研究者データベース

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

基本情報

所属

  • 工学研究院 機械・宇宙航空工学部門 宇宙航空システム

職名

  • 教授

学位

  • 博士(工学)(東京大学)

論文上での記載著者名

  • Harunori Nagata
  • H. Nagata
  • 永田 晴紀

ホームページURL

科研費研究者番号

  • 40281787

J-Global ID

研究キーワード

  • 超音速燃焼   超音速混合   スクラムジェットエンジン   触媒燃焼   触媒反応   固体燃焼   ロケット   燃焼   宇宙インフラ   せん断混合   ハイブリッドロケット   衝突噴流熱伝達   再使用型宇宙輸送機   遺伝的アルゴリズム   デトネーション   スクラムジェット   高エンタルピー流れ   衝撃波誘起燃焼   液滴   保炎   宇宙輸送システム   液滴生成   数値解析とモデリング   微小重力   斜めデトネーションエンジン   乱流混合   コンファインメント   衝撃風洞   地下空間火災   液滴回収器   宇宙推進工学   宇宙推進工学   Space Propulsion   Combustion   Microgravity Engineering   Space Propulsion   

研究分野

  • ものづくり技術(機械・電気電子・化学工学) / 熱工学
  • フロンティア(航空・船舶) / 航空宇宙工学

職歴

  • 2007年 - 2012年 北海道大学 工学(系)研究科(研究院) 教授
  • 1994年 - 1996年 日産自動車株式会社宇宙航空事業部基盤技術部 職員(技術系)
  • 1994年 - 1996年 Technical Staff,Aerospace Div., Nissan Motor Co., Ltd.

研究活動情報

論文

  • K. Komizu, Y. Saito, A. Tsuji, H. Nagata
    Journal of Combustion 2020 2020年 [査読有り][通常論文]
     
    © 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.
  • Investigation of Graphite Nozzle Erosion in Hybrid Rockets Using O2/HDPE
    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 2020年 [査読有り][通常論文]
  • Ryohei Gotoh, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata
    INTERNATIONAL JOURNAL OF HEAT AND MASS TRANSFER 137 1132 - 1140 2019年07月 [査読有り][通常論文]
     
    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.
  • Landon Kamps, Kazuhito Sakurai, Yuji Saito, Harunori Nagata
    AEROSPACE 6 4 2019年04月 [査読有り][通常論文]
     
    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.
  • Yuji Saito, Masaya Kimino, Ayumu Tsuji, Yushi Okutani, Kentaro Soeda, Harunori Nagata
    JOURNAL OF PROPULSION AND POWER 35 2 328 - 341 2019年03月 [査読有り][通常論文]
     
    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.
  • 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][通常論文]
     
    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.
  • 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][通常論文]
     
    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.
  • 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][通常論文]
     

    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.

  • 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][通常論文]
     

    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.

  • 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 [査読無し][通常論文]
     

    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.

  • 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 2018年10月 [査読有り][通常論文]
     
    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.
  • Delburg P. Mitchao, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata
    JOURNAL OF THERMOPHYSICS AND HEAT TRANSFER 32 3 789 - 798 2018年07月 [査読有り][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     

    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.

  • 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 2018年 [査読有り][通常論文]
     
    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.
  • Yuji Saito, Toshiki Yokoi, Hiroyuki Yasukochi, Kentaro Soeda, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata
    JOURNAL OF PROPULSION AND POWER 34 1 247 - 259 2018年01月 [査読有り][通常論文]
     
    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.
  • Harunori Nagata, Hayato Teraki, Yuji Saito, Ryuichiro Kanai, Hiroyuki Yasukochi, Masashi Wakita, Tsuyoshi Totani
    JOURNAL OF PROPULSION AND POWER 33 6 1473 - 1477 2017年11月 [査読有り][通常論文]
     
    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.
  • Landon Kamps, Yuji Saito, Ryosuke Kawabata, Masashi Wakita, Tsuyoshi Totani, Yusuke Takahashi, Harunori Nagata
    JOURNAL OF PROPULSION AND POWER 33 6 1369 - 1377 2017年11月 [査読有り][通常論文]
     
    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.
  • 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 2017年05月 [査読有り][通常論文]
     
