the other hand, the wavelengths of reflectance drop except for the fundamental vibrational mode obtained in the analysis did not accord 7.6 to 8.5 % with experimental data. The reflectance obtained in the analysis was 4 to 11% different from experimental data., 一般社団法人 日本機械学会, 日本語 - 結晶変化を伴う蓄熱材の蓄熱・放熱試験
戸谷剛, 國拓也, 佐藤敏文, 脇田督司, 永田晴紀, 日本機械学会熱工学コンファレンス講演論文集(CD-ROM), 2014, ROMBUNNO.F117, _F117-2_, 2014年
It is desirable that the heat storage materials for nano and micro satellites have the characteristic of not phase change but crystal transformation for the heat storage. Trans-1,4-polybutadien has the desirable characteristic of crystal transformation at the heat storage temperature. The heat storage and release tests were conducted on the ground and in orbit. It is clarified that the heat storage and release characteristics do not change in heat storage and release cycle of 300 times and Trans-1,4-polybutadien can storage and release heat in orbit., 一般社団法人 日本機械学会, 日本語 - A numerical analysis and measurement of radiation heat from a liquid droplet stream under gravitational environment
Tomohiro Takanashi, Tsuyoshi Totani, Yuto Kimura, Harunori Nagata, Masashi Wakita, Proceedings of the International Astronautical Congress, IAC, 1, 519, 522, 2014年01月01日
Copyright © 2014 by the International Astronautical Federation. All rights reserved. A high efficiency system for disposing of large quantities of waste heat is needed to realize large structures in space such as solar power satellites, space factories in orbit and bases on the moon. The Liquid Droplet Radiator (LDR) can be the system instead of a conventional radiator with solid plates. In previous studies, the performance tests of a droplet generator, a linear droplet collector, and a gear pump have been conducted under micro-gravity environment. The radiation heat from liquid droplet streams is due to be measured by a part of research of a LDR. It is needed to conduct an experiment to measure the radiation heat from liquid droplet streams under micro-gravity environment to know properties of a LDR. As a first step for the experiment under micro-gravity environment, an experiment for measuring the radiation heat from a liquid droplet stream was conducted under gravitational environment. In order to take the influence of absorption or dispersion into consideration, a numerical analysis of inside an experimental device was also performed. The appropriateness of a result of the experiment ware evaluated by comparing a experimental value with results of numerical analysis. Working fluids are squirt as liquid droplets into a vacuum inside a shroud that is cooled by liquid nitrogen under 79 K. The radiation heat from a liquid droplet stream is measured by using a radiation sensor (Captec RF-100) pasted on inside wall of the shroud. The shroud is 700 mm high and 100 mm in diameter. Silicon oil (Shin-Etsu Chemical Co., Ltd. KF-96 10 cSt) is used as the working fluid. The experimental value are between 0.617 and 0.739 W/m∧2 when the center-to-center distance and diameter of liquid droplets are changed. The more the center-to-center distance and diameter is decreased, the more experimental results are increase. A trend of the results of numerical analysis fitted in the trend of experimental values. A biggest difference is less than 7 between experimental values and the results of numerical analysis when the radiation factor is 0.70 in the numerical analysis. It can be said that the radiation heat from a liquid droplet stream is measured correctly. Additionally the radiation factor of a liquid droplet can be estimated by using a method of this research. - 円盤状流路に入射する平面デトネーション波の半径方向に伝播する円筒波への遷移にプリデトネータ直径が与える影響
桧物恒太郎, 菊地敬太, 脇田督司, 戸谷剛, 永田晴紀, 燃焼シンポジウム講演論文集, 51st, 544, 545, 2013年11月20日
日本語 - Scale effect on flame spread rate in narrow cylindrical gap
Tsuneyoshi Matsuoka, Yuji Nakamura, Harunori Nagata, Takuya Yamazaki, The 8th International Symposium on Advanced Science and Technology in Experimental Mechanics (8th ISEM’13-Sendai), 2013年11月, [査読有り]
英語 - 1212 CAMUI-Avionicsの打上実験結果に対する故障解析
齊藤 慎之介, 伊藤 那知, 飯野 貴之, 長谷川 貴之, 三橋 龍一, 佐島 新, 永田 晴紀, 千葉 一永, 設計工学・システム部門講演会講演論文集, 2013, 23, "1212, 1"-"1212-5", 2013年10月23日
A students group Hokkaido Space Union had been to space development for nine years. HSU made data handling unit named Sprit-nexus. It has been mounted on the CAMUI hybrid rocket. The objective obtain a telemetry data of the Sprit-nexus for the CAMUI hybrid rocket recovery. Telemetry data is composed of position coordinates and altitude and acceleration and attitude. CAMUI hybrid rocket is being developed by professor Harunori Nagata in Hokkaido University. CAMUI hybrid rocket was launched in Taiki Aerospace Research Field on July 28 2012. It was to go recovery by the ships from the sea after launch but, Sprit-nexus could not obtained a telemetry data. I will tried to find the cause of failure. As a result, the visual observation proved simple faulty electrical wiring., 一般社団法人日本機械学会, 日本語 - 表面微細構造による放射波長制御
戸谷剛, 色川俊雄, 脇田督司, 永田晴紀, 日本機械学会熱工学コンファレンス講演論文集, 2013, 153‐154, 154, 2013年10月18日
A periodic micro structure of the cubic cavity of 6.0μm×6.0μm×6.0μm is built on an ultra violet curable resin by the UV nanoimprint method. A gold film is sputtered on the periodic micro structure. This method can manufacture the periodic micro structure with the metal surface on a large area at a lower cost than etching. It is measured that the maximum normal emissivity of the periodic micro structure is 0.6 near 10μm in the case of the gold film of 200 nm thick on the surface facing a sputter source. The periodic micro structure with the gold film contributes to the prevention of global warming., 一般社団法人日本機械学会, 日本語 - 液滴流の熱放射測定のための数値解析
高梨知広, 戸谷剛, 永田晴紀, 脇田督司, 日本機械学会熱工学コンファレンス講演論文集, 2013, 163‐164, 164, 2013年10月18日
Waste heat from liquid droplet flow is due to be measured by a part of research of the liquid droplet radiator which is a radiator for large-sized spacecrafts. In order to take the influence of absorption or dispersion into consideration, the simulation inside an experimental device was performed. In the case of the single liquid droplet flow, the energy from the liquid droplet flow in a measuring range was set to 0.252 w/m^2. The energy from the surface of a wall was set to 0.116 w/m2. A possibility of separating the amount of thermal radiation obtained in an experiment was able to be shown., 一般社団法人日本機械学会, 日本語 - 結晶転移する蓄熱材の蓄熱量
戸谷剛, 國拓也, 佐藤敏文, 脇田督司, 永田晴紀, 日本機械学会熱工学コンファレンス講演論文集, 2013, 115‐116, 116, 2013年10月18日
In order to increase the heat capacity of nano and micro satellites, the development of a heat storage material for nano and micro satellites has been performed. It is desirable that the heat storage materials for nano and micro satellites have the characteristic of not phase change but crystal transformation for the heat storage. Trans-1,4-polybutadien transforms crystal structure at the temperature of heat storage. Trans-1,4-polybutadien is produced and the heat storage performance is measured. It is clarified that the produced trans-1,4-polybutadien has 80 J/g of heat storage performance. This value corresponding to 70 % of heat storage performance pointed out in the previous literature., 一般社団法人日本機械学会, 日本語 - フライトテストベッドとしてのCAMUI型ハイブリッドロケットの進展
永田 晴紀, 植松 努, 伊藤 献一, 宇宙科学技術連合講演会講演集, 57, 7p, 2013年10月09日
日本航空宇宙学会, 日本語 - 缶サット・超小型衛星を用いた創造的科学技術人育成ネットワークの構築
宮崎 康行, 永田 晴紀, 木村 真一, 宇宙科学技術連合講演会講演集, 57, 5p, 2013年10月09日
日本航空宇宙学会, 日本語 - Development and flight demonstration of 5 kN thrust class CAMUI type hybrid rocket
Harunori Nagata, Masashi Wakita, Tsuyoshi Totani, Tsutomu Uematsu, 49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 12, 29, Ta_1, Ta_4, 2013年09月16日
The authors have been developing CAMUI (Cascaded Multistage Impinging-jet) type hybrid rockets, explosive-flee small rocket motors. This is to downsize the scale of suborbital flight experiments on space related technology development. A key idea is a new fuel grain design to increase gasification rates of a solid fuel. By the new fuel grain design, the combustion gas repeatedly impinges on fuel surfaces to accelerate the heat transfer to the fuel. To demonstrate flight performance of a newly developed 5000 N thrust class motor and accumulate flight data around the sonic speed, a launch test was conducted from a coast to the sea. Basic technologies for the sea recovery are staged braking by parachutes, suspending the fuselage on the ocean, and locating the fuselage by an electric beacon and sea marker. Test results were successful and all of the fuselage was recovered. A typical drag coefficient profile around the sonic speed was obtained., THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 英語 - Proposal of procedure of thermal design on micro and nano satellites pointing to earth
Tsuyoshi Totani, Hiroto Ogawa, Ryota Inoue, Tilok K. Das, Masashi Wakita, Harunori Nagata, 43rd International Conference on Environmental Systems, 2013年09月16日
A procedure for the thermal design of micro and nano satellites is proposed in order to complete the thermal design of micro and nano satellites for about 1 year. Two concepts of thermal design are considered to keep the temperature change of components within the design temperature range of components. One concept is to decrease the temperature change using the whole heat storage of the micro and nano satellite. The other is to decrease the temperature change of the inner structure where the components with the narrow design temperature range are mounted. The temperature of micro and nano satellites designed in the former concept is calculated using one nodal analysis method. The temperature of micro and nano satellites designed in the latter concept is calculated using two nodal analysis method. The combinations of optical properties on structures and components to keep the temperature within the design temperature range of components are clarified using one or two nodal analysis. Then, the multi-nodal analyses are carried out to be designed in detail based on the optical properties clarified from the one-nodal analysis and two nodal analysis. This procedure of thermal design is applied to Hodoyoshi-1 satellite. Hodoyoshi-1 satellite is the micro satellite that is about 50 cm in width, 50 cm in depth, 50 cm in height, is about 50 kg in mass, has two inner plates, has solar cells on the body, flies on the sun-synchronous orbit of the 500 km of altitude and is pointing to the Earth. The thermal design of Hodoyoshi-1 satellite has been completed for about ten months. The validity of this procedure is confirmed and the problems of this procedure are clarified. - 可燃性固体内部の微小空隙を燃え拡がる火炎におよぼすスケールの影響
松岡常吉, 中村祐二, 永田晴紀, 山崎拓也, 日本実験力学会講演論文集, 13, 394, 399, 2013年08月20日
日本実験力学会, 日本語 - Gravity effect on flow field of flame spread in fuel tube
Tsuneyoshi Matsuoka, Harunori Nagata, Yuji Nakamura, Seventh International Symposium on Scale Modeling (ISSM-7), 2013年08月, [査読有り]
英語 - Detonation Transition Around Cylindrical Reflector of Pulse Detonation Engine Initiator
Masashi Wakita, Masayoshi Tamura, Akihiro Terasaka, Kazuya Sajiki, Tsuyoshi Totani, Harunori Nagata, JOURNAL OF PROPULSION AND POWER, 29, 4, 825, 831, 2013年07月, [査読有り]
To achieve reliable transmission of detonation waves to a pulse detonation engine combustor (detonation chamber), the authors propose a pulse detonation engine initiator that uses a cylindrical reflector downstream of a predetonator exit. The detonation wave propagates around the reflector to change the wave shape in three transition stages: from a planar detonation wave in the predetonator to an expanding cylindrical detonation wave, from the cylindrical wave to a planar toroidal detonation wave, and from the toroidal wave to a planar detonation wave in the detonation chamber. The cylindrical wave propagates along a cylindrical path between the reflector and front wall of the detonation chamber, and the toroidal wave propagates along an annular path between the reflector and sidewall of the detonation chamber. The purpose of this study was to examine the influence of the gap width L of the annular path on the transition stages from cylindrical to toroidal and from toroidal to planar. A series of experiments that filled the entire test section with the driver gas mixture (stoichiometric hydrogen oxygen mixture) showed that the expanding cylindrical detonation wave was sufficiently strong to survive the rarefaction waves from the corners of the reflector at all of the investigated annular gap widths (5, 10, 15, and 20 mm) and was transmitted to the planar toroidal wave successfully in all cases. When the strength of the cylindrical detonation wave was under a supercritical condition for diffraction at the reflector corner, the necessary filling distance for the driver gas was predicted well by the Whitham theory. A second series of experiments showed the influence of the annular gap width on the detonation transition from the planar toroidal detonation wave to the planar detonation wave. Two different types of detonation transitions termed "continuous transition" and "temporal quenching" were observed. The threshold value of L/lambda for continuous transition is approximately four., AMER INST AERONAUTICS ASTRONAUTICS, 英語 - Effect of Reynolds number and flow channel geometry on regression formula for forward-end faces in CAMUI type fuel grain
Ryuichiro Kanai, Tatsuya Ishiyama, Masahiro Nohara, Hirokazu Izumo, Masashi Wakita, Tsuyoshi Totani, Harunori Nagata, Advances in the Astronautical Sciences, 146, 79, 84, 2013年04月24日
The authors have been developing fuel regression formulas for CAMUI type hybrid rocket motors. A fuel block in a CAMUI-type fuel grain is a short-axis cylinder having two axial ports. Previous experiments showed that an experimental constant in the regression formula for forward-end faces depends on port length L, mean port diameter D, and the Reynolds number of the flow. In this paper, the authors examined these effects more closely to clarify the basic mechanism of these dependencies. - 大口径パルスデトネーションエンジン用イニシエータにおける円筒デトネーション波の伝播に関する研究
桧物恒太郎, 棧敷和弥, 脇田督司, 戸谷剛, 永田晴紀, 宇宙航空研究開発機構特別資料 JAXA-SP-, 12-010, 73, 78, 2013年03月29日
日本語 - 大口径パルスデトネーションエンジン用イニシエータにおける円筒デトネーション波の伝播に関する研究—Propagation Characteristics of Cylindrical Detonation Wave in Detonation Initiator for Large Bore Pulse Detonation Engines
桧物, 恒太郎, 棧敷, 和弥, 脇田, 督司, 戸谷, 剛, 永田, 晴紀, Himono, Tsunetaro, Sajiki, Kazuya, Wakita, Masashi, Totani, Tsuyoshi, Nagata, Harunori, 宇宙航空研究開発機構特別資料 = JAXA Special Publication: Proceedings of 44th Fluid Dynamics Conference / Aerospace Numerical Simulation Symposium 2012, JAXA-SP-12-010, 73, 78, 2013年03月29日
第44回流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム2012 (2012年7月5日-6日. 富山国際会議場大手町フォーラム), 富山市, 富山県
44th Fluid Dynamics Conference / Aerospace Numerical Simulation Symposium 2012 (July 5-6, 2012. Toyama International Conference Center), Toyama Japan
Detonation initiation is one of the most important problems of pulse detonation engines. A reflector installed near a predetonator exit is effective to maintain detonation waves. The incident detonation waves from the predetonator transform its shape from planer detonation wave to cylindrical detonation wave and from cylindrical detonation wave to toroidal detonation wave around the reflector. To prevent the detonation wave downstream of a reflector from disappearing, it is important to form a cylindrical detonation waves upstream of the reflector. To define the minimal quantity of driver gas to form cylindrical detonation waves and to make sense of the mechanism that determine the quantity, the authors uses a large-bore plate combustor and varied the following parameters to find the quantity to form a cylindrical detonation wave : the quantity of driver gas and nitrogen concentration. To evaluate the quantity of the driver gas, the authors us""" es overfilling radius. The over filling radius is the radius of the cylinder that is equal to the driver gas filled in the combustion chamber. Results indicate that the mixing between driver gas and target gas is critical to successful transmission. To form the cylindrical detonation wave reliably, the concentration of driver gas must be maintained at a high level so that the cell size doesn’t become large.
形態: カラー図版あり
Physical characteristics: Original contains color illustrations
資料番号: AA0061958013
レポート番号: JAXA-SP-12-010, 宇宙航空研究開発機構(JAXA), 日本語 - 平成24年度先進的燃焼技術の調査研究:推進薬の燃焼技術
永田晴紀, 松岡常吉, 和田豊, 福地亜宝郎, 藤里公司, 2013年03月, [招待有り]
日本語, 書評論文,書評,文献紹介等 - 超小型弾道ロケット用液体酸素供給系の開発
永田 晴紀, 混相流, 27, 4, 393, 400, 2013年
A liquid oxygen supply system for a small-scale sounding rockets was developed. The sounding rokets are hybrid type employing a combination of plastics (PMMA or high density polyethylene) and liquid oxygen as propellants. Key points for the miniaturization were using no valve in the liquid oxygen feeding line and omitting precooling treatment of the line. Without precooling treatment, the liquid oxygen in the feeding line becomes multiphase flow until the temperature of the feeding line falls below the boiling temperature of the liquid oxygen. A characteristic time of multiphase flow duration was proposed to evaluate the duration a multiphase flow holds in the feeding line. Static firing tests showed that the multiphase flow ends within the half the characteristic time, showing that the liquid oxygen flow rate history without a precooling treatment is acceptable for an actual operation of the rocket motor. Finally, an impinging type injector was developed to remove a combustion instability, caused by a coupling of the combustion chamber pressure and the propellant (liquid oxygen) feed system., 日本混相流学会, 英語 - 大学でできる再使用型ロケット実験(その6)
米本, 浩一, 相良, 慎一, 松本, 剛明, 永田, 晴紀, 越智, 徳昌, 石本, 真二, 麥谷, 高志, 牧野, 隆, 木元, 健一, Yonemoto, Koichi, Sagara, Shinichi, Matsumoto, Takaaki, Nagata, Harunori, Ochi, Yoshimasa, Ishimoto, Shinji, Mugitani, Takashi, Makino, Takashi, Kimoto, Kenichi, 平成24年度宇宙輸送シンポジウム: 講演集録 = Proceedings of Space Transportation Symposium: FY2012, 2013年01月
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
Space Transportation FY2012 (January 17-18, 2013. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan
九州工業大学では、2005年より有翼ロケットによる無人のサブオービタル飛行システムの研究を進めてきた。航法誘導制御システム、推進システム、回収システムや複合材構造設計技術等の個別要素研究を進める一方、それらの技術実証を目的とした小型有翼ロケット実験機を用いた飛行実験を行ってきた。2010年からは、北海道大学、防衛大学大学校およびJAXA宇宙輸送ミッション本部や航空宇宙機メーカー等と連携し、高々度飛行を目指す有翼ロケット実験機の開発を進めている。これまでに行ってきた飛行実験結果を紹介し、サブオービタル飛行を目指す将来計画についても報告する。
形態: カラー図版あり
形態: PDF
Physical characteristics: Original contains color illustrations
Physical characteristics: PDF
Translation title: Experiment of reusable rocket performable in the university (part 5)
資料番号: AA0061856056
レポート番号: STCP-2012-056, 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS), 日本語 - 推力5000N級CAMUI型ハイブリッドロケットの推進系及び機体構造の開発
五十地, 輝, 前田, 祐義, 橋本, 祐治, 植松, 努, 永田, 晴紀, Isochi, Hikaru, Maeda, Hiroyoshi, Hashimoto, Yuji, Uematsu, Tsutomu, Nagata, Harunori, 平成24年度宇宙輸送シンポジウム: 講演集録 = Proceedings of Space Transportation Symposium: FY2012, 2013年01月
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
Space Transportation FY2012 (January 17-18, 2013. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan
Uematsu Electric Co., Ltd. has been developing CAMUI-type hybrid rockets in collaboration with Hokkaido University. In July 2012 a flight test with a 5000 N thrust class CAMUI-type rocket was conducted. Main objectives of this flight test were technology acquisition concerning supersonic flight, tracking/telemetry and command subsystem, and airframe recovery from sea-water. This paper describes the development outline of the propulsion system and airframe structure of the test vehicle.
形態: カラー図版あり
形態: PDF
Translation title: Development of the propulsion system and airframe structure of 5000 N thrust class CAMUI-type hybrid rocket
Physical characteristics: Original contains color illustrations
Physical characteristics: PDF
資料番号: AA0061856077
レポート番号: STCP-2012-077, 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS), 日本語 - 気球とロケットを組み合わせた揚力飛翔体の極超音速飛行実験の検討
丸 祐介, 澤井 秀次郎, 坂井 真一郎, 坂東 信尚, 小林 弘明, 永田 晴紀, スペースプレーン技術実証機ワーキング・グループ, Maru Yusuke, Sawai Shujiro, Sakai Shinichiro, Bando Nobutaka, Kobayashi Hiroaki, Nagata Harunori, 平成24年度宇宙輸送シンポジウム: 講演集録 = Proceedings of Space Transportation Symposium: FY2012, 2013年01月
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: AA0061856060レポート番号: STCP-2012-060, 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS), 日本語 - S192021 CAMUI型ハイブリッドロケットの超音速フライトデータを用いた空気抵抗係数の算出([S192-02]宇宙システムに関する実践的解決と知見の汎用化(2))
五十地 輝, 前田 祐義, 橋本 祐治, 清尾 陽平, 植松 努, 永田 晴紀, 年次大会, 2013, 0, _S192021, 1-_S192021-4, 2013年
The authors have been developing CAMUI (Cascaded Multistage Impinging-jet) type hybrid rockets, explosive-flee small rocket motors. This is to downsize the scale of suborbital flight experiments on space related technology development. To demonstrate flight performance of a newly developed 5000 N thrust class motor and accumulate flight data around the sonic speed, a launch test was conducted from a coast to the sea. Test results were successful and all of the fuselage was recovered. To obtain drag coefficient, we used flight data, histories of thrust and propellant flow rates obtained by the static firing test. A typical drag coefficient profile around the sonic speed was obtained., 一般社団法人 日本機械学会, 日本語 - 小型有翼ロケットの飛行実験と今後の計画
松本剛明, 米本浩一, 相良慎一, 永田晴紀, 越智徳昌, 石本真二, 麥谷高志, 日本機械学会年次大会講演論文集(CD-ROM), 2013, 0, _S192022, 1-_S192022-5, 2013年
Since 2005, Kyushu Institute of Technology has been conducting researches on a new unmanned sub-orbital system based on the research and development achievements of reusable sounding rocket called HIMES (Highly Maneuverable Experimental Space vehicle), the concept of which was first proposed by the Institute of Space and Astronautical Sciences of former Ministry of Education in the 1980s, but failed its commercialization. Flight experiments have been performed using small test vehicles of winged rocket in order to demonstrate guidance and control system performance and terminal recovery technologies. A larger winged rocket test vehicle that aims at higher altitude flight demonstration and a sub-orbital prototype vehicle are under design by industry-government-academia collaboration since 2010. This paper introduces the future research and development plan., 一般社団法人 日本機械学会, 日本語 - EFFECT OF REYNOLDS NUMBER AND FLOW CHANNEL GEOMETRY ON REGRESSION FORMULA FOR FORWARD-END FACES IN CAMUI TYPE FUEL GRAIN
Ryuichiro Kanai, Tatsuya Ishiyama, Masahiro Nohara, Hirokazu Lzumo, Masashi Wakita, Tsuyoshi Totani, Harunori Nagata, SPACE FOR OUR FUTURE, 146, 79, 84, 2013年, [査読有り]
The authors have been developing fuel regression formulas for CAMUI type hybrid rocket motors. A fuel block in a CAMUI-type fuel grain is a short-axis cylinder having two axial ports. Previous experiments showed that an experimental constant in the regression formula for forward-end faces depends on port length L, mean port diameter D, and the Reynolds number of the flow. In this paper, the authors examined these effects more closely to clarify the basic mechanism of these dependencies., UNIVELT INC, 英語 - 2次元流れ場のよどみ点近傍におけるPMMA燃料の後退特性
斉藤竜也, 松岡常吉, 寺川健, 脇田督司, 戸谷剛, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 57th, ROMBUNNO.2A01, 2013年
日本語 - 環状爆轟波から平面爆轟波への遷移に流路形状が及ぼす影響
菊地敬太, 桧物恒太郎, 脇田督司, 戸谷剛, 永田晴紀, 流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム講演集(CD-ROM), 45th-2013, ROMBUNNO.1C05, 2013年
日本語 - CAMUI型ハイブリッドロケットの作動履歴におけるスケール則の構築
稲場康彦, 石山達也, 金井竜一朗, 脇田督司, 戸谷剛, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 57th, ROMBUNNO.2A03, 2013年
日本語 - 超小型衛星の熱設計手順の提案
戸谷剛, DAS Tilok Kumar, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 57th, ROMBUNNO.3H15, 2013年
日本語 - 高レイノルズ数域におけるCAMUI型固体燃料の後退特性におよぼすスケールの影響
石山達也, 稲場康彦, 寺川健, 遠藤瞳, 永田晴紀, 戸谷剛, 脇田督司, 宇宙科学技術連合講演会講演集(CD-ROM), 57th, ROMBUNNO.2A02, 2013年
日本語 - S192023 短期開発を実現する超小型衛星の熱設計法([S192-02]宇宙システムに関する実践的解決と知見の汎用化(2))
戸谷 剛, DAS Kumar Tilok, 脇田 督司, 永田 晴紀, 年次大会, 2013, 0, _S192023, 1-_S192023-5, 2013年
A guide for the thermal design of micro and nano satellites is proposed in order to complete the thermal design of micro and nano satellites for about 1 year. Two concepts of thermal design are considered to keep the temperature change of components within the design temperature range of components. One concept is to decrease the temperature change using the whole heat storage of the micro and nano satellite. The other is to decrease the temperature change of the inner structure where the components with the narrow design temperature range are mounted. The temperature of micro and nano satellites designed in the former concept is calculated using one-nodal analysis method. The temperature of micro and nano satellites designed in the latter concept is calculated using twonodal analysis method. The combinations of optical properties on structures and components to keep the temperature within the design temperature range of components are clarified using one- or two-nodal analysis. Then, the multinodal analyses are carried out to be designed in detail based on the optical properties clarified from the one-nodal analysis and two-nodal analysis. This guide of thermal design is applied to Hodoyoshi-1 satellite. Hodoyoshi-1 satellite is the micro satellite that is about 50 cm in width, 50 cm in depth, 50 cm in height, is about 50 kg in mass, has two inner plates, has solar cells on the body, flies on the sun-synchronous orbit of the 500 km of altitude and is pointing to the Earth. The thermal design of Hodoyoshi-1 satellite has been completed for about ten months. The validity of this procedure is confirmed and the problems of this procedure are clarified., 一般社団法人 日本機械学会, 日本語 - EMISSIVITY OF WAVELENGTH-SELECTIVE RADIATOR WITH PERIODIC MICROCAVITIES
Tsuyoshi Totani, Minoru Iwata, Masashi Wakita, Harunori Nagata, PROCEEDINGS OF THE 11TH INTERNATIONAL CONFERENCE ON NANOCHANNELS, MICROCHANNELS, AND MINICHANNELS, 2013, 2013年, [査読有り]
A periodic microstructure of the cubic cavity 6.0 mu m wide, 6.0 mu m deep, and 6.0 mu m high is built on an ultraviolet curable resin via UV nanoimprinting. The 200 nm thick gold film is sputtered on the periodic micro structure. The hemispherical spectroscopic transmittance and reflectance of the periodic microstructure with the gold film are measured using a Fourier transform infrared spectrometer with an integrating sphere. The hemispherical spectroscopic transmittance is 0.0 from 2 to 15 mu m wavelength. The hemispherical spectroscopic reflectance is 1.0 from 2 to 8 mu m wavelength and from 12 to 15 mu m wavelength. The bottom of the hemispherical spectroscopic reflectance is 0.4 near 10 mu m. Assuming Kirchhoff's law, the maximum normal emissivity of the periodic microstructure is 0.6 near 10 mu m. It is clarified that the periodic microcavities with a gold film built via UV nanoimprinting and sputtering can enhance maximum spectral emissive power of radiation., AMER SOC MECHANICAL ENGINEERS, 英語 - Effect of reynolds number and flow channel geometry on fuel regression characteristics in camui type hybrid rocket
H. Nagata, M. Nohara, R. Kanai, M. Wakita, T. Totani, Proceedings of the International Astronautical Congress, IAC, 10, 8049, 8055, 2012年12月01日
Regression formulas for solid fuels in CAMUI type hybrid rockcts were developed. An empirical values of the mass flow density exponent m for upstream end faces obtained by two motors with similarity shape and different scaling did not agree with each other. The different Reynolds number (Re) range caused this disagreement. The exponent m coincides with the exponent of Re in a function giving Nusselt number. Static firing tests with various Re and LID (the ratio of the mean port length to the mean port diameter) revealed that the exponent m makes a transition from Mode-1 (stagnation mode, m = 0.5) to Mode-2 (wall jet mode, m = 0.8) with increasing Re and decreasing LID. The results suggest that the transition depends on the heat transfer mechanism. With the increase in Re, Nu in the wall jet region increases more rapidly than that in the stagnation region because of the larger Re exponent m. The mode makes a transition in the smaller Re for smaller L/D because the wall jet region becomes predominant with smaller LID.©2012 by the International Astronautical Federation. - 円筒デトネーション波の遷移及び伝播条件に関する研究
棧敷和弥, 桧物恒太郎, 脇田督司, 戸谷剛, 永田晴紀, 燃焼シンポジウム講演論文集, 50th, 306, 307, 2012年11月20日
日本語 - 一定レイノルズ数のよどみ点近傍における固体燃料の燃焼速度におよぼす酸化剤流速の影響
寺川健, 永田晴紀, 戸谷剛, 脇田督司, 金子雄大, 燃焼シンポジウム講演論文集, 50th, 82, 83, 2012年11月20日
日本語 - 相似形状のCAMUI型ハイブリッドロケットの燃焼特性におけるスケール効果
永田晴紀, 脇田督司, 戸谷剛, 植松努, 燃焼シンポジウム講演論文集, 50th, 78, 79, 2012年11月20日
日本語 - 表面微細周期構造を用いた波長選択性ラジエータ
戸谷剛, 脇田督司, 永田晴紀, 日本機械学会熱工学コンファレンス講演論文集, 2012, 505, 506, 2012年11月16日
A periodic micro structure of the cubic cavity of 6.0 μm x 6.0μm x 6.0μm is built on an ultra violet curable resin by the UV nanoimprint method. A gold film is sputtered on the periodic micro structure. The maximum normal emissivity of the periodic micro structure is 0.6 near 10 μm in the case of the gold film of 200 nm thick on the surface facing the gold source in a sputtering equipment. The resonance wavelength measured in experiments is shorter than that calculated from theory. It may be caused that a part of the bottom of the micro structure is shallower than 6.0 μm and the opening shape is rounded off., 一般社団法人日本機械学会, 日本語 - Scale modeling on flame shape spreading inside fuel tube
Tsuenyoshi Matsuoka, Harunori Nagata, Yuji Nakamura, ISEM-ACEM-SEM-7th ISEM’12-Taipei (The 7th ISEM’ 12-Taipei), 2012年11月, [査読有り]
英語 - Optical properties of wavelength-selective radiator with periodic microcavities
Tsuyoshi Totani, Naoyuki Ishikawa, Minoru Iwata, Masashi Wakita, Harunori Nagata, Proceedings of the 3rd International Forum on Heat Transfer (CD-ROM), IFHT2012-101, 2012年11月, [査読有り]
英語 - 大口径平板燃焼器を伝播する円筒デトネーション波に関する研究
棧敷和弥, 寺坂昭宏, 脇田督司, 戸谷剛, 永田晴紀, 宇宙航空研究開発機構特別資料 JAXA−SP−, JAXA-SP-11-015, 11-015, 27-32, 32, 2012年03月30日
第43回流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム2011 (2011年7月7日-8日. 早稲田大学国際会議場) 東京
43rd Fluid Dynamics Conference / Aerospace Numerical Simulation Symposium 2011 (July 7-8, 2011. Waseda University, International Conference Center), Tokyo Japan
Detonation initiation is one of the important problems of pulse detonation engines. A reflector installed near a predetonator exit is effective to maintain detonation waves. The incident detonation wave from the predetonator transforms its shape from planer to cylindrical and cylindrical to toroidal around the reflector. To prevent the detonation wave downstream of reflector from disappearing, it is important to form a cylindncal detonation wave upstream of the reflector. In this paper, the authors used a large-bore plate combustor and varied the following parameters to find the necessary condition to form a cylindrical detonation wave: the quantity of the driver gas and time from the ball valve opening to the ignition. To evaluate the quantity of the driver gas, an overfilling distance is used. The overfilling distance is the radius of the 10 mm high cylinder whose volume is equivalent to the total amount of the driver gas used in the experiment. Results indicate that the mixing between the driver and target gases affects the ability of successful transition. When a cylindrical detonation wave is formed, an annular soot track structure appears in the vicinity of the predetonator exit. This structure and velocity of detonation waves show that the transformation of the incident detonation wave into the cylindrical detonation wave resuls from the twice reflection from the plates of the combustor. To form a cylindrical detonation Wave, the driver gas should fill the volume much larger than the annular structure area.