    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.
  • 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 2017年 [査読有り][通常論文]
     
    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 2017年 [査読有り][通常論文]
     
    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.
  • 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][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • Tsuneyoshi Matsuoka, Kyohei Kamei, Yuji Nakamura, Harunori Nagata
    INTERNATIONAL JOURNAL OF AEROSPACE ENGINEERING 2017 2017年 [査読有り][通常論文]
     
    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.
  • 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][通常論文]
  • 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 日本航空宇宙学会 2016年09月06日 [査読無し][通常論文]
  • 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][通常論文]
  • Development of Wall Regression Model of Hybrid Rocket Solid Fuel
    脇田督司
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 14 ists30 67 - 72 2016年 [査読有り][通常論文]
  • 脇田督司
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 14 ists30 39 - 44 2016年 [査読有り][通常論文]
  • 永田 晴紀
    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][通常論文]
  • 永田 晴紀
    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][通常論文]
  • 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 2015年11月 [査読有り][通常論文]
  • 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][通常論文]
  • 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 [査読無し][招待有り]
     
    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.
  • Kouichiro Tani, Susumu Hasegawa, Shuichi Ueda, Takeshi Kanda, Harunori Nagata
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 58 4 228 - 236 2015年 [査読有り][通常論文]
     
    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.
  • 永田 晴紀
    Progress in Scale Modeling 2 249 - 263 2014年11月 [査読有り][招待有り]
  • 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][通常論文]
  • Harunori Nagata, Hisahiro Nakayama, Mikio Watanabe, Masashi Wakita, Tsuyoshi Totani
    Advances in Aircraft and Spacecraft Science 1 3 273 - 289 2014年07月 [査読有り][通常論文]
  • Tsuyoshi Totani, Hiroto Ogawa, Ryota Inoue, Tilok K. Das, Masashi Wakita, Harunori Nagata
    JOURNAL OF THERMOPHYSICS AND HEAT TRANSFER 28 3 524 - 533 2014年07月 [査読有り][通常論文]
     
    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.
  • 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 2014年04月 [査読有り][通常論文]
  • 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][通常論文]
     
    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.
  • 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][通常論文]
     
    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.
  • Ken Terakawa, Tatsuya Saito, Yuji Nakamura, Tsuneyoshi Matsuoka, Harunori Nagata, Tsuyoshi Totani, Masashi Wakita
    JOURNAL OF THERMAL SCIENCE AND TECHNOLOGY 9 2 JTST0010 - JTST0010 2014年 [査読有り][通常論文]
     
    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.
  • 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][通常論文]
  • 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 [査読無し][通常論文]
  • 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][通常論文]
  • 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 2013年08月 [査読有り][通常論文]
  • 松岡 常吉, 永田 晴紀, 中村 祐二
    実験力学 13 2 178 - 184 2013年06月 [査読有り][通常論文]
  • 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 [査読無し][招待有り]
  • 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 [査読無し][通常論文]
  • 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][通常論文]
     
    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.
  • 永田晴紀, 飯島直純, 金井竜一朗
    日本燃焼学会誌 54 170 251-258  2012年11月 [査読無し][招待有り]
  • 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][通常論文]
  • 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][通常論文]
  • 戸谷剛, 佐藤敏文, 脇田督司, 永田晴紀
    Thermophys Prop 33rd 326-328  2012年10月 [査読有り][通常論文]
  • 戸谷剛, 脇田督司, 永田晴紀
    Thermophys Prop 33rd 227-229  2012年10月 [査読有り][通常論文]
  • Tsuneyoshi Matsuoka, Shota Murakami, Harunori Nagata
    COMBUSTION AND FLAME 159 7 2466 - 2473 2012年07月 [査読有り][通常論文]
     
    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.
  • 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− 11-015 27-32  2012年03月30日 [査読無し][通常論文]
  • TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 10 ists28 Pa_31-Pa_36  2012年 [査読有り][通常論文]
  • TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 10 ists28 To_1_1-To_1_5  2012年 [査読有り][通常論文]
  • TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 10 ists28 1 - 8 2012年 [査読有り][通常論文]
  • 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][通常論文]
  • Yuki Iwaki, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 54 185-86 212 - 220 2011年11月 [査読有り][通常論文]
     
    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.
  • 永田晴紀
    自動車技術 65 10 56 - 61 2011年10月 [査読無し][招待有り]
  • 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  2011年 [査読有り][通常論文]
  • Masashi Wakita, Ryusuke Numakura, Takatoshi Asada, Masayoshi Tamura, Tsuyoshi Totani, Harunori Nagata
    JOURNAL OF PROPULSION AND POWER 27 1 162 - 170 2011年01月 [査読有り][通常論文]
     
    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.
  • Tsuneyoshi Matsuoka, Harunori Nagata
    ACTA ASTRONAUTICA 68 1-2 197 - 203 2011年01月 [査読有り][通常論文]
     
    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.
  • 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 586 - 589 2010年12月 [査読有り][通常論文]
  • 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][通常論文]
  • 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 2010年07月 [査読有り][通常論文]
  • 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05月 [査読有り][通常論文]
     
    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
    Journal of Space Engineering 3 1 52 - 65 2010年 [査読無し][通常論文]
  • Preliminary Thermal Design of UNITEC-1
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 8 ists27 Pf1-Pf6  2010年 [査読無し][通常論文]
  • 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][通常論文]
     
    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
    APPLICATIONS OF SPACE TECHNOLOGY FOR HUMANITY 138 629 - + 2010年 [査読有り][通常論文]
     
    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
    APPLICATIONS OF SPACE TECHNOLOGY FOR HUMANITY 138 611 - + 2010年 [査読有り][通常論文]
     
    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.
  • 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  2009年 [査読有り][通常論文]
  • Transactions of the Japan Society for Aeronautical and Space Sciences, Aerospace Technology Japan 7 ists26 Tu_1-Tu_5  2009年 [査読無し][通常論文]
  • Transactions of the Japan Society for Aeronautical and Space Sciences, Aerospace Technology Japan 7 ists26 Pa_77-Pa_80  2009年 [査読無し][通常論文]
  • Transactions of the Japan Society for Aeronautical and Space Sciences, Aerospace Technology Japan 7 ists26 Pb_71-Pb_76  2009年 [査読無し][通常論文]
  • 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][通常論文]
     
    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 ...
  • Masashi Wakita, Ryusuke Numakura, Yusuke Itoh, Shigetoshi Sugata, Tsuyoshi Totani, Harunori Nagata
    JOURNAL OF PROPULSION AND POWER 23 2 338 - 344 2007年03月 [査読有り][通常論文]
     
    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.
  • 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 55 646 516 - 526 2007年 [査読無し][通常論文]
  • Nozomu Hashimoto, Harunori Nagata, Tsuyoshi Totani, Isao Kudo
    COMBUSTION AND FLAME 147 3 222 - 232 2006年11月 [査読有り][通常論文]
     
    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.
  • 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][通常論文]
  • T Totani, T Kodama, K Watanabe, K Nanbu, H Nagata, Kudo, I
    ACTA ASTRONAUTICA 59 1-5 192 - 199 2006年07月 [査読無し][通常論文]
     
    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.
  • Harunori Nagata, Mitsunori Ito, Takenori Maeda, Mikio Watanabe, Tsutomu Uematsu, Tsuyoshi Totani, Isao Kudo
    ACTA ASTRONAUTICA 59 1-5 253 - 258 2006年07月 [査読有り][通常論文]
     
    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 10kg vehicle equipped with a 50kgf thrust CAMUI motor have shown the feasibility of the motor system. The meteorological observation model of 400kgf class motor is under development and the development of microgravity experiment class of 1.5-2tonf motor will follow subsequently. The authors plan to complete the development of the 400kgf class motor for meteorological observation model by the end of FY2005. (C) 2006 Elsevier Ltd. All rights reserved.
  • The Challenges of Developing a Hybrid Launch System for Microsatellite Payloads Using CAMUI Hybrid Upperstages with the Rocketplane XP
    Proceedings of the 25th ISTS Selected Papers 821 - 827 2006年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
     
    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 2005年05月 [査読無し][通常論文]
     