形態: カラー図版あり
Physical characteristics: Original contains color illustrations
資料番号: AA0065207005
レポート番号: JAXA-SP-11-015, 宇宙航空研究開発機構, 日本語 - CAMUIロケット開発者 : 火薬を使わないロケットの開発
永田 晴紀, 妹尾 俊明, 三宅 龍馬, 日本機械学會誌 = Journal of the Society of Mechanical Engineers, 115, 1120, 170, 173, 2012年03月05日
一般社団法人日本機械学会, 日本語 - CAMUI型ハイブリッドロケットにおける異常燃焼の克服
永田晴紀, 2, 1, 1, 11, 2012年 - 超小型衛星「ほどよし1号機」の熱設計
小川洋人, 井上遼太, 戸谷剛, 脇田督司, 永田晴紀, 日本機械学会年次大会講演論文集(CD-ROM), 2012, 0, _S192013, 1-_S192013-5, 2012年
"HODOYOSHI-1" is the a micro satellite which is about 60 cm cube and about 60 kg. This satellite is scheduled to be launch at the end of 2012 into a sun-synchronous orbit of the altitude 500 to 600 km for optical remote sensing mission. Using Fortran and thermal analysis software "Thermal Desktop", thermal design has been performed for this satellite. One-nodal analysis and two-nodal analysis are performed using Fortran programs and multi-nodal analysis is carried out by Thermal Desktop for thermal design. One and two-nodal analysis decide optical properties of each surface and thermal conductivity between inside structure and outside structure. As a result, optic properties are decided that the outer surface of outside structure is alodine 1000, and the inner surface of outside structure and the surface of inside structure are black anodized. These analyses decide that GFRP is inserted between the outside structure and the inside structure, and thermal conductive sheet "DENKA BFG20" is inserted between each component and the surface mounted on it. Thermal Desktop calculates temperatures of components on HODOYOSHI-1. Optical property of-X panel is modified from alodine 1000 to white anodized. GFRP is inserted between the battery and the surface mounted on it. The result of these analyses shows that the temperatures of components are within the allowable temperature range. The thermal design of HODOYOSHI-1 have completed for 10 month. This thermal design is useful for the micro and nano satellites producing at a low cost and a short duration., 一般社団法人 日本機械学会, 日本語 - CAMUI型ハイブリッドロケットシステムの小型化手法とその飛行実証
石山達也, 稲場康彦, 井上遼太, 佐々木俊也, 寺川健, 桧物恒太郎, LEE Sang Jun, 金井竜一朗, 脇田督司, 戸谷剛, 永田晴紀, 日本機械学会年次大会講演論文集(CD-ROM), 2012, 0, _S192023, 1-_S192023-5, 2012年
Although many groups are developing Cansat, a can-sized mock satellite, they have few opportunities to test due to difficulties for students to launch Cansats domestically. To provide the chance to launch Cansats, the authors downsized CAMUI type hybrid rocket and created easy-to-use launch system. The new launcher, miniCAMUI, uses gas oxygen (GOX) as oxidizer and high density polyethylene as fuel. Using GOX instead of liquid oxygen contributes to downsizing and weight saving, reduction of turnaround time for launch due to the simplified procedure to fill the oxidizer. A GOX tank connects to a motor through a valve. An air-driven actuator operates the valve miniCAMUI was launched 6 times in June and July 2012. Three of them were serial successful launches with two rockets in a day, with a turnaround time about 45 minutes. Two of the three launches were with the same rocket in the day. With the wind velocity of 1 to 2 m/s, the apogee altitude was about 74 m, being very close to the predetermined altitude of 80 m. This result shows that miniCAMUI was successfully developed as a small launch system with high operability. miniCAMUI is available for launches to various altitudes below 250 m., 一般社団法人 日本機械学会, 日本語 - 極低温に冷却されたCAMUI型固体燃料の点火特性
金井竜一朗, 寺川健, 稲場康彦, 石山達也, 佐々木俊也, 脇田督司, 戸谷剛, 永田晴紀, 日本機械学会年次大会講演論文集(CD-ROM), 2012, 0, _S192022, 1-_S192022-4, 2012年
During development of a large size CAMUI-type hybrid rocket motor in 2006, anomalous combustion associated with large pressure spike frequently occurred. After troubleshooting, low temperature of the fuel was found to have caused anomalous combustion. Because a cryogenic liquid oxygen tank is around the motor, polyethylene fuel grain is cooled to liquid oxygen temperature. The anomalous combustion was duplicated with a downsized motor. The size of the fuel grain was 2/5 of large size motor. Because Damkohler number is proportional to the ratio of pressure and oxygen mass flux, the ratio for the subscale motor was conformed to that for the full-scale motor. In a reproductive experiment, pressure spike also appeared. This result suggests that blow-off occurs before the pressure spike. After the blow-off, gas mixture in the chamber may ignite to cause the pressure spike., 一般社団法人 日本機械学会, 日本語 - 太陽同期軌道を周回する超小型衛星の熱解析と熱設計に関する研究
井上遼太, 小川洋人, 戸谷剛, 脇田督司, 永田晴紀, 日本機械学会年次大会講演論文集(CD-ROM), 2012, 0, _S192011, 1-_S192011-5, 2012年
This paper presents a thermal design method of nano and micro satellites by one-nodal and two-nodal thermal analyses. The orbit is the sun-synchronous and circular orbit at an altitude of 500 km and at the local time at descending node of 11 o'clock. The attitude is pointing to the Earth. The combinations of optical properties, solar absorptivity and infrared emissivity, on surfaces of Inner Structure and Outer Structure under which the temperature of a satellite is within an allowable temperature range are clarified. The results show a correspondence relation between analytical methods and forms of heat transfer (heat conduction or radiation) between Inner Structure and Outer Structure. A thermal design in which there is little difference between analytical value and experimental value is realized by setting appropriate thermal control products on the contact faces between Inner Structure and Outer Structure. The combinations of optical properties are larger in the design focusing on radiation than in that focusing on heat conduction. A configuration of Inner Structure influences the combination. Finally, thermal design method by the one-nodal and two-nodal analyses which allows one to develop a satellite in short period at low cost is proposed., 一般社団法人 日本機械学会, 日本語 - 短期開発を実現する少節点解析と多節点解析を併用した超小型衛星の熱設計法
小川洋人, 井上遼太, 戸谷剛, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD−ROM), 56th, ROMBUNNO.2M07, 2012年
日本語 - 金属膜を付与した表面微細周期構造の垂直放射率特性
戸谷剛, 石川直幸, 脇田督司, 永田晴紀, 日本伝熱シンポジウム講演論文集(CD−ROM), 49th, ROMBUNNO.D125, 2012年
日本語 - CAMUI型燃料グレインの熱伝達特性に及ばす流路形状とレイノルズ数の影響
佐々木俊也, 大島伸行, 永田晴紀, 脇田督司, 宇宙科学技術連合講演会講演集(CD−ROM), 56th, ROMBUNNO.3H01, 2012年
日本語 - 大口径パルスデトネーションエンジン用イニシエータにおける円筒デトネーション波の伝播に関する研究
桧物恒太郎, 棧敷和弥, 脇田督司, 戸谷剛, 永田晴紀, 流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム講演集(CD−ROM), 44th-2012, ROMBUNNO.1E05, 2012年
日本語 - パネルディスカッション:ハイブリッドロケット研究開発の今後を語ろう
嶋田徹, 北川幸樹, 湯浅三郎, 那賀川一郎, 永田晴紀, 福地亜宝郎, 和田豊, 宇宙科学技術連合講演会講演集(CD-ROM), 56th, ROMBUNNO.3H13, 2012年
日本語 - 推力5000N級CAMUI型ハイブリッドロケットの開発と打上げ実証
永田晴紀, 脇田督司, 戸谷剛, 植松努, 安本裕紀, 三橋龍一, 宇宙科学技術連合講演会講演集(CD−ROM), 56th, ROMBUNNO.3H04, 2012年
日本語 - 産官学連携による準軌道型再使用宇宙輸送システム開発の提案
米本浩一, 相良慎一, 松本剛明, 永田晴紀, 越智徳昌, 石本真二, 麥谷高志, 牧野隆, 木元順一, 日本航空宇宙学会年会講演会講演集(CD−ROM), 43rd, ROMBUNNO.B08, 2012年
日本語 - Thermal Analyses of Nano and Micro Satellites on Sun-synchronous Orbit by One Nodal Analysis Method
Tsuyoshi Totani, Hiroto Ogawa, Ryota Inoue, Masashi Wakita, Harunori Nagata, The proceedings of 1st International Conference on Mechanical Engineering and Renewable Energy, ICMERE2011-PI-148, 2011年12月, [査読有り]
英語 - Optimal fuel grain design method for CAMUI type hybrid rocket
Harunori Nagata, Shunsuke Hagiwara, Nasashi Wakita, Tsuyoshi Totani, Tsutomu Uematsu, 47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit 2011, 2011年12月01日
The alternative fuel grain design in CAMUI type hybrid rockets consists of multiple stages of cylindrical fuel blocks with two ports. Regression formulas as functions of local O/F were developed for a 2500 N thrust class flight model motor. Static firing tests with fuel grains of different scaling showed the validity of the similarity rule, which is available for subscale firing tests, based on convective heat transfer mechanisms. Convective heat transfer rate to the downstream end face of the rearmost block is limited comparing with other burning surfaces and radiative heat transfer is not negligible. As a result, the similarity rule is not valid for this burning surface. Because the impinging jet onto the upstream end face of the uppermost block is not high temperature combustion gas but virtually pure oxygen, a similarity about chemical reaction is necessary besides those about convective heat transfer to realize a similarity condition. These results serve as foundation for the methodology to design optimal fuel grain shape for CAMUI type hybrid rockets. © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. - 大口径大推力パルスデトネーションエンジン用イニシエータにおける円筒デトネーション波の伝播に関する研究
寺坂昭宏, 桟敷和弥, 脇田督司, 戸谷剛, 永田晴紀, 燃焼シンポジウム講演論文集, 49th, 328, 329, 2011年11月20日
日本語 - E135 数値解析によるPMMA管内の燃え拡がり速度予測(OS-9:燃焼の最近の進展(4))
松岡 常吉, 村上 翔太, 永田 晴紀, 中村 祐二, 熱工学コンファレンス講演論文集, 2011, 125, 126, 2011年10月28日
We studied numerically the flame spreading phenomena in a narrow circular duct made by PMMA in order to capture the time-depended transport processes during the event. The main aim of this study is to clarify the stability mechanism near the extinction limit (i.e. low Damkohler number limit) which is one of unique feature in this system, which might be different from the conventional flat-plate flame spreading as suggested in our previous works. Numerical model includes time-dependent mass and heat transfer with one-step finite rate of chemical reactions both in gas and solid phase. In this study, it is successfully simulated the nearly steady flame movement under the condition studied. Moreover, it is accomplished to reproduce the two distinctive spreading modes, such as the one belongs in thermal regime and chemical regime, which have already been observed experimentally. Although there is quantitative disagreement between experimental and numerical results, it is confirmed that our numerical model should be satisfactory to capture the qualitative behavior of flame spreading in the narrow duct., 一般社団法人日本機械学会, 日本語 - 9-105 北海道大学工学部における高大連携活動((09)高大院連携,口頭発表論文)
永田 晴紀, 菅原 広剛, 西口 規彦, 工学教育研究講演会講演論文集, 23, 59, 182, 183, 2011年08月22日
公益社団法人日本工学教育協会, 日本語 - 微小重力環境を利用した固体燃焼現象研究(H22研究班WG報告)
藤田修, 中村祐二, 永田晴紀, 菊池政雄, 伊藤昭彦, 鳥飼宏之, 梅村章, 高橋周平, 池田光優, CHUNG Suk Ho, OLSON Sandra L, 宇宙利用シンポジウム, 27th, 31, 32, 2011年03月
第27回宇宙利用シンポジウム (2011年1月24日-25日, 宇宙航空研究開発機構宇宙科学研究所相模原キャンパス), 相模原市, 神奈川県
The Twenty-seventh Space Utilization Symposium (January 24-25, 2011. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan
Since solid combustion is dominated by diffusion process of pyrolyzed gas as well as heat transfer process around the combustion region, flammability limit becomes very different depending on the gravitational conditions. In the present work, an attempt to know the mechanism of the extension of ignition limit of overloaded wire has been made by numerical calculation. The results showed that the Joule energy more than a critical value causes release of degradation gas and its ignition, and the critical value becomes smaller in microgravity than that in normal gravity. Further, the preparation status of the ISS experiment including the wire ignition and flame spread over solid material is introduced.
著者人数: 11人
形態: カラー図版あり
Number of authors: 11
Physical characteristics: Original contains color illustrations
資料番号: AA0065129013, 宇宙航空研究開発機構宇宙科学研究所 (JAXA)(ISAS), 日本語 - 急性肝機能障害を来した24歳女性
本城晴紀, 永田紘子, 渡邉直昭, 野内俊彦, 日本内科学会関東支部関東地方会, 582nd, 2011年 - 大口径平板燃焼器を伝播する円筒デトネーション波に関する研究
棧敷和弥, 寺坂昭宏, 脇田督司, 戸谷剛, 永田晴紀, 流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム講演集(CD-ROM), 43rd-2011, ROMBUNNO.1B03, 2011年
日本語 - CAMUIロケットを利用したエジェクタ・ジェット試験の数値解析
長谷川進, 谷香一郎, 平岩徹夫, 植田修一, 永田晴紀, 流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム講演集(CD-ROM), 43rd-2011, ROMBUNNO.2D02, 2011年
日本語 - 推力900N級CAMUI型ハイブリッドロケットによる打上げ・海上回収実験
永田晴紀, 前田祐義, 鈴木恭兵, 五十地輝, 安中俊彦, 稲石卓也, 清尾陽平, 植松努, 宇宙科学技術連合講演会講演集(CD-ROM), 55th, ROMBUNNO.3B10, 2011年
日本語 - 太陽同期軌道を周回する超小型衛星の熱設計
井上遼太, 小川洋人, 戸谷剛, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD-ROM), 55th, ROMBUNNO.3S12, 2011年
日本語 - CAMUI型固体燃料後退速度式の導出方法に関する検討
野原正寛, 出雲弘一, 脇田督司, 戸谷剛, 永田晴紀, 日本機械学会年次大会講演論文集(CD-ROM), 2011, 0, _S192013, 1-_S192013-5, 2011年
Regression formulas for solid fuels in CAMUI type hybrid rockets were developed. A fuel grain in this rocket consists of multiple stages of cylindrical fuel blocks with two ports. A Fuel block in a CAMUI type grain has three burning surface, i.e., the upstream end face, port inner walls, and the downstream end face. A series of static firing tests by laboratory model motor and 2500 N thrust class flight model motor gave empirical constants in the regression formulas. There regression formulas are used to obtain an optimal design of a grain configuration. However, the empirical constants provided by these two series of firing tests did not agree with each other. To investigate the cause of this disagreement, additional static firing tests by 2500 N thrust class flight model motor was conducted. Results show that the empirical constants depend on Re number. In contrast, the dependence was not observed in the firing tests by the laboratory model motor. This difference may be caused by the difference in L/D (the ratio of the length of the port and the diameter of port)., 一般社団法人 日本機械学会, 日本語 - 太陽同期軌道を周回する超小型衛星の熱設計に関する研究
小川洋人, 井上遼太, 戸谷剛, 脇田督司, 永田晴紀, 日本機械学会年次大会講演論文集(CD-ROM), 2011, 0, _S192022, 1-_S192022-5, 2011年
The thermal analyses of micro and nano satellites on sun-synchronous orbits are performed by using one nodal analysis. Three models of the satellites are considered. The size and mass of model A, B, C are respectively 0.1 m cube and 1 kg, 0.25 m cube and 25 kg, 0.5 m cube and 50 kg. The analyses are carried out under the conditions of the altitude of 300, 500, 700, and 1 000 km, and the local time of descending node (LTDN) from 6 to 12. The combinations of solar absorptivity and infrared emissivity in the case that the temperature of the satellite is within the allowable temperature range (273.15-313.15 K) increase with the larger parameter of the mass times the heat capacity over one area on surface of satellite. The combinations increase with the higher altitudes and decrease with larger LTDN. The combinations in the case of the orbits without an eclipse are larger than with an eclipse. These results are useful for the preliminary thermal design of micro and nano satellites., 一般社団法人 日本機械学会, 日本語 - Automatic Control on Circulation of Working Fluid in Liquid Droplet Radiator
Tsuyoshi Totani, Takuhiro Takekoshi, Masashi Wakita, Harunori Nagata, Proceedings of the 13th Asian Congress of Fluid Mechanics, 10, 28, 586, 589, 2010年12月, [査読有り]
Liquid Droplet Radiator (LDR) is an important candidate for disposing large quantities of waste heat more than 1 MW which will be handled by large space structures such as Space Power Satellite. The working fluid is heated through a heat exchanger by the waste heat generated in a large structure in space. Then, the working fluid is emitted in space through nozzles of the droplet generator toward a droplet collector as multiple streams of droplets. During the flight in space, the droplets lose thermal energy via radiative heat transfer. After the cooled droplets are captured by the droplet collector, the working fluid is recycled to the heat exchanger by a circulating pump. The automatic control system on the circulation of working fluid in a liquid droplet radiator has been built using a programmable logic controller. The proportional control of flow rate with the term of the variation of the counter flow in the gear pump and the relaxation of the change of an target flow rate has succeeded within 5 percent at the automatic circulation control of the working fluid from 100 ml/min to 200 ml/min and from 200 ml/min to 100 ml/min., THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, 英語 - Micro flame spreading in solid fuel ducts
Tsuneyoshi Matsuoka, Harunori Nagata, 61st International Astronautical Congress 2010, IAC 2010, 4, 2863, 2869, 2010年12月01日
Recently, a novel type of flame called a 'micro flame' has been found and studied because of its anomalous properties. It has been know that micro flame is formed when the Reynolds number = O (1 to 2) and Froude number »1 by using a Bunsen burner. Micro flame is a diffusion flame observed under a normal gravitational field; however, the form becomes small-spherical, which is analogous to flame under micro gravitational field. This is because the buoyancy force due to gravity becomes negligible compared to forced convection. Because of the micro flame's pseudo-microgravity behavior, it is expected to be an alternate method for micro gravitational experiments. As energy from combustion is greater than electrical energy, micro flame could be expected to be a small size and high output energy source for such applications as micro- electromechanical system (MEMS). Though micro flame studies have been conducted previously, they have been limited to investigations of conditions, properties, and mechanisms of steady (non-dynamic) micro flames created using a Bunsen burner. Dynamic micro flames created by flame spreading have not previously been characterized. We have found micro flames could be formed within the flame spreading region of solid fuel ducts. In this study, we investigated the conditions and properties of the non-steady (i.e. dynamic) micro flames. These micro flames form and propagate even when the Reynolds number is more than 100 for the flame spreading in the solid fuel ducts. It is believed that in this instance, the effect of heat transfer becomes larger than for the Bunsen burner cases, since the direction of the mass and heat transfer is confined. As the effect of gravity is very small, micro flame spreading is different from normal flame spreading. For example, flame spreading velocity is much lower and the form of the flame becomes small-spherical. Although there are several different kinds of flame spreading in fuel ducts, such as the chemical regime, thermal regime, and stabilized regime, these results suggest that micro flame would be a novel kind of flame spreading. - Effect of jet velocity on scale effect in oxidizer impinging region
Yudai Kaneko, Mitsunori Itoh, Massasi Wakita, Tsuyoshi Totani, Harunori Nagata, Advances in the Astronautical Sciences, 138, 629, 634, 2010年12月01日
Diffusion combustion in a stagnation point boundary layer of a gaseous oxygen jet over a solid fuel was investigated to clarify effects of jet velocity on a similarity condition of fuel regression rates. This combustion field simulates the upstream-end face of the uppermost fuel block of CAMUI type hybrid rocket fuel grain. Increasing the flow velocity from 5.5 m/s to 11.5 m/s caused an increase in the regression rate from 0.22 mm/s to 0.26 mm/s. This result shows that the chemical reaction effect is not negligible in oxidizer impinging region. - Regression progress of fuel grain in CAMUI type hybrid rocket motor
Harunori Nagata, Akihito Kakikura, Mitsunori Ito, Yudai Kaneko, Kazuhiro Mori, Kenta Ueshima, Tsutomu Uematsu, Tsuyoshi Totani, Advances in the Astronautical Sciences, 138, 611, 616, 2010年12月01日
Static firing tests clarified how the fuel flow rate varies with the progress of fuel regression in a 'cascaded multistage impinging-jet' (CAMUI) type hybrid rocket motor. The fuel gasification rate decreases with progressing fuel regression because of two causes. One is decreasing gas flow density in ports. The other is decreasing area of end faces. The fuel gasification rate decreases rapidly when end faces disappear. A simple model of the regression progress was proposed. Fuel grains collected after firing tests with various burning duration approved this model. The model serves as a foundation to develop regression formulas applicable to this unconventional type fuel grain. - Development of 2500 N class CAMUI type hybrid rocket for winged flight experiments
H. Nagata, M. Wakita, T. Totani, T. Uematsu, K. Yonemoto, 61st International Astronautical Congress 2010, IAC 2010, 3, 2144, 2148, 2010年12月01日
The authors have developed CAMUI type hybrid rockets as a non-toxic propellant sounding rocket system. A main purpose is to drastically downsize the cost and scale of rocket experiments and attract potential users in various research fields. A key idea is a distinctive fuel grain design to accelerate gasification rates of solid fuels. The grain design, designated as CAMUI as an abbreviation of "cascaded multi-stage impinging-jet", makes the combustion gas collide repeatedly with fuel surfaces, resulting in intense heat transfer to the fuel. A 2500 N thrust class CAMUI motor was developed for a small scale winged flight test bed. Static firing tests provided successful thrust performance of the motor. Dividing the total impulse by the total propellant consumption gave mean specific impulse to be 254 sec, achieving the target value of 250 sec. This motor is able to launch a winged vehicle of 50 kg to about 1.7 km apogee altitude. The launch experiment is slated in the next fiscal year. Copyright ©2010 by the International Astronautical Federation. All rights reserved. - 流路幅とセルサイズが円環デトネーション波から平面デトネーション波への遷移過程に与える影響
田村正佳, 寺坂昭宏, 脇田督司, 戸谷剛, 永田晴紀, 燃焼シンポジウム講演論文集, 48th, 486-487, 2010年11月20日
日本語 - 亜酸化窒素の触媒分解反応を用いたハイブリッドロケット用点火器の開発
榎本剛矩, 永田晴紀, 戸谷剛, 脇田督司, 徳留真一郎, 燃焼シンポジウム講演論文集, 48th, 542-543, 2010年11月20日
日本語 - 512 樹脂製表面微細周期構造の分光特性(環境工学)
石川 直幸, 戸谷 剛, 永田 晴紀, 脇田 督司, 北海道支部講演会講演概要集, 2010, 49, 133, 134, 2010年11月07日
一般社団法人日本機械学会, 日本語 - 612 遺伝的アルゴリズムを用いたCAMUI型燃料グレインの最適設計(ロボティクス・メカトロニクス/宇宙工学)
野原 正寛, 金子 雄大, 萩原 俊輔, 脇田 督司, 戸谷 剛, 永田 晴紀, 北海道支部講演会講演概要集, 2010, 49, 161, 162, 2010年11月07日
一般社団法人日本機械学会, 日本語 - 613 拡大する環状流路を伝播する円環デトネーション波に流路幅が与える影響(ロボティクス・メカトロニクス/宇宙工学)
寺坂 昭宏, 田村 正佳, 脇田 督司, 戸谷 剛, 永田 晴紀, 北海道支部講演会講演概要集, 2010, 49, 163, 164, 2010年11月07日
一般社団法人日本機械学会, 日本語 - 115 無火薬式小型ロケットによる宇宙工学研究プロジェクトのロバスト化(宇宙構造物・宇宙用材料の安全性と信頼性,信頼性フォーラム)
永田 晴紀, 学術講演会講演論文集, 59, 0, 219, 220, 2010年05月21日
社団法人日本材料学会, 日本語 - 微小重力環境を利用した固体燃焼現象研究 (H21研究班WG報告)—Solid Combustion Research in Microgravity(2009 Research WG Report
藤田, 修, 中村, 祐二, 永田, 晴紀, 菊池, 政雄, 伊藤, 昭彦, 鳥飼, 宏之, 梅村, 章, 高橋, 周平, 池田, 光優, Fujita, Osamu, Nakamura, Yuji, Nagata, Harunori, Kikuchi, Masao, Ito, Akihiko, Torikai, Hiroyuki, Umemura, Akira, Takahashi, Shuhei, Ikeda, Mitsumasa, Chung, Suk Ho, 宇宙利用シンポジウム = Space Utilization Research: Proceedings of Space Utilization Symposium, 26, 137, 138, 2010年02月
第26回宇宙利用シンポジウム(2010年1月25日-26日, 宇宙航空研究開発機構宇宙科学研究本部相模原キャンパス)
The Twenty-sixth Space Utilization Symposium (January 25-26, 2010: ISAS/JAXA Sagamihara, Japan)
Since solid combustion is dominated by diffusion process of pyrolyzed gas as well as heat balance around combustion area, flammability limit becomes very different depending on the gravitational conditions. In the present work, the attempt to obtain ignition limit for overloaded electric wire and extinction limits for spreading flame over flat sheet has been made under the limit of 4.5 sec microgravity time provided by MGLAB. According to the experiments, it is found that the flammable limits, ignition and extinction limits, significantly extend in microgravity in comparison with those in normal gravity.