    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.
  • 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 2005年 [査読無し][通常論文]
  • T Totani, T Kodama, K Watanabe, H Nagata, Kudo, I
    MICROGRAVITY SCIENCE AND TECHNOLOGY 17 3 31 - 38 2005年 [査読無し][通常論文]
     
    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.
  • 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][通常論文]
  • 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 2004年11月04日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • Nakamura, D., Nagata, H., Totani, T. and Kudo, I.:"Relationship between Platinum Wire Temperature and Catalytic Heat Release Rate on Platinum in Unsteady-State Hydrogen-Air Mixture", Heat Transfer - Asian Research, 33(1): 1-11 (2004)*
    2004年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
  • 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01月25日 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
  • 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 2003年 [査読無し][通常論文]
  • 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 2003年 [査読無し][通常論文]
  • 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 2003年 [査読有り][通常論文]
     
    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.
  • 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10月25日 [査読無し][通常論文]
     
    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)...
  • 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 [査読無し][通常論文]
  • 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04月25日 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
  • 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 2002年02月 [査読有り][通常論文]
     
    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 [査読無し][通常論文]
  • 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 2002年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 2002年 [査読無し][通常論文]
  • 「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 [査読無し][通常論文]
  • 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 2001年10月01日 [査読無し][通常論文]
  • 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04月25日 [査読無し][通常論文]
     
    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...
  • 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03月25日 [査読無し][通常論文]
     
    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 ...
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 2001年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 2001年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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08月25日 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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 ...
  • 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 [査読無し][通常論文]
     
    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.
  • Harunori NAGATA, Keiji OKADA, Takashi SAN'DA, Takahiro KATO, Ryojiro AKIBA, Shin SATORI, and Isao KUDO, "Combustion Characteristics of Fibrous Fuels for Dry Towel Hybrid Rocket Motor," Journal of Space Technology and Science, Vol.13, No.1, pp.1 (1999).
    1999年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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」
    日本マイクログラビティ応用学会誌 Vol.16 No.3 191 - 197 1999年 [査読無し][通常論文]
  • 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 1999年 [査読無し][通常論文]
  • 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 1999年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
     
    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 ...
  • 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 [査読無し][通常論文]
     
    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...
  • 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 1998年 [査読無し][通常論文]
  • 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 1998年 [査読無し][通常論文]
  • 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 1998年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 1997年 [査読無し][通常論文]
  • 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 1997年 [査読無し][通常論文]

書籍

講演・口頭発表等

その他活動・業績

  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 2019年11月 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     

    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.

  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 津地 歩, 永田 晴紀 年次大会 2018 (0) J1920101 2018年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    © 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%.
  • Yushi Okutani, Yuji Saito, Masaya Kimino, Ayumu Tsuji, Kentaro Soeda, Harunori Nagata 2018 Joint Propulsion Conference 2018年01月01日 [査読無し][通常論文]
     
    © 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.
  • Yuji Saito, Landon Kamps, Kodai Komizu, Kentaro Soeda, Daniele Bianchi, Francesco Nasuti, Harunori Nagata 2018 Joint Propulsion Conference 2018年01月01日 [査読無し][通常論文]
     
    © 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.
  • Ryohei Gotoh, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata International Heat Transfer Conference 2018-August 4425 -4432 2018年01月01日 [査読無し][通常論文]
     
    © 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.
  • Erika Uchiyama, Yurika Kiyotani, Landon Kamps, Harunori Nagata Advances in the Astronautical Sciences 166 109 -115 2018年01月01日 [査読無し][通常論文]
     
    © 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.
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 永田 晴紀 翼 : 航空自衛隊連合幹部会機関誌 41 (0) 105 -109 2017年10月 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 片野田 洋, 永田 晴紀 鹿児島大学工学部研究報告 (58) 1 -6 2017年03月 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     

    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.

  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • Landon Kamps, Shota Hirai, Yassine Ahmimache, Raymond Guan, Harunori Nagata 53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017 2017年01月01日 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • Daniele Bianchi, Landon Kamps, Francesco Nasuti, Harunori Nagata 53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017 2017年01月01日 [査読無し][通常論文]
     
    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.
  • T. Viscor, H. Nagata 53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017 2017年01月01日 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 永田 晴紀 翼 : 航空自衛隊連合幹部会機関誌 40 (0) 62 -64 2016年07月 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    © 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.
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 年次大会 2016 (0) S1920202 2016年 [査読無し][通常論文]
     

    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.