形態: カラー図版あり
共催: 日本学術会議
Physical characteristics: Original contains color illustrations
joint hosting: The Science Council of Japan
資料番号: AA0064730020, 宇宙航空研究開発機構宇宙科学研究本部, 日本語 - CAMUI型ハイブリッドロケット燃料の後退速度に及ぼす圧力の影響
永田晴紀, 金子雄大, 萩原俊輔, 伊藤光紀, 脇田督司, 戸谷剛, 植松努, 宇宙科学技術連合講演会講演集(CD−ROM), 54th, ROMBUNNO.2B15, 2010年
日本語 - 紫外線硬化樹脂に形成された表面微細周期構造の透過率特性
戸谷剛, 石川直幸, 脇田督司, 永田晴紀, 日本伝熱シンポジウム講演論文集(CD−ROM), 47th, ROMBUNNO.C321, 285, 2010年
UVナノインプリント法によって作成された表面微細周期構造の太陽光吸収率と全半球赤外線放射率について報告する。表面微細周期構造の周期は、1 マイクロメートルと10 マイクロメートルである。, 社団法人 日本伝熱学会, 日本語 - 超小型宇宙探査機「しんえん(UNITEC‐1)」の熱設計と通信途絶原因の一考察
戸谷剛, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD−ROM), 54th, ROMBUNNO.2C01, 2010年
日本語 - 流路幅と流路傾斜角が円環デトネーション波の伝播に与える影響
寺坂昭宏, 田村正佳, 脇田督司, 戸谷剛, 永田晴紀, 流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム講演集(CD−ROM), 42nd-2010, ROMBUNNO.1B2, 2010年
日本語 - 液滴ラジエータの作動流体循環における自動制御可能範囲の拡大
竹腰卓博, 戸谷剛, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD−ROM), 54th, ROMBUNNO.1K08, 2010年
日本語 - S1901-1-1 超小型宇宙機UNITEC-1の熱設計と軌道上温度(小型宇宙システム(1),社会変革を技術で廻す機械工学)
戸谷 剛, 脇田 督司, 永田 晴紀, UNITEC-1開発チーム, 年次大会講演論文集, 2010, 0, 359, 360, 2010年
UNITEC-1 (UNIsec Technology Experiment Carrier-1) is an interplanetary probe that is developed by 22 universities in UNISEC (University Space Engineering Consortium). The dimensions and weight of UNITEC-1 are about 35cm×35cm×40cm and 22kg, respectively. UNITEC-1 was launched by H-IIA rocket on May 21, 2010 as a piggyback payload of Planet-C (Akatsuki) Venus probe developed by JAXA/ISAS. It is expected that the temperature of UNITEC-1 in the worst hot condition near Earth is 40.6 degree Celsius at the battery. UNITEC-1 will pass outside the orbital path of Earth partly in our expecting orbit and will become the lowest temperature. The temperature of UNITEC-1 in the worst cold condition at the furthest point from the sun will be 3.8 degree Celsius. The temperature of UNITEC-1 in the worst hot condition near Venus will be 84.8 degree Celsius., 一般社団法人 日本機械学会, 日本語 - S1901-1-2 小型CAMUI式ハイブリッドロケットの開発と運用(小型宇宙システム(1),社会変革を技術で廻す機械工学)
竹腰 卓博, 佐藤 峻哉, 田村 正佳, 萩原 俊輔, 脇田 督司, 戸谷 剛, 永田 晴紀, 植松 努, 年次大会講演論文集, 2010, 0, 361, 362, 2010年
The authors developed a 1.3-m-long small scale CAMUI hybrid rocket of 4.7 kg in total weight to obtain an easy test launch system for CanSat. Two test launches without CanSat made altitude less than 250 m. Launches below 250 m is not regulated by the Civil Aeronautics Act in Japan. The avionics system loaded in the rocket consisted of a microcomputer, acceleration sensor, angular velocity sensor and altitude sensor. The avionics system can do wireless communication using the Bluetooth technology, real-time onboard data were obtained during the test flights successfully., 一般社団法人 日本機械学会, 日本語 - S1901-1-5 CAMUI型固体燃料の燃料後退予測式取得方法に関する検討(小型宇宙システム(1),社会変革を技術で廻す機械工学)
萩原 俊輔, 金子 雄大, 野原 正寛, 永田 晴紀, 戸谷 剛, 脇田 督司, 松岡 常吉, 植嶋 健太, 植松 努, 年次大会講演論文集, 2010, 0, 367, 368, 2010年
The authors have been developing CAMUI type hybrid rockets which have a new fuel grain design to accelerate burning rate. Regression formulae for CAMUI type rockets have been developed as functions of local O/F. The authors carried out combustion tests with various oxidizer flow rate and burning duration to obtain empirical constants of these regression formulae. Additionally, a simulation model of static firing was built using the fuel regression formulae. In these studies, regression rates were mean values during firing. This paper discusses the effect of firing duration on the accuracy of regression formulae by using the simulation model., 一般社団法人 日本機械学会, 日本語 - S1901-2-5 ハイブリッドエンジン搭載型小型有翼ロケットの飛行実験(小型宇宙システム(2),社会変革を技術で廻す機械工学)
米本 浩一, 永田 晴紀, 渡辺 大地, 年次大会講演論文集, 2010, 0, 379, 380, 2010年
The Space Systems Laboratory of Kyushu Institute of Technology has been studying unmanned suborbital winged rocket as a research subject of future fully reusable space transportation system since 2005. The flight tests of a small scaled winged rocket were conducted five times from 2008 to 2009. A larger winged rocket with a hybrid rocket developed by Hokkaido University, which will reach to a higher altitude, is under development for validating INS/GPS hybrid navigation system, real time trajectory generation and guidance algorithm using GA implemented on FPGA, H_∞ and adaption control theory. This paper reports current design, development and flight test plan of the winged rocket., 一般社団法人 日本機械学会, 日本語 - G1900-1-3 ラバールノズルの超音速域における伝熱が推力および比推力に与える影響(宇宙工学部門一般講演,社会変革を技術で廻す機械工学)
岩城 裕樹, 戸谷 剛, 脇田 督司, 永田 晴紀, 年次大会講演論文集, 2010, 0, 407, 408, 2010年
The variation of the thrust and the specific impulse was revealed numerically with calculating the change of the exhaust velocity. Both cooling and heating the propellant flowing in the divergent section of Laval nozzle were treated. The expansion ratio were 30, 600, and 2000. The specific heat ratio was 1.3. Two types of heat profile were considered; pulsed heat transfer (PHT) and distributed heat transfer (DHT). The relations of Rayleigh flow and isentropic change were used for PHT. The exhaust velocity is higher than the isentropic value in the case that the heat is provided near the throat. In other cases, the exhaust veolocity is less than the isentropic case. The equivalent point of heat transfer is introduced for DHT. Results of DHT is coincident with PHT by using this equivalent point. This results indicates that the effect with DHT can be predicted from PHT., 一般社団法人 日本機械学会, 日本語 - Combustion characteristics of the end burning hybrid rockets at laminar flow
Tsuneyoshi Matsuoka, Harunori Nagata, 60th International Astronautical Congress 2009, IAC 2009, 8, 6293, 6300, 2009年12月01日
In this study, we aim to clarify the blowoff mechanism for flame spreading in an opposed laminar flow in narrow solid fuel ducts. To clarify this mechanism we conducted two experiments. The first, we observed the changes of flame spread rate at various oxygen velocity, ambient pressure and port diameter. For the flame spreading at laminar flow, combustion modes could be classified into 3 distinct regimes based on the strength of the opposed flow, i..e. chemical regime, thermal regime, stabilized regime. This result is consistent with the result at turbulent flow. In stabilized regime, quenching distance is almost constant despite the oxygen velocity. In order to investigate the effect of ambient pressure and port diameter of fuels on blowoff limit, transition oxygen velocity and the parameters are obtained. As a result, transition oxygen velocity is proportional to the logarithm of the ambient pressure and port diameter. This relation is applicable despite the flow condition. Furthermore, we calculated velocity gradient at the fuel sur1ce to reveal the determining 1ctor of the blowoff limit at laminar flow. Consequently, velocity gradient, which is considered to dominate flow separation at laminar flow would not be at constant. This result are not coincident to the fact that friction velocity, which dominates flow separation in the turbulent flow, and thus blowoff limit. - Numerical simulation of the ejector-jet flowfield around small rocket exhaust
Susumu Hasegawa, Kouichiro Tani, Kenji Kudo, Noboru Sakuranaka, Shuichi Ueda, Harunori Nagata, 16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference, 2009年12月01日
In order to get experimental data in subsonic Mach number flight region, two small hybrid rockets installed the ejector-jet were launched, and then the flight analysis was performed. Numerical simulations were conducted for clarification of the flowfields and prediction of the ejector-jet performance. The CFD results revealed the compound compressible flow phenomena involving oblique shock waves and the subsonic flows. The effects of the flight Mach numbers and the rocket pressures to the suction performances were also investigated. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc. - Subsonic flight experiments of ejector-rocket using hybrid-rocket CAMUI
Shuichi Ueda, Tetsuo Hiraiwa, Masao Takegoshi, Kouichiro Tani, Takeshi Kanda, Harunori Nagata, 16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference, 2009年12月01日
This paper presents an overview of the subsonic flight experiments for ejector-rocket using CAMUI-type hybrid-rocket. The flight experiment was planned to obtain ejector characteristics under subsonic flight condition, for developing design technology of rocket based combined-cycle engine. An ejector-duct was attached to the engine of the rocket. Two experimental vehicles were launched at 16 th March 2009, from the coast of Hokkaido island of Japan. Flight experiments were successfully conducted and ejector characteristics data at the flight condition were acquired. Copyright © 2009 by the authors. - Improvement of thrust and specific impulse by convective heat transfer in Laval nozzle
Yuuki Iwaki, Tsuyoshi Totani, Tetsushi Naganuma, Syunya Sato, Masashi Wakita, Harunori Nagata, 60th International Astronautical Congress 2009, IAC 2009, 8, 6563, 6573, 2009年12月01日
The effect of the convective heat transfer from the nozzle wall to the flow in the supersonic region of Laval nozzle has been investigated using one-dimensional numerical analysis program. Nitrogen is assumed as the propellant. The nozzles that have the throat diameter of 0.5 mm and 2.0 mm are used. The ranges of the thrust and the specific impulse are 0.35 to 10 N and 73 to 140 s under adiabatic theory, respectively. The prediction formula for the ratio of the increment of the stagnation temperature in the nozzle to the stagnation temperature at the nozzle inlet is obtained from the basic equations. This prediction formula satisfies the trends of the analysis results; the more the heat transfer in the nozzle is, the more the improvement of the exhaust velocity is. However, the more energy loss, or which cannot be converted to the kinetic energy increases with the increment of the heat transfer. The maximum ratio of the increment of the stagnation temperature in this analysis is 0.50 and the exhaust velocity improved by a factor of 1.12 when the throat diameter is 0.5 mm, the inlet temperature is 300 K, inlet pressure is 1 MPa, and the expansion ratio is 100. - Numerical simulation of CAMUI-type hybrid rocket combustor
Kouichi Kishida, Kouichi Kishida, Yudai Kaneko, Yudai Kaneko, Nobuyuki Oshima, Nobuyuki Oshima, Harunori Nagata, Harunori Nagata, 45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 2009年12月01日
The combustor of new type of hybrid rocket, Cascaded Multistage Impinging Jet(CAMUI), is studied numerically. The rocket has a characteristic combustor so that it can overcome the traditional issue of hybrid rockets, low thrust power. The numerical method is expected to clarify the physical states in the combustion chamber during its operation and to accelerate the development of rockets in the future. This paper will discuss the preliminary stage of numerical simulation using a computational mesh that was based on real measured changes in the chamber shape with the consumption of fuel obtained with a 3-dimensional measuring system. The flow structure related to the changing chamber shape is clarified using this numerical method. © 2009 by the American Institute of Aeronautics and Astronautics, Inc. - 619 CAMUI式ハイブリッドロケットの小型化および繰り返し運用(移動・輸送のメカトロニクス)
佐藤 峻哉, 竹腰 卓博, 田村 正佳, 萩原 俊輔, 脇田 督司, 戸谷 剛, 永田 晴紀, 北海道支部講演会講演概要集, 2009, 48, 211, 212, 2009年11月28日
一般社団法人日本機械学会, 日本語 - 620 小型ハイブリッドロケットにおける能動的ロール制御システムの構築(移動・輸送のメカトロニクス)
田村 正佳, 佐藤 俊哉, 竹腰 卓博, 萩原 俊輔, 脇田 督司, 戸谷 剛, 永田 晴紀, 北海道支部講演会講演概要集, 2009, 48, 213, 214, 2009年11月28日
一般社団法人日本機械学会, 日本語 - ライデンフロスト温度以下に冷却された固体燃料の燃焼特性
飯島直純, 金子雄大, 脇田督司, 戸谷剛, 永田晴紀, 燃焼シンポジウム講演論文集, 47th, 302-303, 2009年11月18日
日本語 - 反射板を有する円管状燃焼器におけるデトネーション遷移過程の研究
浅田隆利, 脇田督司, 戸谷剛, 永田晴紀, 坪井伸幸, 林光一, 衝撃波シンポジウム講演論文集, 2008, 211-214, 2009年03月17日
日本語 - 円錐形状反射板を用いたPDEイニシエータによるドライバーガス削減
脇田督司, 浅田隆利, 戸谷剛, 永田晴紀, 衝撃波シンポジウム講演論文集, 2008, 239-240, 2009年03月17日
日本語 - 微小重力環境を利用した固体燃焼現象研究(H20研究班WG報告)—Solid combustion research in microgravity (2008 Research WG Report)
藤田, 修, 中村, 祐二, 永田, 晴紀, 菊池, 政雄, 伊藤, 昭彦, 鳥飼, 宏之, 梅村, 章, 高橋, 周平, 池田, 光優, Fujita, Osamu, Nakamura, Yuji, Nagata, Harunori, Kikuchi, Masao, Ito, Akihiko, Torikai, Hiroyuki, Umemura, Akira, Takahashi, Shuhei, Ikeda, Mitsumasa, Chung, Suk Ho, 宇宙利用シンポジウム = Space Utilization Research: Proceedings of Space Utilization Symposium, 25, 2009年03月
第25回宇宙利用シンポジウム(2009年1月14日-15日, 宇宙航空研究開発機構宇宙科学研究本部相模原キャンパス)
The Twenty-fifth Space Utilization Symposium (January 14-15, 2009: ISAS/JAXA Sagamihara, Japan)
Since solid combustion is dominated by diffusion process of pyrolyzed gas as well as heat balance around combustion area, which are strongly affected by convective flow, micorgravity could be an effective tool to understand its mechanism. One of the most important contributions of solid combustion research in microgravity is fire safety in space. In the present report, some researches on going regarding fire safety in space will be introduced.
資料番号: AA0064297124, 宇宙航空研究開発機構宇宙科学研究本部, 日本語 - ノズル加熱による推力および比推力の向上に関する評価方法の検討
長沼哲史, 岩城裕樹, 佐藤峻哉, 戸谷剛, 脇田督司, 永田晴紀, 宇宙科学技術連合講演会講演集(CD−ROM), 53rd, 3G03, 2009年
日本語 - 自己加圧供給を利用した液体酸素供給開始方式の検討
和久宏之, 金子雄大, 飯島直純, 萩原俊輔, 脇田督司, 戸谷剛, 永田晴紀, 宇宙科学技術連合講演会講演集(CD−ROM), 53rd, 2B08, 2009年
日本語 - ラバールノズルにおける推進剤への対流熱伝達による推力・比推力の向上
岩城裕樹, 長沼哲史, 佐藤峻哉, 戸谷剛, 脇田督司, 永田晴紀, 日本航空宇宙学会北部支部講演会ならびに再使用型宇宙推進系シンポジウム講演論文集, 2009-10th, 112-117, 2009年
日本語 - 環状流路を有するPDEイニシエーター内のデトネーション波の挙動
田村正佳, 脇田督司, 浅田隆利, 戸谷剛, 永田晴紀, 流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム講演集, 41st-2009, 357-360, 2009年
日本語 - CAMUI型ハイブリッドロケット用固体燃料の後退履歴特性
永田晴紀, 柿倉彰人, 伊藤光紀, 金子雄大, 森一大, 植嶋健太, 植松努, 戸谷剛, 宇宙科学技術連合講演会講演集(CD−ROM), 53rd, 2B05, 2009年
日本語 - UNITEC‐1の初期熱設計
伊井晴明, 田島康晴, 戸谷剛, 永田晴紀, 脇田督司, 日本航空宇宙学会北部支部講演会ならびに再使用型宇宙推進系シンポジウム講演論文集, 2009-10th, 281-286, 2009年
日本語 - S1901-2-5 CAMUI型ハイブリッドロケット燃焼室の数値解析(小型宇宙システム(2))
岸田 耕一, 金子 雄大, 大島 伸行, 永田 晴紀, 年次大会講演論文集, 2009, 0, 259, 260, 2009年
CAMUI (Cascaded Multistage Impinging-jet) type hybrid rocket has been developed in Hokkaido University. The 3-dimensional shape of combustion chamber of hybrid rocket changes considerably during the operation because it is constructed of solid fuel itself. The shape of combustion chamber is related with the performance of the rocket directly, so it is important to understand the relationship between flow and shape. To clarify this issue, numerical simulations were conducted using three different 3D shapes. These shapes were measured from partially burned fuel blocks by the non-contact measuring system. These blocks were obtained by the combustion test where the blocks were taken out before they completely burn out., 一般社団法人 日本機械学会, 日本語 - S1904-2-2 ラバールノズルでの対流熱伝達を利用した推力・比推力の向上(大気突入・減速技術(2))
岩城 裕樹, 長沼 哲史, 佐藤 峻哉, 戸谷 剛, 脇田 督司, 永田 晴紀, 年次大会講演論文集, 2009, 0, 297, 298, 2009年
The one-dimensional analysis has been performed using the analysis program for the flow of the propellant of Laval nozzles. The propellant is nitrogen. Two kinds of nozzle that has a throat diameter of 0.5 mm and 1 mm are assumed. Heat flow to the propellant by convective heat transfer in the supersonic region is increased by elongating of the divergent section of the nozzle. The ratio of the exhaust velocity with the heat transfer in the nozzle to the exhaust velocity without the heat transfer is not affected by the throat diameter, inlet pressure, and temperature difference between the nozzle wall and the propellant. In the case that the throat diameter of 0.5 mm, inlet pressure of 1 MPa, temperature difference of 300 K and length of the divergent section of 200 mm, the ratio of the exhaust velocity becomes 1.05., 一般社団法人 日本機械学会, 日本語 - The improvement of the thrust and specific impulse by heating and elongating of De Laval nozzle
Yuuki Iwaki, Tsuyoshi Totani, Tetsushi Naganuma, Harunori Nagata, International Astronautical Federation - 59th International Astronautical Congress 2008, IAC 2008, 10, 6403, 6410, 2008年12月01日
The numerical analysis program for the flow in the Laval nozzle was constructed to research the effect of the heat transfer in the nozzle. The nitrogen is assumed as the propellant and the aluminum alloy is adopted as the material of the nozzle. Supplied heat to the propellant in the nozzle increases with elongating the length of the divergent nozzle. The heat can contribute to the improvement of the specific impulse and the thrust. The effect of the heat supplied in the nozzle appears significantly under the condition of the low enthalpy of the propellant. The elongation of the length of the divergent nozzle from 10 mm to 100 mm changes the thrust from 14.41 mN to 17.02 mN and changes the specific impulse from 84.43 s to 99.69 s in the condition that the diameter of the nozzle throat is 0.1 mm, the inlet pressure is 1 MPa, the inlet temperature of the propellant is 300 K, and the inlet temperature of the nozzle wall is 600 K. - CAMUI型ハイブリッドロケット固体燃料の燃焼速度に及ぼすスケールの影響
永田晴紀, 伊藤光紀, 金子雄大, 柿倉彰人, 森一大, 植松努, 戸谷剛, 燃焼シンポジウム講演論文集, 46th, 56-57, 2008年11月20日
日本語 - 二段燃焼式ハイブリッドロケット一次燃焼室における燃料定常分布に初期粒径が及ぼす影響
羽柴健太, 堺裕哉, 戸谷剛, 永田晴紀, 燃焼シンポジウム講演論文集, 46th, 516-517, 2008年11月20日
日本語 - 605 二段燃焼式ハイブリッドロケット一次燃焼室における燃料後退速度の粒径依存性(宇宙工学・環境)
羽柴 健太, 堺 裕哉, 戸谷 剛, 永田 晴紀, 北海道支部講演会講演概要集, 2008, 47, 149, 150, 2008年09月27日
一般社団法人日本機械学会, 日本語 - 606 反射板を用いたPDEイニシエーターにおける反射板径の最適化に関する研究(宇宙工学・環境)
浅田 隆利, 脇田 督司, 戸谷 剛, 永田 晴紀, 北海道支部講演会講演概要集, 2008, 47, 151, 152, 2008年09月27日
一般社団法人日本機械学会, 日本語 - CAMUI型ハイブリッドロケットの開発 : 異常燃焼の発生とその克服
永田 晴紀, 伝熱 : journal of the Heat Transfer Society of Japan, 47, 199, 23, 29, 2008年04月01日
日本伝熱学会, 日本語 - 反射板を用いたPDEイニシエーターにおけるドライバーガス削減に関する研究
浅田隆利, 脇田督司, 沼倉龍介, 戸谷剛, 永田晴紀, 衝撃波シンポジウム講演論文集, 2007, 275-278, 2008年03月17日
日本語 - B-2-24 90・250kgf級CAMUI型ハイブリッドロケットへの超小型モデル衛星搭載実験(B-2. 宇宙・航行エレクトロニクス,一般セッション)
三橋 龍一, 佐藤 立博, 竹浪 恭平, 安部 潤一郎, 吉尾 直人, 戸谷 剛, 永田 晴紀, 電子情報通信学会総合大会講演論文集, 2008, 1, 266, 266, 2008年03月05日
一般社団法人電子情報通信学会, 日本語 - 高Ohnesorge数領域における液滴流の捕集と飛散を分ける閾値—Threshold between spreading and splashing of droplet streams in the region of high Ohnesorge number
戸谷, 剛, 南部, 航太, 川上, 哲人, 由利, 泰史, 永田, 晴紀, Totani, Tsuyoshi, Nanbu, Kota, Kawakami, Akihito, Yuri, Yasufumi, Nagata, Harunori, 宇宙利用シンポジウム 第24回 平成19年度 = Space Utilization Research: Proceedings of the Twenty-fourth Space Utilization Symposium, 117, 120, 2008年03月
This research has aimed to prevent the working fluid from dispersing at the droplet collector of liquid droplet radiators by specifying the threshold in which splashing and spreading of droplets in space are divided. The following results have been achieved by this research. (1) It was clarified not to be able to arrange the threshold in which splashing and spreading were divided only by a past arrangement type of the threshold. (2) The threshold approved to fluids with the different viscosity is discovered. (3) It is discovered that the droplets splashes more hardly under microgravity than under normal gravity. (4) The cause that the threshold between splashing and spreading under microgravity are different from that under normal gravity has been specified. The above-mentioned results greatly contribute to the proof examination on the orbit of the liquid droplet radiator.
資料番号: AA0063706028, 宇宙航空研究開発機構宇宙科学研究本部, 日本語 - バルブレス技術による小型ロケットの開発
永田晴紀, バルブ技報, 23, 1, 15, 20, 2008年 - CAMUI方式によるハイブリッドロケット燃料後退速度向上の定量的評価
森一大, 伊藤光紀, 柿倉彰仁, 金子雄大, 植嶋健太, 室井典和, 植松努, 戸谷剛, 永田晴紀, 日本航空宇宙学会北部支部講演会講演論文集, 2008, 135-138, 2008年
日本語 - CAMUIハイブリッドロケットの起動特性シミュレーション
植嶋健太, 伊藤光紀, 前田剛典, 柿倉彰仁, 金子雄大, 森一大, 室井典和, 植松努, 戸谷剛, 永田晴紀, 日本航空宇宙学会北部支部講演会講演論文集, 2008, 139-142, 2008年
日本語 - 液滴ラジエータにおける液滴流の飛散と捕集を分ける閾値
川上哲史, 由利泰史, 仁木雄大, 戸谷剛, 永田晴紀, 日本航空宇宙学会北部支部講演会講演論文集, 2008, 149-153, 2008年
日本語 - CAMUIハイブリッドロケットの性能履歴予測に関する研究
柿倉彰仁, 伊藤光紀, 金子雄大, 森一大, 植嶋健太, 飯嶋直純, 室井典和, 植松努, 戸谷剛, 永田晴紀, 日本航空宇宙学会北部支部講演会講演論文集, 2008, 143-147, 2008年
日本語 - 4037 ラバールノズルの伸長および加熱による推力・比推力の向上(S60-1 小型宇宙システム(1),21世紀地球環境革命の機械工学:人・マイクロナノ・エネルギー・環境)
岩城 裕樹, 長沼 哲史, 戸谷 剛, 永田 晴紀, 年次大会講演論文集, 2008, 0, 395, 396, 2008年
The one-dimensional analysis has been performed using the analysis program for the flow of the propellant and the wall temperature of Laval nozzles. The material of the nozzles is A5056 and the propellant is water. In the case that the diameter of the nozzle inlet is 0.6 mm, the diameter of the throat is 0.1 mm, the mass flow rate of the propellant is 1 g/min, the inlet temperature of the propellant and wall are 500 K, Outlet pressure of the propellant is 400 Pa, the radial thickness of the nozzle is 2 mm, and the length of the divergent nozzle is 36 mm, the 151 s of the specific impulse has been achieved whereas the specific impulse with an adiabatically change is 137 s., 一般社団法人 日本機械学会, 日本語 - 4038 大学連携プロジェクト有翼ロケット実験機開発計画の現状(S60-2 小型宇宙システム(2),21世紀地球環境革命の機械工学:人・マイクロナノ・エネルギー・環境)
脇田 督司, 米本 浩一, 麻生 茂, 幸節 雄二, 永田 晴紀, 鵜沢 潔, 年次大会講演論文集, 2008, 0, 397, 398, 2008年
The project of Winged Experimental Rocket described here is a proposal by the alliance of universities (University Consortium) expanding and integrating the research activities of reusable space transportation system performed by individual universities, and is the proposal that aims at flight proof of the results of advanced research conducted by the universities and JAXA using the university-centered experimental launch systems. This paper verifies the validity of the winged experimental rocket by surveying the technical issues that should be demonstrated and by estimating the airframe scale, weight and structural type. To minimize the development risks of winged experimental rocket, two kinds of airframe with different scales are developed., 一般社団法人 日本機械学会, 日本語 - 4040 CAMUIハイブリッドロケットの性能履歴予測モデルを用いた燃料グレイン設計(S60-2 小型宇宙システム(2),21世紀地球環境革命の機械工学:人・マイクロナノ・エネルギー・環境)
植嶋 健太, 永田 晴紀, 伊藤 光紀, 柿倉 彰仁, 金子 雄大, 森 一大, 飯島 直純, 室井 典和, 植松 努, 戸谷 剛, 年次大会講演論文集, 2008, 0, 401, 402, 2008年
In order to enlarge the system of CAMUI hybrid rocket, it is essential to estimate its performance with a minimum of combustion experiments to reduce research and development costs. The authors had built up "Prediction Model for Startup Characteristic", which simulate the pressure history of each part of the system, and "Fuel Regression History Model", which predict the weight and shape history of fuel under combustion. In this study, the authors combined the two models mentioned above, and built up the program that predict the pressure history of each part of the system, and the weight and shape history of the fuel directly from the initial parameters. The result obtained by the program agreed well with the experimental value and got adequacy. With the program, the authors designed the CAMUI fuel grain which realize the optimal fuel flow rate on 90kgf thrust class motor, of which fuel has not been optimized yet., 一般社団法人 日本機械学会, 日本語 - 4041 超音波パルス反射法を用いたハイブリッドロケット燃料の後退履歴の取得(S60-2 小型宇宙システム(2),21世紀地球環境革命の機械工学:人・マイクロナノ・エネルギー・環境)
金子 雄大, 伊藤 光紀, 植嶋 健太, 戸谷 剛, 永田 晴紀, 年次大会講演論文集, 2008, 0, 403, 404, 2008年
A series of the firing rest was conducted to obtain the history of the instantaneous value of the fuel thickness during the firing test. An ultrasonic pulse-echo method was used to the firing test for that purpose. Gaseous oxygen and polyethylene were used to the firing test about the combustion of the solid fuel in the oxygen jet collision area. A preliminary experiment revealed that the variation of the propagation time caused by the change of fuel temperature is negligible small because the thermal boundary layer in the solid fuel is thin. The fuel regression rate at the stagnation point of oxidizer jet depends on Reynolds number., 一般社団法人 日本機械学会, 日本語 - 推力250kgf級CAMUI型ハイブリッドロケットの燃焼特性
永田晴紀, 植松努, 伊藤光紀, 柿倉彰人, 金子雄大, 森一大, 戸谷剛, 燃焼シンポジウム講演論文集, 45th, 508-509, 2007年11月20日
日本語 - 特集 Reusable Launch Vehicle (RLV) CAMUIロケット開発の現状とその将来構想
伊藤 献一, 永田 晴紀, 植松 努, 日本航空宇宙学会誌, 55, 646, 292, 295, 2007年11月
日本航空宇宙学会, 日本語 - 250kgf級CAMUI型ハイブリッドロケットへの超小型モデル衛星搭載実験
三橋龍一, 佐藤立博, 竹浪恭平, 安部潤一郎, 吉尾直人, 戸谷剛, 永田晴紀, 電気・情報関係学会北海道支部連合大会講演論文集(CD−ROM), 2007, ROMBUNNO.47, 2007年10月27日
日本語 - パネルディスカッション『宇宙輸送系に将来はあるのか?』
久保田 弘敏, 棚次 亘弘, 有田 誠, 永田 晴紀, 稲谷 芳文, 苅田 丈士, 日本航空宇宙学会誌 = Journal of the Japan Society for Aeronautical and Space Sciences, 55, 645, 275, 285, 2007年10月05日
日本航空宇宙学会, 日本語 - 503 液滴ラジエータの作動流体循環流量の自動制御に関する研究(宇宙工学)
由利 泰史, 川上 哲史, 戸谷 剛, 永田 晴紀, 北海道支部講演会講演概要集, 2007, 46, 129, 130, 2007年09月29日
一般社団法人日本機械学会, 日本語 - 504 水を推進剤とした太陽熱推進軌道変換機によるペイロード輸送ミッションの設計(宇宙工学)
岩城 裕樹, 戸谷 剛, 永田 晴紀, 北海道支部講演会講演概要集, 2007, 46, 131, 132, 2007年09月29日
一般社団法人日本機械学会, 日本語 - 505 傾斜壁面に衝突する液滴流の飛散と捕集を分ける閾値に関する研究(宇宙工学)
川上 哲史, 由利 泰史, 戸谷 剛, 永田 晴紀, 北海道支部講演会講演概要集, 2007, 46, 133, 134, 2007年09月29日
一般社団法人日本機械学会, 日本語 - 506 二段燃焼式ハイブリッドロケット一次燃焼室における粒状燃料のガス化履歴(宇宙工学)
片野 光, 羽柴 健太, 戸谷 剛, 永田 晴紀, 北海道支部講演会講演概要集, 2007, 46, 135, 136, 2007年09月29日
一般社団法人日本機械学会, 日本語 - 507 CAMUI方式を用いたハイブリッドロケット燃料の燃焼速度に及ぼすスケールの効果(宇宙工学)
森 一大, 伊藤 光紀, 柿倉 彰仁, 金子 雄大, 植嶋 健太, 室井 典和, 戸谷 剛, 植松 努, 永田 晴紀, 北海道支部講演会講演概要集, 2007, 46, 137, 138, 2007年09月29日
一般社団法人日本機械学会, 日本語 - 502 CAMUIハイブリッドロケット燃料グレインの形状履歴の予測(宇宙工学)
柿倉 彰仁, 伊藤 光紀, 金子 雄大, 森 一大, 植嶋 健太, 室井 典和, 植松 努, 戸谷 剛, 永田 晴紀, 北海道支部講演会講演概要集, 2007, 46, 127, 128, 2007年09月29日
一般社団法人日本機械学会, 日本語 - カムイロケット燃焼室における固体燃料の燃焼
永田 晴紀, 日本燃焼学会誌 = Journal of the Combustion Society of Japan, 49, 149, 163, 169, 2007年08月31日
日本燃焼学会, 日本語 - CAMUI型ハイブリッドロケットの小型可動翼による自律制御システムの開発に関する研究
武岡 和彦, 佐藤 立博, 安部 潤一郎, 竹浪 恭平, 三橋 龍一, 佐鳥 新, 大滝 誠一, 豊田 国昭, 中村 明広, 永田 晴紀, 北海道工業大学研究紀要, 35, 0, 361, 364, 2007年03月26日
CAMUI hybrid rocket is safety, low-cost and low-pollution rocket that was developed in Hokkaido. In this research, it aims at the construction of the system that lands in a target place, it safely and surely of rocket. The basic research is described, change of control wing by enlargement of rocket and development of autonomous control system., 北海道工業大学, 日本語 - 反射板を用いたPDEイニシエーターによる爆轟波の伝播促進
脇田督司, 沼倉龍介, 菅田成俊, 戸谷剛, 永田晴紀, 衝撃波シンポジウム講演論文集, 2006, 47-48, 2007年03月15日
日本語 - LESを用いたCAMUIハイブリッドロケット燃焼器内の熱流体解析
脇田督司, 伊藤光紀, 金子雄大, 中島卓司, 永田晴紀, 大島伸行, 流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム講演集, 39th-2007, 373-376, 2007年
日本語 - 4131 超音波パルス反射法を用いたハイブリッドロケット燃料後退履歴の取得に関する研究(S74-3 小型宇宙システム(3),S74 小型宇宙システム)
金子 雄大, 伊藤 光紀, 柿倉 彰仁, 森 一大, 植嶋 健太, 戸谷 剛, 永田 晴紀, 年次大会講演論文集, 2007, 0, 407, 408, 2007年
A hybrid rocket is a high safety and low cost propulsion system. However, conventional hybrid rockets have a defect of are low thrust because of the low gasification rate of the solid fuel. Therefore, many researches of a hybrid rocket aimed at the improvement of the gasification rate of the solid fuel, being evaluated by a regression rate. We have been developing a new type hybrid rocket named CAMUI that improves the regression rate of solid fuels by using an impinging jet heat transfer. In many cases, a fuel regression rate is obtained as an average during the combustion. However, when the fuel shape changes dramatically during the combustion, this method is not suitable. Therefore, it is necessary to measure the history of the fuel thickness during the combustion. Therefore, we are trying to use ultrasonic pulse-echo measurement system and this paper describes the result of a basic research about the real-time measurement., 一般社団法人 日本機械学会, 日本語 - 4130 CAMUI方式ハイブリッドロケットの燃料後退特性に及ぼすスケール効果(S74-3 小型宇宙システム(3),S74 小型宇宙システム)
伊藤 光紀, 前田 剛典, 柿倉 彰仁, 金子 雄大, 森 一大, 植松 努, 戸谷 剛, 永田 晴紀, 年次大会講演論文集, 2007, 0, 405, 406, 2007年
CAMUI type hybrid rocket has a distinctive configuration of fuel grain to overcome the defect of the low fuel regression rate of conventional hybrid rockets. In CAMUI type fuel grain, a number of surfaces perpendicular to the thrust axis, as well as port surfaces, contribute as burning surfaces. Static firing tests with three scales of analogous motors were conducted to investigate the scaling effect on the fuel regression characteristics of CAMUI type fuel grain. LOX and Polyethylene were used as a propellant, and tests were conducted with the same Reynolds number condition, at the chamber pressure of 1MPa and the mass flux of 150-300kg/(m^2s)., 一般社団法人 日本機械学会, 日本語 - SP1 推力400kgf級CAMUIロケット打上げ実証試験(1)(特別講演)
永田 晴紀, 流体工学部門講演会講演論文集, 2006, 0, "SP1, a", 2006年10月28日
一般社団法人日本機械学会, 日本語 - SP1 推力400kgf級CAMUIロケット打上げ実証試験(2)(特別講演)
永田 晴紀, 流体工学部門講演会講演論文集, 2006, 0, "SP1, 1"-"SP1-3", 2006年10月28日
A joint research team of universities and private companies in Hokkaido, Japan has been organized to develop a small-scale reusable launch system based on CAMUI hybrid rocket. The main purpose is to drastically reduce the cost of rocket experiments and thus attract potential users such as metrological and microgravity researchers. The meteorological observation model of 400-kgf class motor is under development and a launch test in 10 km altitude is planned. Anomalous combustion occurred in the last four static firing tests out of seven serial tests with a flight model motor. This anomalous ..., 一般社団法人日本機械学会, 日本語 - 117 HIT-SATフライトモデルの構造設計と機械環境試験(宇宙工学)
榊原 隆浩, 戸谷 剛, 安中 俊彦, 佐鳥 新, 永田 晴紀, 北海道支部講演会講演概要集, 2006, 45, 33, 34, 2006年09月25日
一般社団法人日本機械学会, 日本語 - 116 液滴ラジエータにおける液滴流回収時の飛散と捕集に関する研究(宇宙工学)
南部 航太, 川上 哲史, 由利 泰史, 戸谷 剛, 永田 晴紀, 北海道支部講演会講演概要集, 2006, 45, 31, 32, 2006年09月25日
一般社団法人日本機械学会, 日本語 - 118 二段燃焼式ハイブリッドロケットのEMの設計(宇宙工学)
坂本 将司, 片野 光, 戸谷 剛, 永田 晴紀, 北海道支部講演会講演概要集, 2006, 45, 35, 36, 2006年09月25日
一般社団法人日本機械学会, 日本語 - 119 マルチセルインフレータブル構造の剛性に関する研究(宇宙工学)
加藤 隆造, 戸谷 剛, 石村 康生, 永田 晴紀, 北海道支部講演会講演概要集, 2006, 45, 37, 38, 2006年09月25日
一般社団法人日本機械学会, 日本語 - 120 デトネーション波の伝播促進に及ぼすドライバーガス供給量の影響(宇宙工学)
菅田 成俊, 脇田 督司, 沼倉 龍介, 永田 晴紀, 戸谷 剛, 北海道支部講演会講演概要集, 2006, 45, 39, 40, 2006年09月25日
一般社団法人日本機械学会, 日本語 - B-2-1 CAMUI型ハイブリッドロケットスケールモデル機体による落下制御実験(B-2.宇宙・航行エレクトロニクス,一般講演)
武岡 和彦, 佐藤 立博, 難波江 亮, 三橋 龍一, 佐鳥 新, 大滝 誠一, 豊田 国昭, 中村 明広, 永田 晴紀, 電子情報通信学会総合大会講演論文集, 2006, 1, 239, 239, 2006年03月08日
一般社団法人電子情報通信学会, 日本語 - B-2-2 400kgf級CAMUI型ハイブリッドロケットのアビオニクス搭載実験(B-2.宇宙・航行エレクトロニクス,一般講演)
三橋 龍一, 佐藤 立博, 武岡 和彦, 大野 努, 中村 明広, 佐鳥 新, 永田 晴紀, 電子情報通信学会総合大会講演論文集, 2006, 1, 240, 240, 2006年03月08日
一般社団法人電子情報通信学会, 日本語 - 微小重力環境を利用した固体燃焼現象検討WG報告—WG report on Solid Combustion in Microgravity
藤田, 修, 中村, 祐二, 永田, 晴紀, 菊池, 政雄, 伊藤, 昭彦, 岡島, 敏, 梅村, 章, 高橋, 周平, Fujita, Osamu, Nakamura, Yuji, Nagata, Harunori, Kikuchi, Masao, Ito, Akihiko, Okajima, Satoshi, Umemura, Akira, Takahashi, Shuhei, 宇宙利用シンポジウム = Space Utilization Research: Proceedings of Space Utilization Symposium, 22, 2006年03月
第22回宇宙利用シンポジウム(2006年1月17日-19日, 日本学術会議6階会議室 六本木、東京)
The Twenty-second Space Utilization Symposium (January 17-19, 2006: Science Council of Japan, Roppongi, Tokyo, Japan)
Since solid combustion, which has long time scale, is dominated by diffusion process of pyrolyzed gas, micorgravity could be an effective tool to understand its mechanism. One of the most important contributions of solid combustion research in microgravity is fire safety in space. In the present report, some researches attained previously as well as future expected subjects regarding fire safety in space will be introduced. Because individual physical processes included in the solid combustion have different time scale, combination of short-term and long-term microgravity experiments is requisite to proceed the research on solid combustion in microgravity.
形態: カラー図版あり
共催: 日本学術会議
Physical characteristics: Original contains color illustrations
Meeting sponsors: The Science Council of Japan, The Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (ISAS)(JAXA)
資料番号: AA0064113029, 宇宙航空研究開発機構宇宙科学研究本部, 日本語 - The Challenges of Developing a Hybrid Launch System for Microsatellite Payloads Using CAMUI Hybrid Upperstages with the Rocketplane XP
永田 晴紀, Proceedings of the 25th ISTS Selected Papers, 821, 827, 2006年, [査読有り] - CAMUI型ハイブリッドロケット小型可動翼自律制御システムの検討
武岡和彦, 佐藤立博, 安部潤一郎, 竹浪恭平, 三橋龍一, 佐鳥新, 大滝誠一, 豊田国昭, 中村明広, 永田晴紀, 加藤聡貴, 大野努, 宇宙科学技術連合講演会講演集(CD-ROM), 50th, 2006年 - スペースプレーンのための革新的な基盤技術研究開発における大学およびベンチャの役割
棚次亘弘, 麻生茂, 新井隆景, 永田晴紀, 姫野武洋, 日本航空宇宙学会年会講演会講演集, 37th, 2006年 - 1024 CAMUI方式を用いた推力400kgf級サウンディングロケットの研究開発(S87 小型宇宙システム,S87 小型宇宙システム)
前田 剛典, 伊藤 光紀, 柿倉 彰仁, 難波江 亮, 永田 晴紀, 戸谷 剛, 工藤 勲, 植松 努, 年次大会講演論文集, 2006, 0, 341, 342, 2006年
Cascaded Multistage Impinging-jet (CAMUI) is one of the combustion methods of hybrid rocket. This is the way of burning solid fuel and liquid oxidizer at stagnation point. By this way, we can expect to get high regression rate and high combustion efficiency. Now, with this combustion method, we developed the 400kgf thrust class flight model to observe upper air and did experiments to confirm its performance. This summer, our aim is to make this hybrid rocket reach 10km altitude. This paper describes the outline of the mission and the combustion characteristics which we got by the static firing test., 一般社団法人 日本機械学会, 日本語 - Development of CAMUI hybrid rocket to create a market for small rocket experiments
Harunori Nagata, Mitsunori Ito, Takenori Maeda, Mikio Watanabe, Tsutomu Uematsu, Tsuyoshi Totani, Isao Kudo, International Astronautical Federation - 56th International Astronautical Congress 2005, 7, 4763, 4767, 2005年12月01日
By introducing various innovative ideas, the difficult-to-develop small hybrid-type rocket is successfully developed. The main purpose is to drastically reduce the cost of rocket experiments and thus attract potential users such as metrological and microgravity researchers. A key idea is a new fuel grain design to accelerate the gasification rate of solid fuel. The new fuel grain design, designated as CAMUI as an abbreviation of "Cascaded Multistage Impinging-jet", is that the gas flow repeatedly collides with the solid fuel surface to accelerate the heat transfer to the fuel. To install a regenerative cooling system using cryogenic liquid oxygen as coolant in a small launcher, the authors devised a valveless supply system (with no valves in the liquid oxygen flow line). Four serial successful launch verification tests by 10 kg vehicle equipped with a 50 kgf thrust CAMUI motor have shown the feasibility of the motor system. The meteorological observation model of 400 kgf class motor is under development and the development of microgravity experiment class of 1.5 to 2 tonf motor will follow subsequently. The authors plan to complete the development of the 400 kgf class motor for meteorological observation model by the end of FY2005. - Numerical and experimental studies on circulation of working fluid in liquid droplet radiator
Tsuyoshi Totani, Takuya Kodama, Kensuke Watanabe, Kota Nanbu, Harunori Nagata, Isao Kudo, International Astronautical Federation - 56th International Astronautical Congress 2005, 6, 3842, 3852, 2005年12月01日
A model of the circulation of the working fluid in a liquid droplet radiator has been developed. The model is based on Bernoulli's law and the loss of the hydraulic head. The behavior of the circulation of the working fluid calculated from the model is compared with that obtained from experiments in the case that the flow rate of the circulating working fluid is changed. In radiators, the flow rate of the circulating working fluid is changed in order to match the change of the waste heat generated in large-space structures. The flow rates of the circulating working fluid calculated from the model correspond to those obtained from the experiments well. The circulation mechanism of the working fluid in the liquid droplet radiator has been clarified. The model developed in the present work will allow us to control the flow rate of the working fluid in the liquid droplet radiator automatically. - 反射板を用いたPDEにおける爆轟波伝播に及ぼす過供給の影響
沼倉龍介, 脇田督司, 伊藤雄介, 菅田成俊, 永田晴紀, 戸谷剛, 工藤勲, 燃焼シンポジウム講演論文集, 43rd, 498-499, 2005年11月20日
日本語 - 反射板によるデトネーション波の伝播促進効果に対するセルサイズの影響
脇田督司, 沼倉龍介, 伊藤雄介, 菅田成俊, 永田晴紀, 戸谷剛, 工藤勲, 燃焼シンポジウム講演論文集, 43rd, 480-481, 2005年11月20日
日本語 - 216 反射板を用いたPDEイニシエータにおいて入射衝撃波強度が動作特性に及ぼす影響(航空宇宙工学)
伊藤 雄介, 脇田 督司, 沼倉 龍介, 菅田 成俊, 永田 晴紀, 戸谷 剛, 工藤 勲, 北海道支部講演会講演概要集, 2005, 44, 70, 71, 2005年10月08日
一般社団法人日本機械学会, 日本語 - 217 二段燃焼式ハイブリッドロケットの熱設計(航空宇宙工学)
譜久山 尚, 坂本 将司, 永田 晴紀, 戸谷 剛, 工藤 勲, 北海道支部講演会講演概要集, 2005, 44, 72, 73, 2005年10月08日
一般社団法人日本機械学会, 日本語 - 218 液滴ラジエータの循環流量変化時における作動流体循環の過渡特性(航空宇宙工学)
渡辺 健介, 南部 航太, 戸谷 剛, 永田 晴紀, 工藤 勲, 北海道支部講演会講演概要集, 2005, 44, 74, 75, 2005年10月08日
一般社団法人日本機械学会, 日本語 - 219 マウスを用いた回収型動物実験衛星の概念設計検討(航空宇宙工学)
飯田 恭平, 戸谷 剛, 永田 晴紀, 工藤 勲, 矢野 昭起, 北海道支部講演会講演概要集, 2005, 44, 76, 77, 2005年10月08日
一般社団法人日本機械学会, 日本語 - 大学における小型再使用打上げシステムの開発研究 : その1 : CAMUI型ハイブリッドロケットの開発
永田 晴紀, 日本航空宇宙学会誌 = Journal of the Japan Society for Aeronautical and Space Sciences, 53, 616, 142, 146, 2005年05月05日
日本航空宇宙学会, 日本語 - Thermal design of liquid droplet radiator for space solar-power system
T Totani, T Kodama, H Nagata, Kudo, I, JOURNAL OF SPACECRAFT AND ROCKETS, 42, 3, 493, 499, 2005年05月
The waste heat from the space solar-power system, which supplies 5 MW of electricity to a power transmission line on Earth, is estimated, and the liquid droplet radiator for handling the waste heat are examined on the basis of experimental results obtained under microgravity for droplet generation and droplet collection of the liquid droplet radiator. The following results have been obtained. First, an active heat removal system for the power generation unit in the photovoltaic power system is not necessary when the concentration ratio of solar energy is smaller than 1.34, whereas for the liquid droplet radiator, with silicon oil as working fluid, in the solar dynamic power system, the droplet sheet for radiating the waste heat must be 147 m long, 65.1 m wide, and 0.998 m thick. Second, the droplet sheet of the liquid droplet radiator, in which the working fluid is silicon oil, must be 107 m long, 43.2 m wide, and 0.998 m thick to manage the waste heat from the power distribution unit and the power transmission unit in the photovoltaic power system, whereas it must be 107 m long, 65.2 to wide, and 0.998 m thick in the solar dynamic power system., AMER INST AERONAUT ASTRONAUT, 英語 - B-2-38 CAMUI型ハイブリッドロケットによるCanSat実験と自律制御装置の開発(B-2. 宇宙・航行エレクトロニクス, 通信1)
三橋 龍一, 佐藤 立博, 武岡 和彦, 難波江 亮, 下岡 彩子, 中村 明広, 佐鳥 新, 大滝 誠一, 豊田 国昭, 永田 晴紀, 電子情報通信学会総合大会講演論文集, 2005, 1, 324, 324, 2005年03月07日
一般社団法人電子情報通信学会, 日本語 - Research on dynamic response of catalytic heat release rate on platinum wire to a shock wave in hydrogen-air mixture
D Nakamura, H Nagata, T Totani, Kudo, I, JSME INTERNATIONAL JOURNAL SERIES B-FLUIDS AND THERMAL ENGINEERING, 48, 1, 144, 150, 2005年02月
The authors have proposed the use of a hydrogen concentration probe as a new simple method of evaluating hydrogen concentration. We propose a method of evaluating the rate of change of catalytic heat release in order to evaluate hydrogen concentration history for the condition under which the boundary layer is likely to be immature. By using a shock tube, the rate of increase of heat budget of platinum wire was investigated experimentally. Experimental results indicate that the rate of increase of difference due to catalytic heat release is the maximum value when the hydrogen concentration is 30%, which agrees well with a previous result. As a result, this method can be used to evaluate the rate of increase of difference due to catalytic heat release. It is clear that the rate of change due to catalytic heat release is strongly correlated with the rate of change of hydrogen concentration., JAPAN SOC MECHANICAL ENGINEERS, 英語 - CAMUIハイブリッドロケットによる小型ロケット実験市場の創出
永田 晴紀, 渡辺 三樹生, 伊藤 光範, 前田 剛典, 戸谷 剛, 工藤 勲, スペース・エンジニアリング・コンファレンス講演論文集 : Space Engineering Conference, 2004, 13, 1, 4, 2005年01月20日
Small-scale reusable sounding rocket system is under development to provide means of stratosphere observation and three-minutes microgravity experiment. The propulsion system is a hybrid type that uses solid fuel (plastics) and liquid oxygen as propellants and free from explosives, resulting in the dramatically reduced launch cost. To enhance the burning rate of the solid fuel and to augment the thrust, the rocket has employed a new fuel grain design. This new design, named CAMUI as an abbreviation of "Cascaded Multistage Impinging-jet", allows mixing and combustion to occur around stagnati..., 一般社団法人日本機械学会, 日本語 - 急拡大部における爆轟波伝播に及ぼす混合気組成の影響
沼倉龍介, 脇田督司, 伊藤雄介, 菅田成俊, 永田晴紀, 戸谷剛, 工藤勲, 流体力学講演会講演集, 37th, 155-158, 2005年
日本語 - Creation of a market for small rocket experiments through CAMUI hybrid rocket
H Nagata, M Watanabe, M Ito, T Maeda, T Uematsu, T Totani, Kudo, I, 17th ESA Symposium on European Rocket and Balloon Programmes and Related Research, 590, 375, 379, 2005年
By introducing various innovative ideas, the difficult-to-develop small hybrid-type rocket is successfully developed. The main purpose is to drastically reduce the cost of rocket experiments and thus attract potential users such as metrological and microgravity researchers. A key idea is a new fuel grain design to accelerate the gasification rate of solid fuel. The new fuel grain design, designated as CAMUI as an abbreviation of "Cascaded Multistage Impinging-jet", is that the gas flow repeatedly collides with the solid fuel surface to accelerate the heat transfer to the fuel. To install a regenerative cooling system using cryogenic liquid oxygen as coolant in a small launcher, the authors devised a valveless supply system (with no valves in the liquid oxygen flow line). Four serial successful launch verification tests by 10 kg vehicle equipped with a 50 kgf thrust CAMUI motor have shown the feasibility of the motor system. The meteorological observation model of 400 kgf class motor is under development and the development of microgravity experiment class of 1.5 to 2 tonf motor will follow subsequently. The authors plan to complete the development of the 400 kgf class motor for meteorological observation model by the end of FY2005., EUROPEAN SPACE AGENCY, 英語 - CAMUI型(縦列多段衝突噴流型)ハイブリッドロケットの開発と微小重力ロケット実験への応用 (小特集2 短時間微小重力環境を利用した最近の燃焼研究)
永田 晴紀, 日本マイクログラビティ応用学会誌, 22, 1, 47, 50, 2005年
日本マイクログラビティ応用学会, 日本語 - 3321 CAMUI型ハイブリッドロケットによるCanSat実験(S86-1 小型宇宙システム(1),S86 小型宇宙システム)
青柳 賢英, 豊田 国昭, 永田 晴紀, 佐藤 立博, 岩本 隆敏, 竹浪 恭平, 難波江 亮, 下岡 彩子, 佐鳥 新, 三橋 龍一, 大滝 誠一, 年次大会講演論文集, 2005, 0, 369, 370, 2005年
The CanSat project was conducted by undergraduate students who had designed and fabricated a small experimented module called "CanSat". The CanSat was launched by a CAMUI-type hybrid rocket up to the height of 330meters. The CanSat was ejected from the rocket by the releasing mechanism that was installed inside the fairing of rocket, and fell down safely by using a parachute., 一般社団法人 日本機械学会, 日本語 - 3329 太陽熱推進用スラスタの伝熱解析と性能特性(S86-2 小型宇宙システム(2),S86 小型宇宙システム)
高野 千尋, 村木 祐介, 戸谷 剛, 永田 晴紀, 工藤 勲, 年次大会講演論文集, 2005, 0, 385, 386, 2005年
The experiments in which water as propellant is heated and ejected from the nozzle of the solar thermal thruster has been conducted. Thermal analysis of the solar thermal thruster and the propellant has been conducted These results indicate that the highest performance is achieved at a largest mass flow rate for a heat input, where propellant finish boiling at the entrance of the nozzle. Analytical result has indicated that the temperature distribution of the thruster was not ideal for heating propellant because of the heat lost at a nozzle. This is the problem to try to improve propulsive performance. It is possible that this problem is solved by using plug nozzle as a nozzle., 一般社団法人 日本機械学会, 日本語 - 3331 ハイブリッドロケットCAMUIモデルの風洞試験(S86-2 小型宇宙システム(2),S86 小型宇宙システム)