  • 山口 亮, 川端 良輔, 平井 翔大, ケンプス ランドン, 脇田 督司, 戸谷 剛, 永田 晴紀 年次大会 2016 (0) S1920201 2016年 [査読無し][通常論文]
     

    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.

  • Landon Kamps, Yuji Saito, Ryosuke Kawabata, Yusuke Takahashi, Harunori Nagata 52nd AIAA/SAE/ASEE Joint Propulsion Conference, 2016 2016年01月01日 [査読無し][通常論文]
     
    © 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%.
  • Yuji Saito, Toshiki Yokoi, Hiroyuki Yasukochi, Kentaro Soeda, Tsuyoshi Totani, Masashi Wakita, Harunori Nagata Proceedings of the International Astronautical Congress, IAC 2016年01月01日 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
  • 横井俊希, 齋藤勇士, 脇田督司, 戸谷剛, 永田晴紀 燃焼シンポジウム講演論文集 53rd 308 -309 2015年11月04日 [査読無し][通常論文]
  • 高梨知広, 戸谷剛, 脇田督司, 永田晴紀 Thermophysical Properties 36th 180‐182 2015年10月19日 [査読無し][通常論文]
  • 戸谷剛, 戸谷剛, 近藤良夫, 山木宏, 脇田督司, 永田晴紀 Thermophys Prop 36th 213 -215 2015年10月19日 [査読無し][通常論文]
  • 丸 祐介, 澤井 秀次郎, 永田 晴紀 宇宙科学技術連合講演会講演集 59 6p 2015年10月07日 [査読無し][通常論文]
  • 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- (14-010) 173 -178 2015年03月25日 [査読無し][通常論文]
  • 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 2015年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 高梨知広, 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.3M09 2015年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    © 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.
  • 色川俊雄, 戸谷剛, 永田晴紀, 脇田督司 Thermophys Prop 35th 40 -42 2014年11月22日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 川端 良輔, 永田 晴紀, 戸谷 剛 宇宙科学技術連合講演会講演集 58 4p 2014年11月12日 [査読無し][通常論文]
  • 高梨 知広, 戸谷 剛, 永田 晴紀 宇宙科学技術連合講演会講演集 58 3p 2014年11月12日 [査読無し][通常論文]
  • Harunori Nagata, Hisahiro Nakayama, Mikio Watanabe, Masashi Wakita, Tsuyoshi Totani Advances in Aircraft and Spacecraft Science 1 273 -289 2014年07月01日 [査読無し][通常論文]
     
    © 2014 Techno-Press, Ltd. 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.
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 川端 良輔, 稲場 康彦, 石山 達也, 脇田 督司, 戸谷 剛, 永田 晴紀 年次大会 2014 (0) _S1920103 --_S1920103- 2014年 [査読無し][通常論文]
     
    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.
  • 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 2014年 [査読無し][通常論文]
  • 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) 51st ROMBUNNO.H341 2014年 [査読無し][通常論文]
  • 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 2014年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 高梨知広, 戸谷剛, 永田晴紀, 脇田督司 Thermophys Prop 34th 141 -143 2013年11月20日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 戸谷剛, 色川俊雄, 脇田督司, 永田晴紀 Thermophys Prop 34th 307 -309 2013年11月20日 [査読無し][通常論文]
  • 戸谷剛, 國拓也, 佐藤敏文, 脇田督司, 永田晴紀 Thermophys Prop 34th 96 -98 2013年11月20日 [査読無し][通常論文]
  • 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 2013年10月18日 [査読無し][通常論文]
  • 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 2013年10月18日 [査読無し][通常論文]
  • 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 2013年10月18日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 高梨 知広, 脇田 督司, 永田 晴紀 宇宙科学技術連合講演会講演集 57 4p 2013年10月09日 [査読無し][通常論文]
  • Harunori Nagata, Masashi Wakita, Tsuyoshi Totani, Tsutomu Uematsu 49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference 2013年09月16日 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][招待有り]
  • 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 [査読無し][通常論文]
     