難波江 亮, 豊田 国昭, 植松 努, 永田 晴紀, 年次大会講演論文集, 2005, 0, 389, 390, 2005年
The launching project of a hybrid rocket has been conducted by the research stuffs in the universities in Hokkaido. The CAMUI (CAscaded Multistage Impinging-jet) hybrid rocket has been developed in Hokkaido university, and three launching tests of the rocket showed good results. As the next stage, we are developing a more powerful motor and arocket body to reach to 60km altitude. In the present study, we report the results of the wind test which was carried out to obtain the basicmaterials of the aerodynamic performance of rocket body models., 一般社団法人 日本機械学会, 日本語 - 3332 CAMUI方式を用いた推力70kgf級ロケットモーターの燃料設計(S86-2 小型宇宙システム(2),S86 小型宇宙システム)
伊藤 光紀, 前田 剛典, 永田 晴紀, 戸谷 剛, 工藤 勲, 植松 務, 年次大会講演論文集, 2005, 0, 391, 392, 2005年
CAMUI type hybrid rocket has a distinctive configuration of fuel grain resulting in unique characteristics of fuel regression. In CAMUI type fuel grain, not only port surfaces but also a number of surfaces perpendicular to the thrust axis contribute as burning surfaces. The authors conducted static firing tests of 70kg thrust class CAMUI type motor to obtain quantitative formulas of the fuel regression as a function of oxidizer flow rate and the initial configuration of the port. These formulas enable us to calculate the instantaneous configuration of the grain, fuel flow rate, and any other performances of the firing motor. They are indispensable to obtain an optimal design of the grain for any missions., 一般社団法人 日本機械学会, 日本語 - 3422 液滴ラジエータにおける作動流体の循環特性(S89-2 微小重力・宇宙環境利用(2),S89 微小重力・宇宙環境利用)
戸谷 剛, 児玉 拓也, 渡辺 健介, 南部 航太, 永田 晴紀, 工藤 勲, 年次大会講演論文集, 2005, 0, 445, 446, 2005年
Experiments have been carried out under normal gravity in order to examine characteristics on circulation of working fluid in a liquid droplet radiator. The experimental setup had functioned under microgravity. The working fluid is silicon oil. It has been clear that the liquid droplet radiator has the function to stabilize flow rate without flow rate controllers. Flow rate controllers will be used in a real machine. Even if the flow rate controllers is malfunctioned, the liquid droplet radiator circulates the working fluid stably., 一般社団法人 日本機械学会, 日本語 - 航空宇宙工学便覧第三版
日本航空宇宙学会, 2005年 - CAMUIロケットによる小型ロケット実験市場の創出
永田 晴紀, 化学工学 = CHEMICAL ENGINEERING OF JAPAN, 68, 12, 738, 739, 2004年12月05日
日本語 - 反射板を用いたPDEイニシエータにおける爆轟波の消炎と再点火の機構
脇田督司, 沼倉龍介, 永田晴紀, 戸谷剛, 工藤勲, 衝撃波シンポジウム講演論文集, 2003, 129-132, 2004年03月18日
日本語 - B-2-9 有翼ハイブリッドロケット打ち上げ実験(B-2.宇宙・航行エレクトロニクス)
三橋 龍一, 中村 直紀, 難波江 亮, 久保田 昌志, 鈴木 智貴, 佐藤 立博, 佐鳥 新, 大滝 誠一, 豊田 国昭, 永田 晴紀, 電子情報通信学会総合大会講演論文集, 2004, 1, 293, 293, 2004年03月08日
一般社団法人電子情報通信学会, 日本語 - Measurement technique for pumping performance of a centrifugal collector under microgravity
T Totani, M Itami, H Nagata, Kudo, I, A Iwasaki, REVIEW OF SCIENTIFIC INSTRUMENTS, 75, 2, 515, 523, 2004年02月
A measurement technique for obtaining the pumping performance of a centrifugal collector under microgravity has been developed and evaluated through microgravity experiments. These tests have been conducted under conditions such that the pressure sensors cannot easily detect the pressure rise of the liquid working fluid. These conditions have a pressure increase smaller than 400 Pa. The characteristic of the head produced in a centrifugal collector calculated from experimental data agrees well with that predicted theoretically from the velocity and the pressure generated by rotation of the centrifugal collector. It is concluded from this result that the measurement technique can correctly obtain the pumping performance of the centrifugal collector under microgravity. The centrifugal collector has produced the head of 0.041 m at the rotation speed of 223 rpm under microgravity. The working fluid is silicon oil. This head corresponds to the pressure rise of approximately 390 Pa. (C) 2004 American Institute of Physics., AMER INST PHYSICS, 英語 - 2段燃焼式ハイブリッドロケットの酸化剤自己加圧供給系に関する研究(S68-3 ハイブリッドロケットおよび回収技術,S68 学生による宇宙活動)
要 貴浩, 豊田 国昭, 秋葉 鐐二郎, 永田 晴紀, 年次大会講演論文集, 2004, 0, 479, 480, 2004年
The basic experiments have been conducted to develop the staged combustion hybrid rocket engine with two combustion chambers. In the present study, we visualized the flow phenomenon around the orifice used for the flow measurement of the oxidizer N_2O, and the condition to suppress the liquid-gas two-phase flow was obtained. The result is useful to design the self-pressure supply system of N_2O for two-staged combustion hybrid-rocket engine., 一般社団法人 日本機械学会, 日本語 - ハイブリッドロケット用デルタ翼の空力特性(S68-3 ハイブリッドロケットおよび回収技術,S68 学生による宇宙活動)
難波江 亮, 豊田 国昭, 秋葉 鐐二郎, 永田 晴紀, 年次大会講演論文集, 2004, 0, 481, 482, 2004年
The winged-vehicle for the fly-back system of CAMUI (CAscaded Multistage Impinging-jet) hybrid-rocket has been developed. As the first stage, we carried out the gliding test of a Space-shuttle type winged-vehicle. In the test, the winged-vehicle was carried by a radio-controlled airplane, separated, glided and landed. Then, we conducted a launching test using model rocket engine. The results encouraged us to proceed to further development of the fly-back system. As the second stage, we carried out the tests to develop the delta wing assembled to, CAMUI hybrid-rocket. In the wind tunnel test, the aerodynamic characteristics of the delta wing were investigated. Also, the flight performance of the delta wing was checked by the water-rocket test. These results provided useful data to design and make a delta wing of optimal shape. The present results led to the success of the fly-back test of CAMUI hybrid rocket., 一般社団法人 日本機械学会, 日本語 - 火炎先端部の直上流に形成される循環領域が固体燃料の燃え広がりに与える影響
橋本望, 永田晴紀, 戸谷剛, 工藤勲, 燃焼シンポジウム講演論文集, 41st, 509-510, 2003年11月20日
日本語 - 415 高層成層圏観測に用いる完全再使用型 CAMUI ハイブリッドロケットの基本設計
三浦 崇志, 渡辺 三樹生, 永田 晴紀, 戸谷 剛, 工藤 勲, 北海道支部講演会講演概要集, 2003, 43, 154, 155, 2003年09月28日
一般社団法人日本機械学会, 日本語 - 416 マイクロローバーを搭載した火星探査機の概念設計
増田 紀昭, 戸谷 剛, 永田 晴紀, 工藤 勲, 北海道支部講演会講演概要集, 2003, 43, 156, 157, 2003年09月28日
一般社団法人日本機械学会, 日本語 - 419 反射板まわりにおけるデトネーション波の再点火機構
沼倉 龍介, 脇田 督司, 永田 晴紀, 戸谷 剛, 工藤 勲, 北海道支部講演会講演概要集, 2003, 43, 162, 163, 2003年09月28日
一般社団法人日本機械学会, 日本語 - B-2-52 ハイブリッドロケット回収のためのアビオニクス系実験
中村 直紀, 芝 邦明, 下岡 彩子, 三橋 龍一, 佐鳥 新, 大滝 誠一, 豊田 国昭, 永田 晴紀, 電子情報通信学会総合大会講演論文集, 2003, 1, 311, 2003年03月03日
一般社団法人電子情報通信学会, 日本語 - VSRLの利用状況評価および教育プロジェクトにおける用例
吉川 茂雄, 戸谷 剛, 永田 晴紀, 日本ディスタンスラーニング学会会誌, 4, 0, 13, 20, 2003年03月
日本ディスタンスラーニング学会, 日本語 - ハイブリッドロケット回収のためのアビオニクス搭載機の滑空実験
三橋 龍一, 中村 直紀, 芝 邦明, 下岡 彩子, 佐鳥 新, 大滝 誠一, 豊田 国昭, 永田 晴紀, 北海道工業大学研究紀要, 31, 0, 37, 42, 2003年03月
北海道工業大学, 日本語 - Relation of Pt wire temperature and catalytic heat release rate on Pt in unsteady state of H2-air mixture
Daisuke Nakamura, Harunori Nagata, Tsuyoshi Totani, Isao Kudo, Nippon Kikai Gakkai Ronbunshu, B Hen/Transactions of the Japan Society of Mechanical Engineers, Part B, 69, 677, 126, 131, 2003年
The authors have proposed a hydrogen concentration probe using catalytic reaction on Pt wire surface. To use this probe to detect a concentration change in a supersonic mixing layer, the response of the catalytic heat release rate must depend only on concentration change around the probe. Catalytic heat release rate on the Pt wire surface in unsteady state is measured using a constant temperature type hotwire anemometer technique and a shock tube to investigate the relation of the response of the catalytic heat release rate and Pt wire temperature. Catalytic heat release rate begins increasing at the arrival of the shock wave. The increasing rate of the catalytic heat release depends on the Pt wire temperature when the wire temperature is low. However, the dependence is very weak when the wire temperature is over about 680 K. This shows that not the catalytic reaction but molecular transfer from the flow to the Pt wire surface is the controlling step when Pt wire temperature is high enough. As a conclusion, the Pt wire temperature over about 680 K is necessary to use the hydrogen concentration probe in a supersonic mixing layer., Japan Society of Mechanical Engineers, 日本語 - 3316 再使用型 CAMUI ハイブリッドロケットの打上げ・回収試験
渡辺 三樹生, 久保田 勲, 三浦 崇志, 伊藤 光紀, 村木 祐介, 永田 晴紀, 戸谷 剛, 工藤 勲, 芝 邦明, 下岡 彩子, 年次大会講演論文集, 2003, 0, 363, 364, 2003年
To develop a reusable launch system, development study of jet-impinging hybrid rocket has been made. To prove a reliability and safety of a launch-recover system with jet-impinging hybrid rocket motor, a ballistic launch test of CAMUI (Cascaded Multistage Impinging-jet)-02 was performed on a January 13,2003 at TAIKI Hokkaido. The CAMUI-02 went up stably and reached about 500m in altitude. the rocket was recovered safely by parachute. These results prove reliability and safety of the launch-recover system with CAMUI hybrid rocket., 一般社団法人 日本機械学会, 日本語 - 110 ハイブリッドロケット回収系用飛翔体の開発 : 有翼飛翔体の空力特性
下岡 彩子, 芝 邦明, 難波江 亮, 松尾 亮弘, 豊田 国昭, 三橋 龍一, 佐鳥 新, 永田 晴紀, 北海道支部講演会講演概要集, 2002, 42, 20, 21, 2002年10月05日
The launching project of a hybrid rocket from Taiki Multi-purpose Aerospace Park has been conducted by the research groups in the universities in Hokkaido. In the plan, after the flight in outer space, the rocket is taken back using a parafoil-glider with of GPS (Grobal Positioning System). The present study deals with the development of the winged vehicle to obtain the basic materials of the parafoil-glider system., 一般社団法人日本機械学会, 日本語 - 306 反射板によるデトネーション波の伝播促進 : 反射板距離の影響
脇田 督司, 沼倉 龍介, 永田 晴紀, 戸谷 剛, 工藤 勲, 北海道支部講演会講演概要集, 2002, 42, 82, 83, 2002年10月05日
Quick initiation of a detonation wave in a combustion chamber is important to realize a high-performance pulse detonation engine. A possible method is to generate a detonation wave in a shock-tube and release the detonation wave into the chamber. In this paper, a reflecting board is installed in the combustion chamber near the shock-tube exit where the pipe diameter expands sharply. It prevents the detonation wave disappearing at the expanding area near the shock tube exit. The relation of the cell size at the shock-tube exit and the distance between the shock-tube exit and the reflecting b..., 一般社団法人日本機械学会, 日本語 - 308 加圧燃焼器内でのペレット状固体燃料のガス化に関する研究
藤井 篤之, 栗田 慎一郎, 永田 晴紀, 戸谷 剛, 工藤 勲, 北海道支部講演会講演概要集, 2002, 42, 86, 87, 2002年10月05日
The authors have been proposed staged combustion hybrid rocket to overcome defects of conventional hybrid rockets such as low combustion efficiency, Isp loss due to O/F shift, and poor throttling characteristics. This hybrid rocket mainly consists of primary and secondary combustion chambers. The primary combustion chamber, which generates fuel-rich combustion gas, functions as a fuel tank and contains unsaturated polyester resin pellets as solid fuels. Experimental results show that O/F in the primary combustion chamber is independent of oxygen flow rate if the residence time is long enoug..., 一般社団法人日本機械学会, 日本語 - 401 TOPAZ を用いた木星探査衛星の概念設計
加藤 健太郎, 永田 晴紀, 戸谷 剛, 工藤 勲, 北海道支部講演会講演概要集, 2002, 42, 104, 105, 2002年10月05日
The purpose of this conceptual study is to design a Jupiter probe for investigating the origin of Great Red Spot which has continued to exist on the planet for more than 300 years. The probe is equipped with TOPAZ, Russian nuclear reactor which is used for a power source for ion thrusters which must shorten the interplanetary time of flight from an orbit whose radius is the sphere of influence of the Earth to Jovian orbit and for one for mission equipments observing the Great Red where solar power is quite faint., 一般社団法人日本機械学会, 日本語 - 404 ハイブリッドロケット回収系用飛翔体の開発 : 有翼飛翔体打ち上げ装置の開発
芝 邦明, 下岡 彩子, 松尾 亮弘, 難波江 亮, 豊田 国昭, 佐鳥 新, 三橋 龍一, 永田 晴紀, 北海道支部講演会講演概要集, 2002, 42, 110, 111, 2002年10月05日
The launching project of a hybrid rocket from Taiki Multi-purpose Aerospace Park has been conducted by the research groups in the universities in Hokkaido. In the plan, after the flight in outer space, the rocket is taken back using a parafoil-glider with of GPS (Grobal Positioning System). The present study deals with the development of the winged vehicle to obtain the basic materials of the parafoil-glider system., 一般社団法人日本機械学会, 日本語 - Interactive combustion of two-dimensionally arranged quasi-droplet clusters under microgravity
H Nagata, Kudo, I, K Ito, S Nakamura, Y Takeshita, COMBUSTION AND FLAME, 129, 4, 392, 400, 2002年06月
o investigate the mutual Interactions between droplets in the spray combustion, combustion of 2-dimensionally arranged quasi-droplet clusters is studied under microgravity. Quasi-droplet samples, which are solid in room temperature and change into liquid just after the ignition, consist of alcohol (propanol, butanol, pentanol, or hexanol) and polyethylene glycol with a volumetric ratio of 2:1. Seven samples sustained by glass rods form a 2-dimensional quasi-droplet cluster. Electrically heated nichrome wires ignite all samples in the cluster simultaneously. Single envelope flames that surround the clusters appeared, The results show that the sample spacing has a strong effect on the shape and movement of the flame. Sample clusters with large sample spacings come to the external group combustion through the scavenging combustion mode, whereas the small spacing clusters start directly with the external group combustion. At large sample spacings, the distance from the edge of the sample cluster to the flame (flame distance) increases to a maximum value and then decreases with time. The period of flame growth is prolonged with decreasing sample spacing and finally, at a small enough sample spacing, the flame distance keeps increasing until the flame disappears. This flame movement is attributed to the fuel vapor accumulation effect, which becomes more dominant with decreasing sample spacing, The burning lifetime decreases monotonically and approaches the value of the single flame with increasing sample spacing. The flame distance decreases monotonically and approaches the single flame radius with increasing sample spacing also. These results render Important confirmations of the external group combustion phenomena and prove the importance of the two kinds of unsteadiness, that is, the scavenging combustion with large droplet interval and the fuel vapor accumulation effect with small droplet interval, in group combustion. (C) 2002 by The Combustion Institute., ELSEVIER SCIENCE INC, 英語 - ハイブリッドロケットのための回収型有翼飛しょう体実験
三橋龍一, 佐鳥新, 中村直紀, 芝邦明, 下岡彩子, 大滝誠一, 豊田国昭, 永田晴紀, 宇宙科学技術連合講演会講演集, 46th, Pt.3, 2002年 - Performance of droplet generator and droplet collector in liquid droplet radiator under microgravity
T Totani, M Itami, H Nagata, Kudo, I, A Iwasaki, S Hosokawa, MICROGRAVITY SCIENCE AND TECHNOLOGY, 13, 2, 42, 45, 2002年
The Liquid Droplet Radiator (LDR) has an advantage over comparable conventional radiators in terms of the rejected heat power-weight ratio. Therefore, the LDR has attracted attention as an advanced radiator for high-power space systems that will be prerequisite for large space structures. The performance of the LDR under microgravity condition has been studied from the viewpoint of operational space use of the LDR in the future. In this study, the performances of a droplet generator and a droplet collector in the LDR are investigated using drop shafts in Japan: MGLAB and JAMIC. As a result, it is considered that (I) the droplet generator can produce uniform droplet streams in the droplet diameter range from 200 to 280 [mum] and the spacing range from 400 to 950 [mum] under microgravity condition, (2) the droplet collector with the incidence angle of 35 degrees can prevent a uniform droplet stream, in which droplet diameter is 250 [mum] and the velocity is 16 [m/s], from splashing under microgravity condition, whereas splashes may occur at the surface of the droplet collector in the event that a nonuniform droplet stream collides against it., Z A R M TECHNIK PUBLISHING DIV, 英語 - A new era of the hybrid rocket
Ryojiro Akiba, Takashi Nakajima, Harunori Nagata, Advances in the Astronautical Sciences, 110, 325, 329, 2002年01月01日
New type of hybrid rocket is proposed. Keywords are safety, environmentally tender and low cost. Staged combustion hybrid rocket was designed and basic characteristics were measured and its applicability was confirmed to the near future use, for example to fully reusable sounding rockets, space tugs etc. - Opposed-flow flame spread in a circular duct of a solid fuel: Influence of channel height on spread rate
Nozomu Hashimoto, Satoshi Watanabe, Harunori Nagata, Tsuyoshi Totani, Isao Kudo, Proceedings of the Combustion Institute, 29, 1, 245, 250, 2002年
The influence of channel height on flame spread in a circular duct of the solid fuel in an opposed-flow configuration was examined. Polymethylmethacrylate cylinders with a circular duct (diameter of 1, 2, or 3 mm) were used as fuel specimens, and both flame-spreading and stabilized combustion were observed. In the case of stabilized combustion, the flame cannot spread into the duct because of the high oxygen velocity. The flame-traveling velocity is the velocity at which the flame widens the ductby fuel consumption. Therefore, the flame-traveling velocity in stabilized combustion is significantly low compared with flame-spreading combustion. In the case of flame-spreading combustion, the equivalence velocity, which contains channel height information, defines whether the regime is the thermal or the chemical regime. When the equivalent velocity is higher than a certain value, the flame-spread rate is controlled by chemical effects. On the whole, the flame-spread rate decreases with the decrease of channel height in the case of flame-spreading combustion because of the curvature effect. Owing to the curvature effect, the area ratio of the flame to that of the solid surface decreases with decreasing channel height, and this is conspicuous when the channel height is low. The curvature effect is negligible when the channel height is sufficiently large compared with the flame stand-off distance., Elsevier Ltd, 英語 - Performance of droplet emittor for liquid droplet radiator under microgravity
Tsuyoshi Totani, Masahiro Itami, Shigeru Yabuta, Harunori Nagata, Isao Kudo, Akira Iwasaki, Shunsuke Hosokawa, Nippon Kikai Gakkai Ronbunshu, B Hen/Transactions of the Japan Society of Mechanical Engineers, Part B, 68, 668, 1166, 1173, 2002年
The Liquid Droplet Radiator (LDR) has an advantage over comparable conventional radiators in terms of the rejected heat power-weight ratio. Therefore, the LDR has attracted as an advanced radiator for high-power space systems that will be prerequisite for large space structures. In this study, the performance of a droplet emittor under microgravity condition has been investigated from the viewpoint of operational space use of the LDR in the future. From experiments, it is considered that the droplet emittor can produce uniform droplet streams under microgravity condition in the non-dimensional wave number range from 0.215 to 0.490. In this range, the droplet diameter and the spacing range are from 204 to 285 [μm] and from 445 to 1160 [μm] respectively. And it is concluded that these diameter and spacing can be estimated by the equations based on the law of conservation of mass in the process of generating droplets., Japan Society of Mechanical Engineers, 日本語 - Performance test under microgravity on a centrifugal droplet collector for liquid droplet radiator
Tsuyoshi Totani, Masahiro Itami, Shigeru Yabuta, Harunori Nagata, Isao Kudo, Akira Iwasaki, Shunsuke Hosokawa, Nippon Kikai Gakkai Ronbunshu, B Hen/Transactions of the Japan Society of Mechanical Engineers, Part B, 68, 674, 2780, 2787, 2002年
The Liquid Droplet Radiator (LDR) has an advantage over conventional radiators in terms of the rejected heat power-weight ratio. LDR has been taken notice as an advanced radiator for high-power generation systems which will be prerequisite for large space structures. In this study, the performance of a centrifugal droplet collector under microgravity condition has been investigated from the viewpoint of operational space use of LDR in the future. It has been concluded that (1) a centrifugal collector is able to transport working fluid to a recirculating pump under microgravity condition
(2) the ability to pump working fluid is formulated as the sum of pressure head and velocity head generated in the centrifugal collector where the velocity is c (0 <
c ≤ 1) times as fast as in the rigid rotational flow
(3) splashing of the working fluid occurs at that position, when working fluid strikes against part of the entrance of the pitot tube on the centrifugal collector., Japan Society of Mechanical Engineers, 日本語 - 3330 ハイブリッドロケット回収系用飛翔体の開発
芝 邦明, 下岡 彩子, 豊田 国昭, 大滝 誠一, 佐鳥 新, 三橋 龍一, 永田 晴紀, 年次大会講演論文集, 2002, 0, 339, 340, 2002年
The launching project of a hybrid rocket from Taiki Multi-purpose Aerospace Park has been conducted by the research stuffs in the universities in Hokkaido. In this plan, after the flight in outer space, the rocket is taken back using a parafoil-glider with of GPS (Grobal Positioning System). The present study deals with the development of the winged vehicle to obtain the basic materials of the parafoil-glider system., 一般社団法人 日本機械学会, 日本語 - 3332 2 段燃焼方式ハイブリッドロケットに関する基礎研究 : 酸化剤として亜酸化窒素を用いた場合
栗田 慎一郎, 豊田 国昭, 青木 嘉範, 秋葉 鐐二郎, 藤井 篤之, 永田 晴紀, 年次大会講演論文集, 2002, 0, 343, 344, 2002年
The basic experiments have been conducted to develop the staged combustion hybrid-rocket engine with two combustion chambers. In the present study, in order to understand the characteristics of the oxidizer (N_2O), the dependency on temperature was measured and the self-pressurizing supply system was tested. Moreover the combustion test was carried out. The results reveal that N_2O is useful as oxidizer of the hybrid-rocket engine., 一般社団法人 日本機械学会, 日本語 - 3333 噴流衝突型高推力ハイブリッドロケットの開発および打上げ実験
永田 晴紀, 戸谷 剛, 工藤 勲, 伊藤 献一, 大和田 陽一, 中山 久広, 渡辺 三樹生, 佐鳥 新, 高田 毅, 芝 邦明, 豊田 国昭, 中須賀 真一, 宮村 典秀, 年次大会講演論文集, 2002, 0, 345, 346, 2002年
A joint research team of universities has been organized to develop small-scale reusable launch systems based on new type of hybrid rockets. This paper describes the project outline, development study of a new type of hybrid rocket engine, and the result of a ballistic test launch using this engine. The key design of the new type hybrid rocket engine, designated as Jet Impinging type Hybrid Rocket, is that the gas flow collides with the solid fuel surface to accelerate the heat transfer to the fuel, resulting in improved thrust level. Static firing tests with an engineering model engine with a LOX cooling system showed sufficiently prompt ignition and stable combustion. Based on these results, a flight model engine was developed to conduct a ballistic test launch. The engine worked excellently in the test launch and the result was successful, showing that te engine showed expected performance in the flight condition., 一般社団法人 日本機械学会, 日本語 - 1502 微小重力場を利用した高温雰囲気における固体燃料粒子群の点火に関する研究
永田 晴紀, 藤井 篤之, 伊藤 献一, 戸谷 剛, 工藤 勲, 年次大会講演論文集, 2002, 0, 3, 4, 2002年