    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.
  • 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 年次大会 2013 (0) _S192022 -1-_S192022-5 2013年 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 戸谷剛, 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) 50th ROMBUNNO.G123 2013年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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) 57th ROMBUNNO.2I03 2013年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 戸谷剛, 佐藤敏文, 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) 50th ROMBUNNO.F131 2013年 [査読無し][通常論文]
  • 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][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 小川 洋人, 井上 遼太, 戸谷 剛, 脇田 督司, 永田 晴紀 年次大会 2012 (0) _S192013 -1-_S192013-5 2012年 [査読無し][通常論文]
     
    "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.
  • 石山 達也, 戸谷 剛, 永田 晴紀, 稲場 康彦, 井上 遼太, 佐々木 俊也, 寺川 健, 桧物 恒太郎, 李 尚駿, 金井 竜一朗, 脇田 督司 年次大会 2012 (0) _S192023 -1-_S192023-5 2012年 [査読無し][通常論文]
     
    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.
  • 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 年次大会 2012 (0) _S192022 -1-_S192022-4 2012年 [査読無し][通常論文]
     
    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.
  • 井上 遼太, 小川 洋人, 戸谷 剛, 脇田 督司, 永田 晴紀 年次大会 2012 (0) _S192011 -1-_S192011-5 2012年 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 戸谷剛, 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) 56th ROMBUNNO.1N05 2012年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 石川直幸, 戸谷剛, 脇田督司, 永田晴紀 Thermophys Prop 32nd 307 -309 2011年11月21日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 松岡常吉, 村上翔太, 中村祐二, 永田晴紀 燃焼シンポジウム講演論文集 49th 258 -259 2011年11月20日 [査読無し][通常論文]
  • 永田晴紀, 金子雄大, 脇田督司, 戸谷剛 燃焼シンポジウム講演論文集 49th 248 -249 2011年11月20日 [査読無し][通常論文]
  • 永田晴紀 化学工学会大会講演要旨集(CD-ROM) 2011 ROMBUNNO.E122 2011年07月25日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 石川直幸, 戸谷剛, 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) 48th ROMBUNNO.H312 2011年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 戸谷剛, 小川洋人, 井上遼太, 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.2G16 2011年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 野原 正寛, 出雲 弘一, 脇田 督司, 戸谷 剛, 永田 晴紀 年次大会 2011 (0) _S192013 -1-_S192013-5 2011年 [査読無し][通常論文]
     
    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).
  • 小川 洋人, 井上 遼太, 戸谷 剛, 脇田.督司, 永田 晴紀 年次大会 2011 (0) _S192022 -1-_S192022-5 2011年 [査読無し][通常論文]
     
    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.
  • Tsuneyoshi Matsuoka, Harunori Nagata 61st International Astronautical Congress 2010, IAC 2010 4 2863 -2869 2010年12月01日 [査読無し][通常論文]
     
    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.
  • Yudai Kaneko, Mitsunori Itoh, Massasi Wakita, Tsuyoshi Totani, Harunori Nagata Advances in the Astronautical Sciences 138 629 -634 2010年12月01日 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • Koichi Kishida, Yudai Kaneko, Nobuyuki Oshima, Harunori Nagata Nihon Kikai Gakkai Ronbunshu, B Hen/Transactions of the Japan Society of Mechanical Engineers, Part B 76 (765) 789 -794 2010年05月01日 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
  • 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 2010年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • Tsuneyoshi Matsuoka, Harunori Nagata 60th International Astronautical Congress 2009, IAC 2009 8 6293 -6300 2009年12月01日 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 田村正佳, 脇田督司, 永田晴紀, 戸谷剛 燃焼シンポジウム講演論文集 47th 72-73 2009年11月18日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 堺裕哉, 榎本剛矩, 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 300-301 2009年11月18日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 室井 典和, 登坂 茂, 永田 晴紀 北海道工業大学研究紀要 0 (37) 141 -146 2009年03月 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 脇田督司, 田村正佳, 戸谷剛, 永田晴紀 宇宙科学技術連合講演会講演集(CD−ROM) 53rd 2G14 2009年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 仁木雄大, 竹腰卓博, 戸谷剛, 永田晴紀, 脇田督司 宇宙科学技術連合講演会講演集(CD−ROM) 53rd 2J02 2009年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 金子雄大, 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) 53rd 2B06 2009年 [査読無し][通常論文]
  • 伊井晴明, 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 2F11 2009年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 松岡常吉, 戸谷剛, 永田晴紀 燃焼シンポジウム講演論文集 46th 194-195 2008年11月20日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 永田 晴紀 伝熱 : journal of the Heat Transfer Society of Japan 47 (199) 23 -29 2008年04月01日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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) 2008年03月05日 [査読無し][通常論文]
  • 戸谷剛, 南部航太, 川上哲人, 由利泰史, 永田晴紀 宇宙利用シンポジウム 24th 117-120 2008年03月 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 岩城裕樹, 長沼哲史, 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 1-6 2008年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 榊原隆浩, 伊井晴明, 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 55-60 2008年 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 永田 晴紀 エンジンテクノロジー 9 (4) 29 -33 2007年08月 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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).
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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 ...
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • Harunori Nagata, Mitsunori Ito, Takenori Maeda, Mikio Watanabe, Tsutomu Uematsu, Tsuyoshi Totani, Isao Kudo Acta Astronautica 59 253 -258 2006年07月01日 [査読無し][通常論文]
     