Ignition delays of fuel particle clusters that are inserted in a high-temperature environment are measured by changing the ambient temperature and the interval between fuel particles. To eliminate the effect of natural convection on the phenomena, ignition experiments are conducted under microgravity conditions. A qualitative discussion using numerical results about the ignition of a spherical cluster of fuel particles is made also to interpret the experimental results. Main conclusions obtained are in the followings. Ignition delay of a fuel particle cluster quickly immersed into a hot environment has a minimum at a certain interval between particles, and the minimum ignition delay is shorter than that of a single particle. This ignition delay behavior is common to previous results reported for droplet array ignition. The cause of this behavior is the long characteristic reaction time compared to the characteristic fuel mass transfer rate. Depending on the ambient temperature, the following two phenomena occur : (1) The ignition delay becomes minimum for a larger particle interval with increasing the ambient temperature because the volatilization time, which decreases with increasing the particle interval, is more dominant with increasing the ambient temperature. For higher ambient temperatures, the ignition delay is expected to decrease monotonically with increasing the particle interval because the reaction time should be negligible small comparing with the volatilization time. (2) An envelope flame appears for a larger particle interval with decreasing the ambient temperature because more amount of fuel gas is accumulated around the cluster at ignition for lower ambient temperatures., 一般社団法人 日本機械学会, 日本語 - 307 端面燃焼式ハイブリッドロケット用多孔質固体燃料の後退特性
渡辺 賢, 橋本 望, 永田 晴紀, 戸谷 剛, 工藤 勲, 北海道支部講演会講演概要集, 2002, 0, 84, 85, 2002年
Because of some defects, such as low combustion efficiency and O/F shift, Hybrid Rocket Motors have not been practicable yet. To overcome these defects of conventional Hybrid Rocket Motors, End-Burning Hybrid Rocket Motor has been suggested. In this rocket motor, oxidizer gas is injected from one side of a porous solid fuel grain and combustion occurs on the other side. If the oxidizer flow velocity in the gaps of the porous solid fuel is sufficiently high, the flame cannot spread into these gaps. This type of combustion is called Stabilized Combustion. Understanding of fuel regression characteristics is important to develop the End-Burning Hybrid Rocket Motor. In this paper, the fuel regression characteristics of porous solid fuels are investigated experimentally, and feasible arrangement of the gaps in the porous solid fuel for End Burning Hybrid Rocket Motor is shown., 一般社団法人 日本機械学会, 日本語 - Hydrogen concentration measurements of supersonic hydrogen-air shear layer using catalytic reaction
Takakage Arai, Jiro Kasahara, Fuminori Sakima, Junji Miura, Takayuki Ami, Harunori Nagata, 10th AIAA/NAL-NASDA-ISAS International Space Planes and Hypersonic Systems and Technologies Conference, 2001年12月01日
To investigate development of an air-hydrogen supersonic shear layer and distribution of hydrogen concentration, a hydrogen jet was injected into a cold air supersonic free-streem in a paralell direction. The free stream Mach number was about 1.81. Using a catalytic reaction on a thin platinum wire, heat release due to catalytic reaction, a heat transfer coefficient and hydrogen concentration were measured. It was shown that the paralell injection was found to affect on mixing condition. The effect of paralell injection on hydrogen concentration profile was clarified. It seemed that there was the stoichiometric condition at the outer edge of shear layer. It was confirmed that the diffusion of Hydrogen, including turbulent mixing, had an effect of flow configuration. © 2001 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. - Reusable launch vehicle concepts based on hybrid rockets
Kazuhide Mizobataj, Shimon Narita, Jan Nakaya, Hiroshi Yoshida, Takakage Arai, Jiro Kasahara, Harunori Nagata, Ken'ichi Ito, Ryojiro Akiba, Yooichi Oowada, 10th AIAA/NAL-NASDA-ISAS International Space Planes and Hypersonic Systems and Technologies Conference, 2001年12月01日
Hybrid rocket motors, propelled by a combination of a solid fuel and a liquid oxidizer, are best suited to development of small launch systems in university laboratories, because of their advantages such as safety, easy handling, and low costs. The performance of hybrid rocket motors of three classes of thrust - 10tonf, 1tonf, and 200kgf - is estimated. The feasibility of reuseable launch systems based on the three motors is preliminarily analysed for suborbital microgravity experiments. A system with a lOton-thrust-class motor by a coolant bleed cycle with polystyrene and LOx fed by an LE-5B turbopump will be capable of exposing a payload of 360kg to a microgravity environment for three minutes. It is also predicted that a system with a lton-thrust-class motor will be moderately capable and that with a 200kg-thrust-class motor will not be feasible for microgravity missions, mainly because the weight of its helium pressurization system for feeding LOx will spoil its mass ratio and takeoff/climbing performance. © 2001 by Kazuhide Mizobata. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission. - 再使用型有翼ハイブリットロケットシステムの基本設計と飛行性能予測
溝端 一秀, 中谷 淳, 成田 志門, 吉田 拓史, 杉山 弘, 永田 晴紀, 伊藤 献一, 秋葉 鐐二郎, 大和田 陽一, 室蘭工業大学紀要, 51, 0, 91, 97, 2001年11月30日
Hybrid rocket motors, propelled by a combination of a solid fuel and a liquid oxidizer,are best suited to development of small launch systems in university laboratories,becauee af their advantageous characteristics such as Bafety, easy handling. and lawcosts. The performance of hybrid rocket motors of three classes of thrust - 10tonf,1tonf, and 200kgf - is estimated. The feasibility of reusable winged launch systemsbased on the three motors is prelirninarily analyzed for suborbital micro-gravityexperiments. The results tell that a sysbem with a 10tonf-class motor by a coolant bleedcycle wit..., 室蘭工業大学, 日本語 - 713 ハイブリッドロケット回収系用飛翔体の開発 : 有翼飛翔体制御のアビオニクスについて(OS7-5 ハイブリッドロケット(システム,回収系))(OS7 北海道における航空宇宙工学研究の進展と未来)
高田 強, 豊田 国昭, 大滝 誠一, 三橋 龍一, 佐鳥 新, 永田 晴紀, 青木 嘉範, 北海道支部講演会講演概要集, 2001, 41, 254, 255, 2001年09月25日
The launching project of a hybrid rocket from Taiki Multi-purpose Aerospace Park as an Experimental base is developing by the research stuffs in the Universities in Hokkaido. In this plan, after the flight in outer space, rocket is taken back using a Parafoil-glider Making use of GPS (Global Positioning System). To testify the system, an experiment is Conducted by using Radio-controlled model airplanes equipped GPS system., 一般社団法人日本機械学会, 日本語 - 707 微小重力環境下における液滴ラジエータの作動流体循環試験(OS7-2 マイクログラビティー)(OS7 北海道における航空宇宙工学研究の進展と未来)
薮田 茂, 宮本 拓哉, 戸谷 剛, 永田 晴紀, 工藤 勲, 岩崎 晃, 細川 俊介, 北海道支部講演会講演概要集, 2001, 41, 238, 239, 2001年09月25日
Liquid Droplet Radiator (LDR) is an advanced and light weight radiator for high power space systems that will be prerequisite for large space structures superseding the traditional heat-pipe radiator. LDR consists of 3 elements that are a droplet generator, a droplet collector and a gear pump. The results of performance tests of 3 elements conducted respectively under microgravity confirm that these devices can function properly under weightless condition. This paper describes performance tests on circulation of working fluid in LDR conducted under microgravity and supplementary experiments..., 一般社団法人日本機械学会, 日本語 - 709 衝撃波後方の非定常場における水素 : 空気混合気の白金触媒燃焼発熱量の測定(OS7-3 計測・地上実験設備)(OS7 北海道における航空宇宙工学研究の進展と未来)
中村 大輔, 永田 晴紀, 戸谷 剛, 工藤 勲, 北海道支部講演会講演概要集, 2001, 41, 242, 243, 2001年09月25日
Catalytic combustion is a candidate for an apparatus to stabilize and promote combustion for a Scramjet engine that is a strong candidate for thrusters of space planes and supersonic airplane in the next generation. However, the mechanism of catalytic combustion in the supersonic flow has hardly been clarified because there are little cases of the research on catalytic combustion in the supersonic flow, especially at the unsteady Field. In this research, we tried to measure catalytic heat release rate at the unsteady flow behind a shock wave using Ni or Pt prove and a shock tube. The temper..., 一般社団法人日本機械学会, 日本語 - 712 再使用型ハイブリッドロケットシステム基本設計と飛行性能予測(OS7-5 ハイブリッドロケット(システム,回収系))(OS7 北海道における航空宇宙工学研究の進展と未来)
吉田 拓史, 中谷 淳, 前田 直樹, 溝端 一秀, 杉山 弘, 永田 晴紀, 伊藤 献一, 秋葉 鐐二郎, 大和田 陽一, 北海道支部講演会講演概要集, 2001, 41, 252, 253, 2001年09月25日
Hybrid rocket motors, propelled by a combination of a solid fuel and a liquid oxidizer, are best suited to development of small launch systems in university laboratories, because of their advantageous characteristics such as safety, easy handling, and low costs. The performance of hybrid rocket motors of three classes of thrust - 10tonf, 1tonf, and 200kgf- is estimated. The feasibility of reusable winged launch systems based on the three motors is preliminarily analyzed for suborbital micro-gravity experiments. The results tell that a system with a 10tonf-class motor by a coolant bleed cycl..., 一般社団法人日本機械学会, 日本語 - 714 ハイブリッド・ロケット回収系用飛翔体の開発 : パラシュートと有翼飛翔体(OS7-5 ハイブリッドロケット(システム,回収系))(OS7 北海道における航空宇宙工学研究の進展と未来)
芝 邦明, 下岡 彩子, 豊田 国昭, 大滝 誠一, 佐鳥 新, 青木 嘉範, 永田 晴紀, 北海道支部講演会講演概要集, 2001, 41, 256, 257, 2001年09月25日
The launching project of a hybrid rocket from Taiki Multi-purpose Aerospace Park as an experimental base is conducted by the research stufls in the universities in Hokkaido. In this plan, after the flight in outer space, the rocket is taken back using a parafoil-glider making use of GPS(Grobal Positioning System). The present study deals with the development of the parachute and the winged vehicle to obtain the basic materials of the parafoil-glider system., 一般社団法人日本機械学会, 日本語 - 715 ハイブリッド・ロケットの回収系用飛翔体の開発 : GPSによる無線誘導実験(OS7-5 ハイブリッドロケット(システム,回収系))(OS7 北海道における航空宇宙工学研究の進展と未来)
大滝 誠一, 豊田 国昭, 芝 邦明, 三橋 龍一, 今井 規晶, 佐々木 大輔, 佐鳥 新, 青木 嘉範, 永田 晴紀, 北海道支部講演会講演概要集, 2001, 41, 258, 259, 2001年09月25日
The launching project of a hybrid rocket from Taiki Multi-purpose Aerospace Park as an experimental base is conducted by the research stufls in the universities in Hokkaido. In this plan, after the flight in outer space, the rocket is taken back using a parafoil-glider making use of GPS (Grobal Positioning System). To testify the system, an experiment is conducted by using Radio-controlled model airplanes equipped GPS system., 一般社団法人日本機械学会, 日本語 - 719 二段燃焼式ハイブリッドロケットの基礎研究(OS7-6 航空宇宙推進(流体,燃焼))(OS7 北海道における航空宇宙工学研究の進展と未来)
青木 嘉範, 藤井 篤之, 永田 晴紀, 加勇田 清勇, 栗田 慎一郎, 豊田 国昭, 秋葉 鐐二郎, 杉木 光輝, 北海道支部講演会講演概要集, 2001, 41, 266, 267, 2001年09月25日
The staged combustion hybrid rocket is under development by our research group since 1999. This hybrid rocket engine consists of two combustion chambers. The primary combustion chamber is the lower part of the very fuel tank itself filled with granular solid fuels. The fuel rich gas generated by the first stage combustion flows into the secondary combustion chamber, which is located in the bottom core part of the primary combustion chamber. The additional oxidizer is injected to the secondary combustion chamber in order to attain an optimal specific impulse by completing combustion. This ne..., 一般社団法人日本機械学会, 日本語 - 720 衝突噴流式ハイブリッドロケットのフライトモデル設計のための地上燃焼試験(OS7-6 航空宇宙推進(流体,燃焼))(OS7 北海道における航空宇宙工学研究の進展と未来)
中山 久広, 渡辺 三樹生, 永田 晴紀, 工藤 勲, 戸谷 剛, 北海道支部講演会講演概要集, 2001, 41, 268, 269, 2001年09月25日
The authors have proposed a new fuel configuration to overcome defects of conventional hybrid rockets such as low thrust level and low combustion efficiency. This new fuel configuration allows mixing and combustion to occur around jet-impinging points on forward ends of solid fuels. In the previous researches with cylindrical PMMA blocks as fuel, LOx was employed as oxidizer and ignition and combustion characteristics are investigated. As the result, good ignition characteristics and steady combustion with LOx were confirmed. In the present research, the engineering model of jet-impinging h..., 一般社団法人日本機械学会, 日本語 - 熱工学 : 機械工学年鑑(2000年)
西尾 茂文, 長坂 雄次, 鈴木 雄二, 笹口 健吾, 山田 雅彦, 長崎 孝夫, 太田 淳一, 円山 重直, 松島 均, 前川 透, 大中 逸雄, 伊藤 献一, 小林 成嘉, 長谷 耕志, 小住 敏之, 山下 卓也, 神原 信志, 永田 晴紀, 日本機械学會論文集. B編, 67, 660, 1881, 1890, 2001年08月25日
一般社団法人日本機械学会, 日本語 - B-2-9 ハイブリッドロケット回収系のための飛翔体無線誘導実験
三橋 龍一, 今井 規晶, 佐々木 大輔, 芝 邦明, 大滝 誠一, 佐鳥 新, 青木 嘉範, 永田 晴紀, 電子情報通信学会総合大会講演論文集, 2001, 1, 239, 2001年03月07日
一般社団法人電子情報通信学会, 日本語 - 北海道の教育現場における宇宙開発
伊藤献一, 永田晴紀, 佐鳥新, 溝端一秀, 日本航空宇宙学会年会講演会講演集, 32nd, 2001年 - 反射板による爆轟波の再点火機構に関する研究
永田晴紀, 平成12年度衝撃波シンポジウム講演論文集, 383, 386, 2001年 - Hydrogen concentration of 2-D supersonic air-hydrogen mixing layer using platinum catalytic reaction
Takakage Arai, Jiro Kasahara, Junji Miura, Fuminori Sakima, Harunori Nagata, Nihon Kikai Gakkai Ronbunshu, B Hen/Transactions of the Japan Society of Mechanical Engineers, Part B, 67, 656, 934, 939, 2001年
Abstrcat To investigate development of an air-hydrogen supersonic shear layer and distribution of hydrogen concentration, a hydrogen jet was injected into a cold air supersonic free-stream in a paralell direction. The free stream Mach number was 1.81. Using a catalytic reaction on a platinum wire, heat release due to catalytic reaction, a heat transfer coefficient and hydrogen concentration were measured. It was shown that paralell injection was found to affect on mixing condition. The effect of paralell injection on hydrogen concentration profile was clarified. It seemed that there was the stoichiometric condition at the outer edge of shear layer. It was confirmed that the diffusion of Hydrogen, including turbulent mixing, had an effect of flow configuration., Japan Society of Mechanical Engineers, 日本語 - 506 Virtual Space Research Laboratoryの構築と設計支援への適用(北海道における宇宙活動I)(オーガナイズドセッション(d)北海道における宇宙活動)
吉川 茂雄, 戸谷 剛, 永田 晴紀, 工藤 勲, 北海道支部講演会講演概要集, 2000, 40, 192, 193, 2000年09月25日
Engineering changes have been managed by Configuration Management Provision regulated by such a procurement agency as NASDA or USEF. Recently activities of configuration control board which should be established when ECP is issued seem to be neglected because of complicated procedures and pressure of business of CCB members. For revitalizing the activities, we are developing a virtual environment which supports configuration management, especially engineering changes, using functions of Virtual Space Research Laboratory (VSRL) we have been operating these one year and a half., 一般社団法人日本機械学会, 日本語 - 510 衝突噴流式ハイブリッドロケットの点火特性および燃焼安定性(北海道における宇宙活動II)(オーガナイズドセッション(d)北海道における宇宙活動)
渡辺 三樹生, 中山 久広, 永田 晴紀, 戸谷 剛, 工藤 勲, 大和田 陽一, 北海道支部講演会講演概要集, 2000, 40, 200, 201, 2000年09月25日
The authors have proposed a new fuel configuration to overcome defects of conventional hybrid rockets such as low thrust level and low combustion efficiency. This new fuel configuration allows mixing and combustion to occur around jet-impinging points on forward ends of solid fuels. Previous researches with cylindrical PMMA blocks as fuel and gas oxygen as oxidizer revealed that the regression rates of forward ends increase due to the effect of impinging jet. In the present research, LOX was employed as oxidizer and ignition and combustion characteristics are investigated. Additionally, we ..., 一般社団法人日本機械学会, 日本語 - 512 再使用型ハイブリットロケット : システム概要と有翼機体の成立可能性(北海道における宇宙活動II)(オーガナイズドセッション(d)北海道における宇宙活動)
成田 志門, 溝端 一秀, 杉山 弘, 吉田 拓史, 永田 晴紀, 伊藤 献一, 秋葉 鐐二郎, 大和田 陽一, 北海道支部講演会講演概要集, 2000, 40, 204, 205, 2000年09月25日
Hybrid Rocket systems, propelled by a combination of a solid propellant and liquid oxidizer, have significant advantages such as safety, easy handling and low costs. A joint team has been organized by Hokkaido University, Muroran Institute of Technology, Hokkaido Institute of Technology, Tokai University, Tokyo Metropolitan Institute of Technology, National Space Development Agency, Japan, and some private sectors, so as to investigate the feasibility of suborbital/orbital reuseable hybrid rocket systems. A preliminary analysis of the rocket motor performance and winged flight trajectories ..., 一般社団法人日本機械学会, 日本語 - 514 超音速流用水素濃度プローブの応答速度(北海道における宇宙活動III)(オーガナイズドセッション(d)北海道における宇宙活動)
笹木 正裕, 中村 大輔, 永田 晴紀, 戸谷 剛, 工藤 勲, 北海道支部講演会講演概要集, 2000, 40, 208, 209, 2000年09月25日
The authors have proposed a new simple method using catalytic reaction on platinum wire to evaluate hydrogen concentrations in a hydrogen-air supersonic flow without the need for costly apparatus. In this research, in order to verify availability of the probe at turbulent mixing field, we attempt to evaluate the response time of the probe in a shock tube. A shock tube generates a shock wave. The response time means a rising time of supplied power when a surface discontinuity of shock wave reaches the probe. Also, we try to examine the relation between response time and hot wire temperature., 一般社団法人日本機械学会, 日本語 - 516 液滴ラジエータ要素の微小重力実験(北海道における宇宙活動III)(オーガナイズドセッション(d)北海道における宇宙活動)
薮田 茂, 戸谷 剛, 永田 晴紀, 工藤 勲, 伊丹 雅洋, 岩崎 晃, 細川 俊介, 北海道支部講演会講演概要集, 2000, 40, 212, 213, 2000年09月25日
Liquid Droplet Radiator (LDR) is an advanced radiator for high power space systems that will be prerequisite for large space structures (LSS). LDR consists of 3 elements, droplet generator, droplet collector and gear pump. Many studies on performance of LDR have been done at the ground. But LDR must work in space, under weightless condition. Only few performance tests of LDR have been made under microgravity. We have made performance tests of LDR's 2 elements which are droplet generator and droplet collector under microgravity, and have got a certain results. In this year, we will test abou..., 一般社団法人日本機械学会, 日本語 - 517 微小重力場における模擬石炭粒子群の点火に関する研究(北海道における宇宙活動III)(オーガナイズドセッション(d)北海道における宇宙活動)
丹羽 由樹子, 永田 晴紀, 戸谷 剛, 工藤 勲, 北海道支部講演会講演概要集, 2000, 40, 214, 215, 2000年09月25日
The purpose of this study is to investigate mutual influences of fine coal particles in a coal dust cloud on the ignition process. Ignition processes of a cluster of four coal particles in high temperature atmosphere are observed. Spherical active carbons impregnated with salicylic acid are employed as fuel particles. Four fuel particles are arranged on tops of a regular tetrahedron. To remove the influence of the natural convention, experiments are made under microgravity. As a result, a bright luminous flame is observed with a small interval of fuel particles. This is because the density ..., 一般社団法人日本機械学会, 日本語 - A NEW ERA OF THE HYBRID ROCKET
AKIBA Ryojiro, NAKAJIMA Takashi, NAGATA Harunori, The Journal of space technology and science : a publication of Japanese Rocket Society, 16, 2, 1, 5, 2000年09月01日, [査読有り], [国際誌]
英語 - Simulation for deployment of an inflatable disk in orbit
K Takahashi, H Nagata, Kudo, I, JOURNAL OF SPACECRAFT AND ROCKETS, 37, 5, 707, 708, 2000年09月
Two deployment methods of an inflatable tube and a disk model are compared. The first method is carried out under conditions in which a test specimen is placed on the floor. It is found that the inexpensive and simple deployment of an inflatable structure on the floor has a substantial flaw because a portion of the disk interacted with the floor. At the second simulation, in which the test specimen is placed in the air, data obtained at both the dropshaft test and 1-g ground test before a microgravity test showed satisfactory simulation., AMER INST AERONAUT ASTRONAUT, 英語 - B110 EFFECTS OF COAL PARTICLES SEEDING ON BURNING VELOCITY OF METHANE-AIR MIXTURES(Droplet/particle combustion-2) :
NAGATA Harunori, FUJITA Osamu, ITO Kenichi, KUDO Isao, TAKESHITA Yasuhiro, Proceedings of the ... JSME-KSME Thermal Engineering Conference, 1, "1, 151"-"1-155", 2000年
This paper describes experimental and numerical studies about effects of coal particles seeding on flame propagation velocity and burning velocity of lean methane-air mixtures. To produce uniform coal dust clouds and to eliminate the buoyancy effect, experiments are carried out under microgravity environment, being obtained with the 500-m drop shaft of Japan Microgravity Center (JAMIC). Experimental and numerical results reveal that seeding of an appropriate amount of coal particles increases flame propagation velocity. In spite of this, the increase of burning velocity is almost zero or very small. Accordingly, the main cause of the enhancement of flame propagation velocity is the gas expansion due to coal combustion behind the methane-air flame. This is because coal combustion is nearly isolated from methane combustion in conditions investigated in this study., 日本機械学会, 英語 - Evaluation of mass transfer coefficient and hydrogen concentration in supersonic flow by using catalytic reaction
H Nagata, M Sasaki, T Arai, T Totani, Kudo, I, PROCEEDINGS OF THE COMBUSTION INSTITUTE, 28, 713, 719, 2000年
The authors propose a new simple method to evaluate hydrogen concentrations in a hydrogen/air supersonic mixing layer without the need for costly apparatus. Catalytic reaction occurs on an electrically heated platinum wire in the supersonic flow of a hydrogen/air mixture. By adopting the technique of a constant-temperature hot-wire anemometer, the heat transfer coefficient and the catalytic heat release rate are measures. A series of experiments with different platinum wire temperatures shows that the platinum wire temperature does not affect the catalytic heat release rate, implying that the rate of mass transfer from the flow to the platinum wire surface is the controlling factor. This means that the catalytic heat release rate gives the mass transfer coefficient of the controlling species, which is hydrogen/oxygen in lean/rich mixtures. It is found that the effect of hydrogen concentration on the ratio of heat and mass transfer coefficients is very weak, suggesting that the mass transfer coefficient is obtained with reasonable accuracy from the heat transfer coefficient by assuming the equivalent spatial distributions of heat and mass transfer. Based on this result, a method to translate the catalytic heat rate into the hydrogen concentration of the flow is proposed. To prove the accuracy of this method, hydrogen concentrations of hydrogen/air premixed supersonic flows were measured successfully. Finally, one example applying this method to an actual supersonic mixing layer is presented., ELSEVIER SCIENCE INC, 英語 - 511 端面燃焼式ハイブリッドロケットの燃焼安定性に関する研究(北海道における宇宙活動II)(オーガナイズドセッション(d)北海道における宇宙活動)
橋本 望, 永田 晴紀, 戸谷 剛, 工藤 勲, 北海道支部講演会講演概要集, 2000, 0, 202, 203, 2000年
To overcome defects of conventional hybrid rockets such as low combustion efficiency and the O/F shift during the combustion, the authors have proposed a new form of hybrid rocket fuel. The fuel is a fibrous bed in which oxidizer gas flows. Stable diffusion flame appears at the exit surface. Previous researches show that sudden increase of the fuel regression rate occurs with the increase of ambient pressure. This sudden increase is attributed to the flame spreading between fuel fibers. To clarify the limit of fuel gap space the diffusion flame can spread into, experimental study was made. Critical gap space, which means the minimum gap space the diffusion flame can spread into, was obtained experimentally as a function of oxygen gas flow velocity and ambient pressure. Using this result, necessary conditions to realize a stable combustion with this new fuel form are shown., 一般社団法人 日本機械学会, 日本語 - Kazuhide MIZOBATA, Harunori NAGATA, Kenichi ITO, Ryojiro AKIBA, Isao KUBOTA, "A Reusable Hybrid Rocket System: Concept Outlines and Feasibility of Winged Flights," Proceedings of the 22nd International Symposium on Space Technology and Science, Vol.2, ・・・
2000年
Kazuhide MIZOBATA, Harunori NAGATA, Kenichi ITO, Ryojiro AKIBA, Isao KUBOTA, "A Reusable Hybrid Rocket System: Concept Outlines and Feasibility of Winged Flights," Proceedings of the 22nd International Symposium on Space Technology and Science, Vol.2, pp.1334-1340, 2000. - Harunori NAGATA, Mikio WATANABE, Takashi SANDA, Shin SATORI, Yoshinori AOKI, Isao KUDO, Ryojiro AKIBA, Isao KUBOTA, "Improvement of Fuel Regression Rate by Impinging Jet for High Thrust Hybrid Rocket Motors," Proceedings of the 22nd International Sympo・・・
2000年
Harunori NAGATA, Mikio WATANABE, Takashi SANDA, Shin SATORI, Yoshinori AOKI, Isao KUDO, Ryojiro AKIBA, Isao KUBOTA, "Improvement of Fuel Regression Rate by Impinging Jet for High Thrust Hybrid Rocket Motors," Proceedings of the 22nd International Symposium on Space Technology and Science, Vol.1, pp.121-126, 2000. - Shin SATORI, Harunori NAGATA, Yoshinori AOKI, Ryojiro AKIBA, Isao KUDO, Isao KUBOTA, "Preliminary Study for Staged Combustion Hybrid Rocket," Proceedings of the 22nd International Symposium on Space Technology and Science, Vol.1, pp.116-120, 2000.
2000年 - Yoshinori AOKI, Shin SATORI, Kunihiko TAHARA, Hayato MAEDA, Takehisa MURATA, Harunori NAGATA, Ryojiro AKIBA, Isao KUBOTA, "Measuring of the Hybrid Rocket Combustion by Spectrum Mesurement," Proceedings of the 22nd International Symposium on Space Techn・・・
2000年
Yoshinori AOKI, Shin SATORI, Kunihiko TAHARA, Hayato MAEDA, Takehisa MURATA, Harunori NAGATA, Ryojiro AKIBA, Isao KUBOTA, "Measuring of the Hybrid Rocket Combustion by Spectrum Mesurement," Proceedings of the 22nd International Symposium on Space Technology and Science, Vol.2, pp.127-132, 2000. - 円筒形固体燃料での拡散火炎の燃え広がり限界直径に関する研究
加藤隆博, 橋本望, 永田晴紀, 工藤勲, 日本機械学会北海道支部講演会講演概要集, 39th, 132-133, 1999年09月25日
日本語 - Evaluation of supersonic turbulent mixing using catalytic combustion of constant temperature Pt wire
T Arai, H Nagata, A Endo, H Sugiyama, S Morita, H Hosokawa, JSME INTERNATIONAL JOURNAL SERIES B-FLUIDS AND THERMAL ENGINEERING, 42, 1, 65, 70, 1999年02月
Supersonic combustion using catalytic wire at constant temperature in a cold supersonic flow field was investigated in a square duct with a backward-facing Step. The free stream Mach number was M(m) = 1.81. Hydrogen was injected transversely behind a backward-facing step into a cold air free-stream. The heat released from the catalytic combustion had no effect on the temperature of the catalyst. This indicates that the reaction rate of the catalytic combustion observed in this study was determined by the concentration of H(2) and/or O(2) on the surface of the catalyst. The spatial distribution of heat released from the catalytic combustion in supersonic turbulent mixing layer, corresponds to the spatial distribution of concentration of H(2) and/or O(2) in local, was obtained. It was found that the most suitable position for supersonic combustion was at the outer edge of the mixing layer., JAPAN SOC MECHANICAL ENGINEERS, 英語 - 触媒反応を利用した超音速流中水素濃度プローブに関する研究
永田晴紀, 第9回ラム/スクラムジェットシンポジウム議論集, 239, 244, 1999年 - Experimental investigation of inclined hydrogen injection into a supersonic flow
Takakage Arai, Hideo Fukuzoe, Junji Miura, Harunori Nagata, Hiroshi Hosokawa, 9th International Space Planes and Hypersonic Systems and Technologies Conference, 1, 7, 1999年01月01日
© 1999 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Effect of the injection angle on supersonic mixing was conducted. The pressure loss, penetration height and hydrogen concentration were measured. It was shown that, as the injection angle decreased, pressure loss decrease. However, in the far field, injection atigle was found to have little effect in the penetration height. The hydrogen concentration was introduced by using catalytic reaction on Ptwire, and discussed quantitatively. It was clarified that the longitudinal vortex took an important part on mixing process. It was also clarified that the new technique proposed in the present paper was very useful and easy to evaluate the hydrogen concentration in supersonic mixing flow field. - Evaluation of hydrogen concentration of supersonic flow using catalytic reaction on Pt wire
Harunori Nagata, Yasuhiko Sakai, Moto-Omi Iwatsuki, Syuji Morita, Isao Kudo, Nihon Kikai Gakkai Ronbunshu, B Hen/Transactions of the Japan Society of Mechanical Engineers, Part B, 65, 636, 2666, 2671, 1999年
The authors propose a new simple method which can be used to evaluate hydrogen concentration in hydrogen-air supersonic mixing layers without the need for costly apparatus. The catalytic reaction occurs on an electrically heated platinum wire in hydrogen-air supersoic mixing layers. By- adapting the technique of constant temperature type hotwire anemometers, a catalytic heat release rate is measured. A series of experiments with different Pt wire temperatures shows that Pt wire temperature has little effect on the catalytic heat release rate, implying that the rate of transfer of molecules to the Pt wire surface is the controlling factor. Accordingly, the heat release rate is related to the hydrogen concentration in the flow. The profile of hydrogen concentration is obtained by assuming the equivalent spatial distribution of heat and mass transfer. Stoichiometric conditions are found to be realized in the mixing layer., Japan Society of Mechanical Engineers, 日本語 - Evaluation of supersonic turbulent mixing using catalytic combustion of constant temperature pt wire
Takakage Arm, Harunori Nagata, Akira Endo, Hiromu Sugiyama, Shuji Morita, Hiroshi Hosokawa, Nihon Kikai Gakkai Ronbunshu, B Hen/Transactions of the Japan Society of Mechanical Engineers, Part B, 64, 619, 793, 799, 1998年12月01日
Supersonic combustion using catalytic wire at constant temperature in a cold supersonic flow field was investigated in a square duct with a backward-facing step. The free stream Mach number was of M m = 1.81. Hydrogen was injected transversely behind a backward-facing step into a cold air free stream. The heat release due to the catalytic combustion has no effect of the temperature of catalyst. It indicates that the reaction rate of the catalytic combustion observed in this study was determined by the consentration of H 2 and/or O 2 on the surface of the catalyst. The spatial distribution of heat release due to the catalytic combustion in supersonic turbulent mixing layer, that corresponds to the spatial distribution of consentration of H 2 and or/0 2 in local, was obtained. It was found that there exists the most suitable position for the supersonic combustion at the outer edge of mixing layer., 一般社団法人 日本機械学会, 日本語 - 反射板によるデトネーション波の伝播促進に関する研究
永田晴紀, 平成10年度衝撃波シンポジウム, 1998年 - H
2 concentration profile in cold supersonic hydrogen-air mixing layer*
Takakage Arai, Shuji Mortta, Harunori Nagata, Hiromu Sugiyama, 8th AIAA International Space Planes and Hypersonic Systems and Technologies Conference, 645, 650, 1998年01月01日
© 1998, by the American Institute of Aeronautics and Astronautics, Inc. All Rights Reserved. Hydrogen was injected normally to a Mach 1.8 cold air stream with a backward-facing step to investigate mixing flow field in a scramjet combustor. Catalytic reaction on constant temperature Pt wire was used to measure the mixing condition of H 2 and O 2 . It was new technique to investigate H 2 mixing condition. The amount of heat release due to the catalytic reaction, which corresponds to the concentration of H 2 and/or O 2 on the surface of the catalyst, was measured spatially so that the local mixing condition of H 2 and O 2 was cleared. The results showed that there were two core regions, which would have good mixing condition for supersonic combustion, in the mixing layer. The development of core regions along flow direction was also clarified. - 短形断面模擬燃焼器におけるスロット噴射による超音速燃焼に関する研究(共著)
富岡 定毅, 田口 秀之, 永田 晴紀, 高橋 周平, 氏家 康成, 河野 通方, 日本航空宇宙学会誌, 42, 483, 243, 250, 1994年
Supersonic combustion in a rectangular duct was investigated experimentally. Hydrogen fuel was injected parallel to an air flow from a two-dimensional slot injector. Two-dimensional flame was established downstream of the injector and so-called precombustion shock was observed. The shock wave originated at separation region in air flow boundary layer upstream of the injector. The relationship between the strength and the position of the shock wave was different from that in perpendicular injection case. In slot injection case, it is supposed that the separation region and reacting region interacted to act as a large wedge and generate the rather strong shock. Combustion without generation of the precombustion shock was achieved with conversion duct. However, the combustion efficiency was low and the ignition delay distance was large in this case. The shock wave was expected to enhance the reaction by rising the static temperature and pressure., 日本航空宇宙学会, 日本語 - メタン-水素混合燃料を用いた超音速燃焼に関する研究(共著)
田口 秀之, 富岡 定毅, 永田 晴紀, 高橋 周平, 氏家 康成, 河野 通方, 日本航空宇宙学会誌, 42, 483, 224, 231, 1994年
In order to attain combustion of methane fuel in supersonic airstreams of relatively low temperature, hydrogen fuel was added to the fuel as a ignition promotor. In experiments, rectangular combustor with backward step was adopted and the methane-hydrogen mixture fuel was injected into supersonic airstreams. The region where the self-ignition occurs was confirmed by high speed direct photographs, and self-ignition characteristics were investigated with parameters of equivalence ratio and fraction of hydrogen in the fuel. The injection form was refined with use of the obtained results to attain self-ignition with low fraction of hydrogen. The combustion of methane-hydrogen mixture fuel was compared with that of hydrogen fuel by using wall static pressure distributions and OH/CH emission images of the flame. As results, it is confirmed that the combustion of methane fuel in the airstreams is obtained by addition of hydrogen fuel, and it is considered that the self-ignition in refined injection form is obtained under same conditions as that of hydrogen fuel, independently of the fraction of hydrogen in total injected fuels. The static pressure distribution and the flame shape in methane-hydrogen mixture fuel case were similar to those in hydrogen fuel case., 一般社団法人 日本航空宇宙学会, 日本語 - An experimental and numerical investigation on the hot surface ignition of premixed gases under microgravity conditions
H. Nagata, H. M. Kim, J. Sato, M. Kono, Symposium (International) on Combustion, 25, 1, 1719, 1725, 1994年
There have been many studies of hot surface ignition of premixed gases. However, using the experimentalresults obtained up to this time, it is difficult to understand the basic mechanism of ignition by heated surfaces because these experiments were made under the normal gravity condition, suffering from the effect of natural convection. In the present study, experiments were made on the ignition of mixtures by heated nickel, tungsten,and platinum wires under normal gravity and microgravity conditions. With a methane-oxygen mixture ignited by a heated tungsten wire, a strong gravity effect on the ignition delay is observed. On the contrary, there is almost no gravity effect observed with a methane-air mixture. Results of calculations show that ignition experiments under the microgravity condition are simulated successfully by this numerical model. It is shown from the numerical results that the ignition point in the methane-oxygen case is apart from the hot-wire surface, while, in the methane-air case, this point exists near the hot-wire surface. Therefore, natural convection has little effect on the ignition delay of a methane-air mixture. Experimental results of hydrogen-air mixtures ignited by heated platinum wires show that chemicalspecies are supplied to the platinum wire surface by natural convection, and therefore, the surface reaction is promoted under the normal gravity condition. Experimental and numerical results show that the surface reaction is the most energetic when the equivalence ratio of the mixture is 0.3, which conflicts with the result obtained by Coward and Guest that the rate of catalytic reaction attains a maximum at stoichiometric composition. This contradiction may be due partially to the gravity effect in their experiments. © 1994 Combustion Institute., 英語 - 微小重力環境を利用した高温白金線による水素-空気予混合気の点火に関する研究
永田晴紀, 第31回燃焼シンポジウム講演集, 319, 321, 1993年
書籍等出版物
講演・口頭発表等
- 冷却によるグラファイトノズルの浸食抑制効果
吉丸 利,永田 晴紀,伊藤 聖司,Landon T. Kamps
第57回燃焼シンポジウム講演論文集, 2019年11月
2019年11月 - 2019年11月 - アブレータ材料によるハイブリッドロケットノズル浸食抑制に関する研究
奥田 晃崇,Landon T. Kamps,櫻井 和人, 井上 卓,Tor Viscor,内山 絵里香,池田 華子, 吉丸 利,脇田 督司,永田 晴紀
第57回燃焼シンポジウム講演論文集, 2019年11月
2019年11月 - 2019年11月 - Using equivalent burn time to improve regression characterization of the CAMUI type hybrid rocket engine
T. Viscor, H. Nagata
53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017, 2017年01月01日
2017年01月01日 - 2017年01月01日, Burn time errors caused by various start-up transient effects have a large effect on the regression simulation model of the Cascaded Multi Impinging Jet hybrid rocket engine. This paper analyses these burn time errors and their effect on the regression simulations for short burn time engines. To address these the equivalent burn time is then defined as the time the engine would burn if it was burning at steady state level throughout the burn time to achieve the measured total impulse. The accuracy of the regression simulation with and without the use of equivalent burn time are then finally compared. Equivalent burn time alone without at the same time addressing other errors is found to be inadequate to clearly address the burn time errors. - Numerical and experimental investigation of nozzle thermochemical erosion in hybrid rockets
Daniele Bianchi, Landon Kamps, Francesco Nasuti, Harunori Nagata
53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017, 2017年01月01日
2017年01月01日 - 2017年01月01日, Despite some inherent disadvantages, hybrid rockets are today considered as having a great potential to become one of the future generation of propulsion systems, partly due to recent emphasis on propulsion safety, reliability, low development cost, reduced environmental pollution, and greater operability. Nevertheless, the hybrid rocket development has not achieved the same level of maturity as solid and liquid traditional systems. An aspect that has not been much dealt with in the open literature is that of nozzle erosion, whose minimization or reduction is one of the challenges in hybrid rocket propulsion. To this goal, a joint numerical and experimental investigation of nozzle throat erosion has been performed using a computational fluid dynamics approach compared to static firing tests carried out on a 2kN-class lab-scale hybrid rocket burning liquid oxygen and highdensity polyethylene. The numerical approach is able to capture the main features of the nozzle throat erosion behavior, fairly reproducing the throat erosion rate values and its dependence upon the oxidizer to fuel mixture ratio and motor chamber pressure. - Investigation of regression characteristics under relatively high-pressure in Axial-injection End-Burning Hybrid Rockets
Yuji Saito, Masaya Kimino, Tsuji Ayumu, Yushi Okutani, Kazunobu Omura, Hiroyuki Yasukochi, Kentaro Soeda, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata
53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017, 2017年01月01日
2017年01月01日 - 2017年01月01日, This study is an investigation of Axial-Inection End-Burning Hybrid Rockets aimed at revealing fuel regression characteristics under relatively high pressure conditions. Firing tests were conducted using gaseous oxygen as oxidizer at chamber pressures and oxidizer port velocities ranging from 0.22 MPa to 1.05 MPa and 31 m/s to 103 m/s, respectively. The results of fifteen static firing tests show that fuel regression rate increases as the chamber pressure increases, and regression rates ranged from approximately 1.1 mm/s at 0.25 MPa to 5.4 mm/s at 0.71 MPa. Furthermore, it is observed that the fuel regression rate is not influenced by oxidizer port velocity. The athours encountered a problem refered to as backfiring in this paper, and developed a calculation model to investigate this problemanalytically. The calculation model explains why the back-firing problem tends to occur in relatively high-pressure conditions, and leads to the conclusion that increasing nozzle throat diameter is an effective means of preventing back-firing from occuring. - Investigation of graphite nozzle-throat-erosion in a laboratory-scale hybrid rocket using GOX and HDPE
Landon Kamps, Shota Hirai, Yassine Ahmimache, Raymond Guan, Harunori Nagata
53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017, 2017年01月01日
2017年01月01日 - 2017年01月01日, The authors of this paper employ a recently developed reconstruction technique titled nozzle-throat reconstruction technique to investigate graphite nozzle-throat-erosion in a laboratory-scale hybrid rocket motor using gaseous oxygen as an oxidizer and high density polyethylene as a fuel. Fifteen static firing tests were conducted under varying experimental conditions to confirm the validity of the reconstruction technique results, and to collect a wide range of nozzle-throat-erosion data. Furthermore, a technique for carrying out classical finite difference calculations for 1D convective and conductive heating based on the time histories of gas properties as determined by the reconstruction technique is introduced and used to estimate nozzle throat wall temperature history and convective heat transfer coefficient history. Results show a distinct trend where nozzle erosion rates increase in the beginning of a firing test, and subsequently decrease for the remainder of the firing test, even though nozzle throat temperatures continue to increase. It is shown that the decreasing nozzle-throat-erosion rates coincide with decreasing mass fluxes and that the erosion rates in this regime may be sensitive to oxidizer to fuel mass ratio. - 小型衛星放出機構から放出される50kg級衛星の熱設計
PANES Mitchao Delburg, 戸谷剛, 脇田督司, 永田晴紀
宇宙科学技術連合講演会講演集(CD-ROM), 2017年, 日本語
2017年 - 2017年 - 2‐amino‐2‐methyl‐1,3‐propanediolの固相‐固相結晶転移による潜熱を利用した蓄熱材の開発 過冷却状態の結晶化による発熱とガラス転移点の関係
後藤凌平, 永田晴紀, 戸谷剛, 脇田督司
日本伝熱シンポジウム講演論文集(CD-ROM), 2017年, 日本語
2017年 - 2017年 - ふく射センサの低温環境下でのふく射率および単一液滴流からの排熱量
嶋田泰三, 高梨知広, 両門健人, 戸谷剛, 永田晴紀, 脇田督司
日本伝熱シンポジウム講演論文集(CD-ROM), 2017年, 日本語
2017年 - 2017年 - 酸化剤にN2Oを用いたハイブリッドロケットにおける燃焼室圧力と特性排気速度効率の関係について
清谷優理香, 山口亮, 櫻井和人, KAMPS Landon, 井上卓, 脇田督司, 戸谷剛, 永田晴紀
宇宙科学技術連合講演会講演集(CD-ROM), 2017年, 日本語
2017年 - 2017年 - 液体酸素を用いた端面燃焼式ハイブリッドロケットの実現可能性
津地歩, 齋藤勇士, 尾村和信, 君野正弥, 戸谷剛, 脇田督司, 永田晴紀
宇宙科学技術連合講演会講演集(CD-ROM), 2017年, 日本語
2017年 - 2017年 - 気体亜酸化窒素を酸化剤とした端面燃焼式ハイブリッドロケットの燃料後退特性
尾村和信, 津地歩, 齋藤勇士, 君野正弥, 奥谷勇士, 小水弘大, 戸谷剛, 脇田督司, 永田晴紀
宇宙科学技術連合講演会講演集(CD-ROM), 2017年, 日本語
2017年 - 2017年 - 端面燃焼式ハイブリッドロケットの推力制御時におけるヒステリシス特性に関する研究
君野正弥, 齋藤勇士, 津地歩, 尾村和信, 安河内裕之, 添田建太郎, 戸谷剛, 脇田督司, 永田晴紀
宇宙科学技術連合講演会講演集(CD-ROM), 2017年, 日本語
2017年 - 2017年 - デトネーション波の衝突・反射による遷移促進効果に関する研究
大関敦, 松岡将司, 桧物恒太郎, 脇田督司, 戸谷剛, 永田晴紀
燃焼シンポジウム講演論文集, 2016年11月26日, 日本語
2016年11月26日 - 2016年11月26日 - 燃料微小管内を燃え拡がる火炎への雰囲気圧力の影響
横井俊希, 齋藤勇士, 尾村和信, 津地歩, 安河内裕之, 添田建太郎, 脇田督司, 戸谷剛, 永田晴紀
燃焼シンポジウム講演論文集, 2016年11月26日, 日本語
2016年11月26日 - 2016年11月26日 - 端面燃焼式ハイブリッドロケットの燃料後退モデルに関する考察
齋藤勇士, 横井俊希, 津地歩, 尾村和信, 安河内裕之, 添田建太郎, 戸谷剛, 脇田督司, 永田晴紀
燃焼シンポジウム講演論文集, 2016年11月26日, 日本語
2016年11月26日 - 2016年11月26日 - 液体酸素を用いた固体燃料管内燃え拡がりに関する研究
津地歩, 齋藤勇士, 横井俊希, 尾村和信, 嶋田泰三, 戸谷剛, 脇田督司, 永田晴紀
燃焼シンポジウム講演論文集, 2016年11月26日, 日本語
2016年11月26日 - 2016年11月26日 - ハイブリッドロケットにおけるノズルスロートエロージョン抑制材料に関する研究
山口亮, 川端良輔, 川端良輔, 平井翔大, KAMPS Landon, 脇田督司, 戸谷剛, 永田晴紀
日本機械学会年次大会講演論文集(CD-ROM), 2016年09月10日, 日本語
2016年09月10日 - 2016年09月10日 - 端面燃焼式ハイブリッドロケットの推力制御特性
永田晴紀, 齋藤勇士, 横井俊希, 嶋田泰三, 安河内裕之, 添田建太郎, 戸谷剛, 脇田督司
日本機械学会年次大会講演論文集(CD-ROM), 2016年09月10日, 日本語
2016年09月10日 - 2016年09月10日 - 超小型衛星の短期開発を目指す熱設計法の妥当性の検証
戸谷 剛, 毛利 正宏, 脇田 督司, 永田 晴紀
宇宙科学技術連合講演会講演集, 2016年09月06日, 日本航空宇宙学会, 日本語
2016年09月06日 - 2016年09月06日 - F117 結晶変化を伴う蓄熱材の蓄熱・放熱試験
戸谷 剛, 國 拓也, 佐藤 敏文, 脇田 督司, 永田 晴紀
熱工学コンファレンス講演論文集, 2014年11月08日, 日本語
It is desirable that the heat storage materials for nano and micro satellites have the characteristic of not phase change but crystal transformation for the heat storage. Trans-1,4-polybutadien has the desirable characteristic of crystal transformation at the heat storage temperature. The heat storage and release tests were conducted on the ground and in orbit. It is clarified that the heat storage and release characteristics do not change in heat storage and release cycle of 300 times and Trans-1,4-polybutadien can storage and release heat in orbit. - E215 金属膜を持つ微小キャビティの構造と放射ピークの関係
戸谷 剛, 色川 俊雄, 脇田 督司, 永田 晴紀
熱工学コンファレンス講演論文集, 2014年11月08日, 日本語
Micro-cavities were made from Zr dispersing liquid on SiO_2 bases and were coated by Au for generating resonance phenomena. The theoretical wavelength of reflectance drop differed 4% from the experiment data. A numerical analysis taking into account the pitch was carried out to find causes of the difference. The wavelength of reflectance drop in the analysis were closer to experiment data than theoretical predictions. On the other hand, the wavelengths of reflectance drop except for the fundamental vibrational mode obtained in the analysis did not accord 7.6 to 8.5 % with experimental data. The reflectance obtained in the analysis was 4 to 11% different from experimental data. - 高レイノルズ数域におけるCAMUI型固体燃料の後退特性におよぼすスケールの影響
石山達也, 稲場康彦, 寺川健, 遠藤瞳, 永田晴紀, 戸谷剛, 脇田督司
宇宙科学技術連合講演会講演集(CD-ROM), 2013年, 日本語 - 缶サット・超小型衛星を用いた創造的科学技術人育成ネットワークの構築
宮崎康行, 永田晴紀, 木村真一, 佐原宏典, 坂本啓, 川島レイ, 安藤恵美子
宇宙科学技術連合講演会講演集(CD-ROM), 2013年, 日本語 - 環状爆轟波から平面爆轟波への遷移に流路形状が及ぼす影響
菊地敬太, 桧物恒太郎, 脇田督司, 戸谷剛, 永田晴紀
流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム講演集(CD-ROM), 2013年, 日本語 - CAMUI型ハイブリッドロケットの作動履歴におけるスケール則の構築
稲場康彦, 石山達也, 金井竜一朗, 脇田督司, 戸谷剛, 永田晴紀
宇宙科学技術連合講演会講演集(CD-ROM), 2013年, 日本語 - 微小重力環境を利用した固体燃焼現象研究(H23研究班WG報告)
藤田修, 中村祐二, 永田晴紀, 菊池政雄, 伊藤昭彦, 鳥飼宏之, 梅村章, 高橋周平, 池田光優, CHUNG Suk Ho, OLSON Sandra L
宇宙利用シンポジウム, 2012年03月, 日本語 - サブオービタル機開発に向けての飛行実験と今後の計画
米本浩一, 相良慎一, 松本剛明, 永田晴紀, 越智徳昌, 石本真二, 麥谷高志, 牧野隆, 木元健一
飛行機シンポジウム講演集(CD−ROM), 2012年, 日本語 - 大口径パルスデトネーションエンジン用イニシエータにおける円筒デトネーション波の伝播に関する研究
桧物恒太郎, 棧敷和弥, 脇田督司, 戸谷剛, 永田晴紀
流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム講演集(CD-ROM), 2012年, 日本語
共同研究・競争的資金等の研究課題
- 液体酸素を用いる端面燃焼式ハイブリッドロケットの実証研究
科学研究費助成事業
2022年04月01日 - 2026年03月31日
永田 晴紀, 添田 建太郎, KAMPS LANDON, 脇田 督司
日本学術振興会, 基盤研究(A), 北海道大学, 22H00240 - 亜酸化窒素を酸化剤とするハイブリッドロケット用再生冷却グラファイトノズルの研究
科学研究費助成事業
2023年03月08日 - 2025年03月31日
永田 晴紀, GALLO GIUSEPPE
液体亜酸化窒素を酸化剤とするハイブリッドロケットエンジンにおける再生冷却システムの信頼性と実現可能性を実験的に調べた。昨年度は極低温の液体酸素を用いて再生冷却ノズルの良好な作動を実験的に実証したが、その結果を更に発展させたものである。本年度の研究の新規性は、ロケットノズルで発生する高い熱流束を適切に管理するための冷却剤として、自身の蒸気圧により加圧供給される液体亜酸化窒素を使用することにある。具体的には、燃焼中にグラファイトノズルを目標温度(1500 K程度)以下に保つことにより、ノズル壁面での浸食を完全に抑止することを目指す。
実験においては、直径が異なるオリフィスを通過させることで液体亜酸化窒素にキャビテーションを発生させることで質量流量を適切に制御し、10回の燃焼試験を実施した。燃焼中に、酸化剤流路の複数個所で冷却剤(亜酸化窒素)の圧力と温度の履歴を取得した。グラファイトノズルに埋め込んだ熱電対により、ノズル壁面からグラファイトノズルへの熱流束の履歴も取得した。
燃焼実験においてグラファイトノズル内部の温度は安定しており、燃焼室圧力が5~30 barの範囲において700~1200 Kの範囲内であった。円筒一次元の定常熱伝導を仮定すると、ノズル壁面の温度は1500 Kを充分に下回っていると予想された。実験条件は過去の研究においてはノズル浸食が発生した条件であるが、全ての実験においてノズル浸食は全く観測されなかった。
キャビテーションオリフィスと冷却流路に沿って温度と圧力の分布を取得し、流路に沿って亜酸化窒素の相変化と熱伝達特性の変化の関係を明らかにした。液体酸素と比較して液体亜酸化窒素は高い冷却能力を示した。最後に、キャビテーションオリフィスの位置がシステム全体の流量特性に及ぼす影響を明らかにした。
日本学術振興会, 特別研究員奨励費, 北海道大学, 22KF0013 - 革新的超小型衛星による機動的で高頻度な深宇宙探査領域の開拓
科学研究費助成事業 学術変革領域研究(B)
2020年10月02日 - 2023年03月31日
船瀬 龍, 永田 晴紀, 尾崎 直哉
本研究領域のため総括班として,2021年度は,引き続き研究コミュニティへの研究成果の発信などの広報活動を行いながら,複数の計画研究から構成される研究領域全体の研究を円滑に進行させた.
本研究領域が最終的に目指すキックモーターを有する超小型探査機システムの構築を見据えて,3つの計画研究(通称,推進班,バス班,軌道班)は,システム全体として物理的,機械的,時間的な境界条件を満たしながら全体としての設計を最適化する必要がある.総括班として,各種オンラインコミュニケーションツールを活用して頻度高く研究会合を開催し,各班の検討の整合性をとりながらタイムリーに各班の進む方向性にフィードバックをかけて,領域全体の研究を進めることができた.
2021年11月には,宇宙科学技術連合講演会にてオーガナイズドセッション「超小型探査機を用いた月以遠深宇宙探査」を開催し,超小型衛星による深宇宙探査を目指す関連コミュニティへ本研究領域の成果発信を実施するとともに,超小型探査機による深宇宙探査に必要な技術シーズや技術実証計画に関する発表,超小型探査機で実施したい科学ミッションのニーズに関する発表も実施され,ニーズ側とシーズ側との間で活発な意見交換が行われた.
また,2022年2月には,本研究領域の成果を活用した具体的な深宇宙ミッションを計画しているGEO-X研究グループの研究会に招待され本研究領域の活動内容等を発表し,利用側のユーザーコミュニティとのコミュニケーションも実施した.その他,各種研究会での招待講演等も通じて,本研究領域の目指す世界の認知向上とコミュニティの拡大に向けて大きな成果をあげることができた.
日本学術振興会, 学術変革領域研究(B), 国立研究開発法人宇宙航空研究開発機構, 20H05746 - 小型宇宙機に革新的軌道変換能力を与えるハイブリッドキックモータの開発
科学研究費助成事業 学術変革領域研究(B)
2020年10月02日 - 2023年03月31日
永田 晴紀, KAMPS LANDON, 脇田 督司
静止トランスファー軌道まで相乗りし、近地点で1 km/s前後の増速を与えて月、火星、金星等へ向かう超小型深宇宙探査機用のキックモータ(軌道変換用上段ロケット)をハイブリッドロケットにより実現するため、高比推力(適切な燃料/酸化剤流量比)を維持する固体燃料形状設計の自在性、ノズルスロート浸食履歴を予測した上での最適ノズル形状設計、およびノズルスロート浸食を抑制する作動条件およびノズル材料、を得るのが本研究の目的である。R3年度はプリフライトモデルモータを開発して高空燃焼実験装置を用いた地上燃焼実験を実施したのが主な実施項目である。概要は以下に説明する通りである。
・高空燃焼実験装置とプリフライトモデルモータとのインタフェース構造を製作した。
・プリフライトモデルモータを製作し、高空燃焼実験装置を用いて0.1気圧程度の低圧力環境雰囲気で点火実験を行った。推力が良好に立ち上がり、真空の宇宙環境での点火シーケンス構築の見通しを得た。併せて、導電性プラスチックを用いて点火する新たな手法を考案し、高い信頼性で点火可能である事を確認した。
・高空燃焼実験装置を用いて10秒間の燃焼実験を実施し、ロケット噴流により低圧力環境が維持されることを確認した。
・プリフライトモデルモータによる60秒の燃焼実験を実施し、設計要求を満たす推力履歴が得られることを確認した。
・亜酸化窒素の蒸気圧を利用したコールドガススラスタを製作し、高空燃焼実験装置を用いて低圧力環境下で推力を計測した結果、設計通りの推力が得られていることを確認した。
日本学術振興会, 学術変革領域研究(B), 北海道大学, 20H05747 - ハイブリッドロケットノズル浸食の機構解明
科学研究費助成事業 基盤研究(B)
2019年04月01日 - 2022年03月31日
永田 晴紀, 脇田 督司
本研究は、提案者らが過去の研究で開発に成功した、ハイブリッドロケットのノズルスロート面積と燃料流量の各時間履歴を同時に取得する手法であるNozzle Throat Reconstruction Technique(以下、NTRT)を用いて、ノズル浸食(ノズルスロートが浸食により拡大し、ノズル膨張比が下がる現象)の物理/化学機構を解明し、浸食速度予測手法を構築することを目指すものである。R1年度は以下の内容を実施した。
1. 酸化剤の種類がノズル浸食に及ぼす影響:ガス酸素および亜酸化窒素(N2O)を酸化剤として地上燃焼実験を実施した。推力、燃焼室圧力、および酸化剤流量の各履歴を取得し、NTRTを用いてOF比およびノズルスロート面積の各履歴を得た。モータは200 N級を用い、燃料には高密度ポリエチレンを使用した。
2.モータスケールがノズル浸食に及ぼす影響:スケールの影響は、主に酸素モータにより取得する。R1年度は1スケール(200 N級)のみでデータを取得した。
3.共同研究の発展的展開:研究協力者であるローマ大学のD. Bianchiらとの共同研究により、ノズル浸食の予測を可能とする技術を構築する。R1年度は8/19-22にインディアナポリスで開催されたPropulsion and Energy Forumの会場で打合せを行い、今後の研究方針を決定した。並行して、ノズル浸食を普遍的に評価できる無次元数を見出し、ハイブリッドロケットにおけるノズル浸食速度の定式化に成功した。
4.ノズル浸食を抑制する技術:ノズル浸食に強い材料の候補として、IHIエアロスペース社からC/Cノズルの提供を受け、浸食特性を評価した。
日本学術振興会, 基盤研究(B), 北海道大学, 19H02336 - 液体酸素-固体燃料の拡散燃焼機構の解明と端面燃焼式ハイブリッドロケットへの適用
科学研究費助成事業 基盤研究(C)
2019年04月01日 - 2022年03月31日
脇田 督司, 永田 晴紀
本研究は、2015年度採択の挑戦的萌芽研究、および2017年度採択の基盤研究(C)おいて、ガス酸素供給で実証されてきた「端面燃焼式ハイブリッドロケット」を液体酸化剤供給でロケットとして成立させることを目指し、本年度は下記の3つについて研究を進めてきた。
1. 燃料後退速度(燃料が燃焼によって後退してく速度)の圧力依存性:ミリオーダの比較的大きな単一のポートを有するアクリル燃料を用いた液体酸素供給の燃焼を実現し、燃焼室の圧力に比例して燃料後退速度が増加することを明らかにし、酸化剤流速と燃料圧力の関数として燃料後退速度の実験式を構築した。これはガス酸素供給の燃焼特性と一致し、液体酸素を用いた場合でも高圧な燃焼条件下で高推力なロケットを実現可能なことを示した。
2. 複数の微小ポートを有する燃料を用いたロケット燃焼の予備検討:光造形3Dプリンタによって造形されたサブミリオーダの複数のポートを有する燃料に、液体酸素を供給した場合の燃焼実験を実施した。しかしながら、液体酸素の極低温による熱応力と3Dプリンタ造形特有の燃料の脆さのために、液体酸素による燃料の予冷段階から燃料に亀裂が発生し破損してしまうことが明らかになった。この問題を解決するために、次年度以降は熱応力が発生しない常温の推進剤である液体亜酸化窒素などを用いた燃焼の実現を目指す。
3. 圧力依存性の効果を考慮した燃料後退モデルの構築:燃料表面から火炎までの距離(消炎距離)に圧力の関数である化学反応速度を導入することで、圧力によって燃料への熱入力量が変化するようなモデルを構築し、圧力に比例した燃料後退速度の増加を再現できた。しかしながら、消炎距離の計算に用いられる火炎先端上流の未燃の燃料酸化剤混合ガスの流速のモデル化次第で、大きく結果が変わってくるため、次年度以降は数値計算による流速のオーダの見積もりとガス流速のモデル化を行う。
日本学術振興会, 基盤研究(C), 北海道大学, 19K04832 - 端面燃焼式ハイ ブリッドロケットの実用化研究
科学研究費助成事業 基盤研究(C)
2017年04月01日 - 2020年03月31日
添田 建太郎, 永田 晴紀, 田丸 博晴
本研究は,高い燃焼効率や優れたスロットリング特性等の多くの利点を有し,H27-28 年度の挑戦的萌芽研究において本提案者らによって初めて実証された「端面燃焼式ハイブリッドロケット」について,以下の成果が得られた。1) 3Dプリンタによる燃料造形において,高い生産性で,高精度,大型化が可能な造形条件を見出した,2) 様々な推力履歴でスロットリング燃焼実験を行い,推力応答特性を明らかにした,3) 酸化剤の種類を変えた燃焼実験および燃焼モデルの構築により,燃焼機構を明らかにした.
日本学術振興会, 基盤研究(C), 東京大学, 17K06943 - 超小型火星探査機用Ne計測装置の基礎開発
科学研究費助成事業 基盤研究(A)
2017年04月01日 - 2020年03月31日
杉田 精司, 笠原 慧, 永田 晴紀, 吉岡 和夫, 黒澤 耕介, 黒川 宏之, 三浦 弥生
火星大気に含まれるNe同位体比測定では、同じ質量電価比をもつ“Ar二価イオン”が妨害成分となり、測定精度が出ない。本研究では、分子種により透過特性の異なる分離膜を用いて、質量分析部に導入する前にNeとArを分離し、上記の問題を解決することを目指している。膜材候補として、バイトンとポリイミドを選択し透過特性を評価した。その結果、0.125mm厚のポリイミド膜を用いた分離膜に地球大気(~10^5Pa)を透過させると、20Ne/40Ar比を10^-3(地球大気の値)から10^2まで向上できることが判明した。1mm厚のバイトン膜では20Ne/40Ar比が3桁向上した。これらの結果が、膜内の分子拡散理論で説明できることも確認した。火星大気組成を仮定した透過の理論計算と質量分析計内で作られるAr++/Ar+比(約0.1)とから、0.1mm厚のポリイミド膜でNe測定が可能なことがわかった。これにより、火星大気進化を知る上で重要な20Ne/22Ne比を精度5~10%で測定できる見通しが立った。
また、火星探査に必要な宇宙空間航行期間を想定して、放射線によるポリイミド膜の透過特性劣化度を調べた。50 kradのガンマ線を照射する前後でポリイミド膜に対するNe、Arの透過量に劣化は見られず、十分な放射線耐性を有していることが確認できた。
さらに、NASAの火星探査機搭載装置での使用実績もあるSAES社製のST175を準備して、材質からの脱ガス、標準大気を用いてのガス精製効率(精製時間や精製結果が必要充分であるか等)、固定方法や耐久性、などについて試験や評価を進めた。ゲッター材に7Aで数10分以上電流を流して脱ガスを行うと要求ブランクに近くなること、0.3 Pa程度の地球大気を精製するにはゲッター材に電流を流す必要なないことなどがわかった。
日本学術振興会, 基盤研究(A), 東京大学, 17H01175 - ハイブリッドロケットノズル浸食の高精度時系列データ取得と機構解明
科学研究費助成事業 基盤研究(B)
2015年04月01日 - 2018年03月31日
永田 晴紀, 戸谷 剛, 脇田 督司
ハイブリッドロケット燃焼実験データの新しい解析手法として、ノズルスロート再現法を開発し、燃焼中の燃料-酸化剤比とノズルスロート面積の各履歴を同時に取得することに成功した。推力2 kN級モータを用いて、燃料-酸化剤比を幅広く変えて地上燃焼実験を行い、本再現法を適用した。当量比0.4から1.4の範囲内で、ノズル浸食速度と当量比の関係が得られ、過去に化学反応律速を仮定して予測された数値計算結果と一致する実験結果を示した。また、耐酸化性を念頭にノズル浸食抑制材料を選定し、地上燃焼実験により抑制効果を評価した。結果、グラファイトの表面をSiCでコーティングしたノズルが最も高い浸食抑制効果を示した。
日本学術振興会, 基盤研究(B), 北海道大学, 研究代表者, 競争的資金, 15H04197 - スペースプレーン技術の極超音速飛行実証システムの開発研究
科学研究費助成事業 基盤研究(A)
2013年05月31日 - 2018年03月31日
澤井 秀次郎, 坂井 真一郎, 坂東 信尚, 丸 祐介, 永田 晴紀, 後藤 健, 小林 弘明, 吉光 徹雄
空気吸込式エンジンを用いるスペースプレーンの実現に向けて,飛行実証を通して基盤となる技術を獲得するために,気球による高高度からの落下と小型ロケットブースターによる加速を組み合わせた高速飛行実証システムの構築を目指した.飛行軌道検討を主としたシステム概念検討を行った.その結果を踏まえ,飛行実験機の試作研究を行った.試作研究を通して,システム統合および飛行制御系技術の実践研究を行った.さらに,スペースプレーンに必要な技術として,空力設計技術の研究を行った.実験オペレーションまで想定した実験計画を検討し,本システムのメリットに加え,課題も整理した.
日本学術振興会, 基盤研究(A), 国立研究開発法人宇宙航空研究開発機構, 25249125 - 光造形技術を利用した端面燃焼式ハイブリッドロケットの実現と性能実証
科学研究費助成事業 挑戦的萌芽研究
2015年04月01日 - 2017年03月31日
永田 晴紀
著者らは、軸方向に多数の微小ポートを有する固体燃料のポート出口端面で微小拡散火炎群を保持する「端面燃焼式ハイブリッドロケット」を提案してきたが、燃料の製作が困難なため実証実験を見送って来た。近年の3Dプリンタの発展により複雑な燃料形状が製作可能となり、世界で初めて端面燃焼式ハイブリッドロケットの実証実験を実施した。燃焼実験の結果、初期の非定常期間を経て、燃焼中に燃料-酸化剤比が一定に保たれる定常燃焼への移行が確認された。ポート内径が0.2~0.5 mmの単ポート燃料試料を用いた燃焼実験も実施した。燃え広がり燃焼と安定燃焼の2つのモードが観察され、両モードを分ける臨界摩擦速度が確認された。
日本学術振興会, 挑戦的萌芽研究, 北海道大学, 研究代表者, 競争的資金, 15K14243 - 高レイノルズ数域におけるCAMUI型ハイブリッドロケットの燃料後退機構の解明
科学研究費助成事業 基盤研究(A)
2012年04月01日 - 2016年03月31日
永田 晴紀, 大島 伸行, 戸谷 剛, 脇田 督司
火薬類や液体燃料等の危険物を使用せず、安全管理が安価で小型化が容易なCAMUI型ハイブリッドロケットの実用化を目的として、レイノルズ数で50万以上、推力で10 kN級まで地上燃焼実験を実施して燃料後退特性を取得し、燃料後退速度のスケール依存性を明らかにした。合わせて数値計算を実施し、燃料後退特性がレイノルズ数(スケール)に依存する機構を明らかにした。これらの成果により、大型モータにおいて燃料グレイン形状の最適設計が可能となり、大型化開発のために必要な基盤知識を確立することが出来た。本設計手法を適用して推力15 kN級モータを開発して燃焼実験を行い、予想通りの燃焼特性が得られることを確認した。
日本学術振興会, 基盤研究(A), 北海道大学, 研究代表者, 競争的資金, 24246135 - CAMUI型ハイブリッドロケット燃料グレイン最適形状設計手法の開発
科学研究費助成事業 基盤研究(B)
2009年 - 2011年
永田 晴紀, 戸谷 剛, 脇田 督司
申請者らが開発を進めているCAMUI型ハイブリッドロケットの燃料形状を最適に設計するための手法の構築を目指して,以下の成果を得た.1)数回の燃焼実験で全ての燃料ブロックの後退速度式を局所当量比の関数として得る手法を確立した.2)最上流前端面の燃焼特性を基礎燃焼実験により再現することに成功した.3)遺伝的アルゴリズムにより,最適燃料形状を探索する手法を開発し,目標とする最適解の探索に成功した.
日本学術振興会, 基盤研究(B), 北海道大学, 研究代表者, 競争的資金, 21360410 - 酸化剤噴流衝突場における固体燃料の燃焼機構の解明
科学研究費助成事業 基盤研究(B)
2006年 - 2008年
永田 晴紀, 戸谷 剛, 大島 伸行
申請者らが開発を進めている無火薬式小型ロケット「CAMUI型ハイブリッドロケット」の特徴的な燃焼特性を実験および数値計算により明らかにし, 以下の成果を得た.
固体燃料のガス化速度を燃料形状, 酸化剤供給量, および燃焼室圧力の関数として予測する手法を確立した.
ロケットモータのスケールが燃焼特性に与える影響を解明し, 小型モータによる燃焼実験で実機モータの燃焼特性を明らかにする手法を開発した.
日本学術振興会, 基盤研究(B), 北海道大学, 研究代表者, 競争的資金, 18360402 - 微小重力燃焼研究に基づく地下空間火災安全に関する基礎研究
科学研究費助成事業 基盤研究(C)
2004年 - 2004年
藤田 修, 永田 晴紀, 伊東 弘行, 梅村 章, 伊藤 昭彦, 菊池 政雄
本企画調査では、地下空間の火災を想定しその抑制に貢献できる研究プロジェクトを立ち上げることを目標として実施したものである。その中で特に、テグの地下鉄火災を一つの例として取り上げその災害を構成する燃焼現象を調査することで、閉鎖空間の火災現象の研究シナリオを構築した。とくに、一連の現象が浮力に強く支配されることに着目して、研究ツールとして微小重力環境利用を取り込み、閉鎖空間特有の境界条件影響を燃焼科学的な側面から明らかにしようとする研究の提案を行った。また、本研究のもう一つの特徴として、国際的な研究ネットワークを構築し、この海外研究者の協力により研究の推進を図る体制について検討した。このため、以下のような項目について調査を進め、最終的に、米国、フランス、韓国の研究者を研究グループに取り込んだ平成17年度発足特定領域研究の提案に至った。
(1)当該研究課題の研究シナリオとアプローチ法に関する検討:研究分担者と5回程度の議論の場を持ったほか、著名な火災研究者による地下空間火災事例に関する講演会を実施し、さらに研究分担者による数値計算・モデリングに関する方向性の考察を加えた研究シナリオを提案した。
(2)当該研究課題への微小重力環境利用有効性の検討:宇宙航空研究開発機構の研究者と当該研究について微小重力利用の有効性について検討した。
(3)海外研究者の研究協力体制:米国、フランス、韓国の研究者と各2回程度の議論の機会を持ち、提案予定の研究への参画見込み、海外研究者の研究資源の調査を行った。また、テグ地下鉄火災に関しては韓国研究者の協力を得て現地調査を実施した。
(4)平成17年発足特定領域研究「微小重力燃焼研究に基づくコンファインメント火災の科学」の提案:本調査研究の具体的成果として本提案を行った。
日本学術振興会, 基盤研究(C), 北海道大学, 連携研究者, 競争的資金, 16636003 - 超音速燃焼ラムジェットにおける衝撃波保炎機構の解明
科学研究費助成事業 基盤研究(C)
2001年 - 2002年
新井 隆景, 笠原 次郎, 溝端 一秀, 杉山 弘, 松尾 亜紀子, 永田 晴紀
高エンタルピー衝撃風洞内にスクラムジェットエンジンモデルを設置して,マッハ7,高度約30キロメートルの飛行条件で燃焼試験を行った。スクラムジェットエンジンモデルでは、2段ランプの1段目の圧縮部あるいは2段目のランプから燃料(ガス水素)を噴出した.上記のモデルでは、マッハ7の流れが燃焼器内部ではマッハ数約1.7まで減速された。燃焼器内の流れの静温は約1600K、静圧は約30kPaであった。実験では、流れの可視化と圧力測定を行った。実験結果は、2段目のランプ壁から燃料を噴射した場合,着火は燃焼器入り口で生じた。これは、当初の予定通りであった。特に、燃焼器入り口のカウルからの反射衝撃波後方で強い発光が認められた。このことから、流れの圧縮を受け持つインテーク部からの燃料噴射を行い、かつ、インテークから燃焼器にいたる衝撃波システムで着火を行う機構が実現できることをしめした。保炎に対しては、試験時間が短いことから、さらに詳しい実験が必要と考えられる。モデルが小さいことから、燃焼器内部での圧力上昇は観察されなかった。今後、大スケールの実験が必要である。燃料の混合促進と混合評価の観点からも研究を進め、燃料噴流と超音速流との干渉現象、境界層と超音速流の干渉現象、触媒反応を利用した混合評価方法の開発、等の研究を精力的に進めた。特に,触媒反応を用いた混合評価法では,混合状態の時間変動を捉えることに成功した.本研究結果は、日本航空宇宙学会やAIAA主催の国際会議等で、発表済みまたは発表予定である。
日本学術振興会, 基盤研究(C), 室蘭工業大学, 連携研究者, 競争的資金, 13650960 - 高速度濃度履歴が取得可能な超音速流用水素濃度プローブに関する研究
科学研究費助成事業 基盤研究(B)
2000年 - 2002年
永田 晴紀, 新井 隆景, 戸谷 剛, 工藤 勲
次世代の極超音速機用エンジンとして注目されているスクラムジェットエンジンの技術的な課題として,極超音速流中でいかに速やかに空気と燃料の混合を行うかということが挙げられる.この課題を解決するためには,超音速混合場の定量的な評価が必要である.研究者らはこれまで,安価かつ簡便な水素濃度評価方法として,触媒反応を利用した水素濃度プローブを提案してきた.これまでの研究では,触媒反応発熱量から定常状態を仮定して水素濃度の推算を行ってきたが,このプローブを乱流混合場に適用し,乱流構造に起因する高速な濃度変動履歴を計測するためには,触媒反応発熱量の非定常な変化量から水素濃度変化量を計測する手法が必要となる.そこで,水素濃度変化量が触媒反応発熱量の変化量(dQ/dt)に与える影響を明らかにすることを目的として,白金線とニッケル線(共に,直径25μm,長さ2mm)をX字型に張ったダブルプローブを用い,触媒反応が起こる条件と起こらない条件の同時測定を行い,触媒反応の応答速度を測定した.衝撃波管により水素濃度が不連続に変化する波面を発生させ,この波面に対するダブルプローブの応答を測定する.実験の結果,プローブの設定温度を概ね680K程度にすれば,発熱量の立ち上がりの勾配(dQ/dt)が設定温度に依存しなくなることが明らかとなった.このとき,(dQ/dt)は主流から白金線表面への化学種の輸送速度が決めており,(dQ/dt)から主流水素濃度を求めることができる.また,触媒反応速度は熱や化学種の輸送速度に比較して充分に早いと言え,本水素濃度プローブは熱線風速計クラスの応答速度(数十kHz〜100kHz)を達成できていると考えられる.
日本学術振興会, 基盤研究(B), 北海道大学, 研究代表者, 競争的資金, 12450390 - 高エンタルピ衝撃風洞を用いた水素の超音速混合と燃焼
科学研究費助成事業 基盤研究(C)
1999年 - 2000年
新井 隆景, 笠原 次郎, 溝端 一秀, 杉山 弘, 永田 晴紀
本研究では,将来型宇宙輸送機である完全再使用型宇宙往還機の推進システムと期待されるスクラムジェットエンジン(超音速燃焼ラムジェットエンジン)の極超音速飛行状態における性能予測と現象の把握を目的として,高エンタルピ衝撃風洞を用いた水素の超音速混合と燃焼について解明した.
まず,スクラムジェットエンジンが作動する極超音速状態を,室蘭工業大学機械システム工学科航空基礎工学講座所有の小型高エンタルピ衝撃風洞(ノズル出口直径60mm)を改良して実現した.実現できた飛行状態はマッハ数7,高度30kmであり,約2MJ/kgのエンタルピーを持つ流れを約400μs維持できる.この飛行条件は,スクラムジェットエンジンの作動条件十分満たしている.
次に,機体圧縮型のスクラムジェットエンジンモデルを製作した.流れの可視化から,機体圧縮は設計どおり行われ,マッハ約7の流れが,燃焼器内ではマッハ約1.7に減速され,静温約1,400K,圧力0.8気圧まで圧縮できることを確認した.この条件は水素の自発着火に条件を満たしている.
上述の流れ条件の対して,燃焼器内に水素を噴射し,超音速混合と超音速燃焼試験を行った.超音速混合実験では主流に窒素を用いることで,非燃焼場を実現した.流れの可視化によれば,水素は十分速やかに混合していることが確認できた.一方,主流に空気を用いた燃焼試験では,燃焼器内に発光が観察された.さらに,燃焼器内の圧力測定と高速度ビデオによる紫外波長領域の観察を行った.その結果,
1.風洞の作動時間内に紫外波長領域の発光があった.
2.そのとき,燃焼器内の壁面静圧は上昇した.
3.燃焼器内の壁面静圧の上昇は当量比が大きいほど高い.
が判明した.このことは,本研究で用いたスクラムジェットエンジンモデルで超音速燃焼が実現していることを示している.航空宇宙技術研究所角田宇宙推進研究センターのHIESTを用いてサブスケールスクラムジェットエンジンモデルの燃焼試験が行われようとしているが,本研究グループが日本で最初に高エンタルピ衝撃風洞を用いたスクラムジェットエンジンモデルの燃焼試験に成功した.
今後,本研究を発展させ,より効率的な超音速混合と燃焼を行うための燃焼器の改良やインテークと燃焼器の整合性がエンジン性能にいかに影響を及ぼすか,等について解明する.
日本学術振興会, 基盤研究(C), 室蘭工業大学, 連携研究者, 競争的資金, 11650938 - 微小重力環境下における液滴の放出と捕集に関する研究
科学研究費助成事業 基盤研究(B)
1998年 - 2000年
工藤 勲, 戸谷 剛, 永田 晴紀
平成10年度から平成12年度にかけて微小重力環境下で液滴の放出と捕集に関する実験を行った。
平成10年度は,液滴生成器を作成し,微小重力下で液滴生成器の性能試験を行った。その結果,
1.微小重力環境では表面張力が顕在化するため,ノズル出口を塞ぎ,液滴生成ができないことが懸念されていたが,微小重力下でも通常重力下と同様,液滴直径および液滴間隔が均一である均一液滴流が生成されることが確認された。
平成11年度は,回収器自体を回転させて遠心流を発生させ,遠心流に直角に当たるように入口を配置した回収管から作動流体を回収する回収器を製作し,微小重力での性能試験を行った。その結果,
2.作動流体循環で回収器の下流にあるギアポンプへ作動流体を送り込む能力があることがわかった。また,液滴の捕集については,液滴をアルミ平板に衝突させ,飛散の有無を確認した。その結果,
3.液滴直径200μm程度の均一液滴流では飛散発生せず,不均一液滴流の場合,飛散が発生することがわかった。これは,不均一液滴流内に直径の大きい液滴が含まれているためであると推測された。
平成12年度は,平成11年度の不均一液滴流内に液滴直径が大きい液滴が含まれていたため飛散したのではないかという推測を裏付けるため,ノズル直径を大きくするとともに,平成11年度の液滴回収装置を改修し,回収面角度を変更できるようにするとともに回収面上に液膜を生成できるようにして,液滴捕集に及ぼす回収面角度依存性,液膜の有無の影響を調べた。その結果,
4.微小重力環境と通常重力環境下で結果に大きな違いが見られないこと,
5.液滴直径が400μm程度の液滴であると均一液滴流でも飛散すること,
6.厚さ数mm程度の液膜が液滴直径400μm程度の液滴の飛散を防止する効果があること,
7.液滴捕集におよぼす回収面角度依存性が小さいこと,などがわかった。
日本学術振興会, 基盤研究(B), 北海道大学, 連携研究者, 競争的資金, 10450369 - 超音速流中における白金触媒燃焼に関する研究
科学研究費助成事業 奨励研究(A)
1998年 - 1999年
永田 晴紀
超音速流れ場における白金表面での触媒反応の機構を基礎的に理解し、スクラムジェットエンジンでの点火・保炎機構として触媒燃焼を応用することを目指すとともに、触媒燃焼を利用して超音速混合場を評価する手段を確立することを目的とした研究を行った。定温度型熱線風速計の原理を応用し、触媒反応発熱量と白金温度、雰囲気水素濃度の関係を明らかにした。白金線温度が十分に高い条件では、発熱量は白金線温度に依存しないことを実験的に明らかにした。発熱量は水素濃度が理論混合比のときに最大となり、酸素と水素の拡散係数の違いが白金表面への分子輸送の差にほとんど影響を及ぼさないことが判った。ほぼ全域の水素濃度において、熱伝達係数と物質伝達係数の相似性が実験的に示された。これを利用して、熱伝達係数から物質伝達係数を見積る手法が提案され、触媒反応発熱量から高精度で水素濃度を見積ることが可能であることが示された。
日本学術振興会, 奨励研究(A), 北海道大学, 研究代表者, 競争的資金, 10750654 - スクラムジェットモデル内の低温超音速混合層における触媒燃焼
科学研究費助成事業 基盤研究(C)
1997年 - 1998年
新井 隆景, 永田 晴紀, 杉山 弘
将来型宇宙輸送システムで用いられると考えられるスクラムジェットエンジンにおいて,低飛行マッハ数(マッハ5〜6程度)の場合,特に全温が低い場合においても着火・保炎が容易と考えられる触媒燃焼に着目し,超音速流中でのその特性を明らかにした.次に,この触媒燃焼をによる発熱量を定量的に測定する方法を考案し,発熱量が水素と空気の混合状態に依存することを利用して,混合状態を評価した.最後に,熱伝達と物質伝達の相似性を仮定することで,この触媒燃焼による発熱量から超音速流れ中の水素濃度を測定する新しい方法を提案し,その有用性を明らかにした。
具体的には,今年度は,昨年度の研究結果を踏まえ,マッハ1.8の後ろ向きステップを過ぎる超音速流れ中に,種々の方法で水素を噴出し,その混合状態と水素の空間分布を測定した.その結果,流れ場に対応した混合状態と水素濃度分布が得られ,本測定法の有用性を示した。
日本学術振興会, 基盤研究(C), 室蘭工業大学, 連携研究者, 競争的資金, 09651006 - 新形式ハイブリッドロケットの研究 パルスデトネーションエンジンに関する基礎研究
競争的資金 - Development of Advanced Hybrid Rockets Development of Pulse Detnation Engine
競争的資金
産業財産権
- ハイブリッドロケット
特許権, 秋 葉, 鐐二郎, 長 島, 隆 一, 棚 次, 亘 弘, 永 田, 晴 紀, 佐, 鳥, 新, 山 本, 洋, 一, 秋葉 鐐二郎, 独立行政法人 宇宙航空研究開発機構, 棚次 亘弘, 永田 晴紀, 佐鳥 新, 株式会社IHIエアロスペース
特願平11-174936, 1999年06月22日
特開2001-003813, 2001年01月09日
特許第4224599号
2008年12月05日
200903076540210472 - ハイブリッドロケット
特許権, 永田 晴紀, 独立行政法人科学技術振興機構
特願2003-285514, 2003年08月04日
特開2005-054649, 2005年03月03日
特許第4205520号
2008年10月24日
200903066296453400 - バルブレス液体供給装置
特許権, 永田 晴紀, 金子 雄大, 国立大学法人 北海道大学
特願2007-082739, 2007年03月27日
特開2008-240643, 2008年10月09日
200903066003720183 - ハイブリッドロケットエンジンの固体モータとその燃焼方法
特許権, 棚次 亘弘, 永田 晴紀, 佐鳥 新, 渡辺 裕之, 株式会社IHIエアロスペース
特願2002-302641, 2002年10月17日
特開2004-137956, 2004年05月13日
特許第4183475号
2008年09月12日
200903017661817971 - ガスタービンの燃焼器
特許権, 藤森 俊郎, 佐藤 公美, 秋葉 鐐二郎, 伊藤 献一, 永田 晴紀, 石川島播磨重工業株式会社
特願2000-317693, 2000年10月18日
特開2002-130679, 2002年05月09日
200903087204144897 - パルスデトネーションエンジン用多孔微細管燃料酸化剤供給プレート
特許権, 笠原 次郎, 新井 隆景, 永田 晴紀, 松尾 亜紀子, 笠原 次郎, 新井 隆景, 永田 晴紀, 松尾 亜紀子
特願2000-258181, 2000年07月24日
特開2002-039012, 2002年02月06日
200903098647827840