    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.
  • 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) 2006年03月08日 [査読無し][通常論文]
  • 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) 2006年03月08日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 永田 晴紀 日本機械学會誌 108 (1043) 2005年10月05日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • T Totani, T Kodama, H Nagata, Kudo, I JOURNAL OF SPACECRAFT AND ROCKETS 42 (3) 493 -499 2005年05月 [査読無し][通常論文]
     
    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.
  • 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) 2005年03月07日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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...
  • 永田 晴紀 日本機械学會誌 108 (1034) 2005年01月05日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 航空宇宙工学便覧第三版
    日本航空宇宙学会 2005年 [査読無し][通常論文]
  • 永田 晴紀 化学工学 = CHEMICAL ENGINEERING OF JAPAN 68 (12) 738 -739 2004年12月05日 [査読無し][通常論文]
  • 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 42nd 275-276 2004年11月20日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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) 2004年03月08日 [査読無し][通常論文]
  • T Totani, M Itami, H Nagata, Kudo, I, A Iwasaki REVIEW OF SCIENTIFIC INSTRUMENTS 75 (2) 515 -523 2004年02月 [査読無し][通常論文]
     
    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.
  • 沼倉龍介, 脇田督司, 永田晴紀, 戸谷剛, 工藤勲 日本航空宇宙学会北部支部講演会講演論文集 2004 49-54 2004年 [査読無し][通常論文]
  • Daisuke Nakamura, Harunori Nagata, Tsuyoshi Totani, Isao Kudo Heat Transfer - Asian Research 33 1 -11 2004年01月01日 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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) 2003年03月03日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • Daisuke Nakamura, Harunori Nagata, Tsuyoshi Totani, Isao Kudo Nippon Kikai Gakkai Ronbunshu, B Hen/Transactions of the Japan Society of Mechanical Engineers, Part B 69 126 -131 2003年01月01日 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 橋本望, 渡辺賢, 永田晴紀, 戸谷剛, 工藤勲 燃焼シンポジウム講演論文集 40th 163-164 2002年11月01日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 2780 -2787 2002年10月01日 [査読無し][通常論文]
     
    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.
  • H Nagata, Kudo, I, K Ito, S Nakamura, Y Takeshita COMBUSTION AND FLAME 129 (4) 392 -400 2002年06月 [査読無し][通常論文]
     
    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.
  • 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 1166 -1173 2002年04月01日 [査読無し][通常論文]
     
    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.
  • T Totani, M Itami, H Nagata, Kudo, I, A Iwasaki, S Hosokawa MICROGRAVITY SCIENCE AND TECHNOLOGY 13 (2) 42 -45 2002年 [査読無し][通常論文]
     
    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.
  • Ryojiro Akiba, Takashi Nakajima, Harunori Nagata Advances in the Astronautical Sciences 110 325 -329 2002年01月01日 [査読無し][通常論文]
     
    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.
  • Nozomu Hashimoto, Satoshi Watanabe, Harunori Nagata, Tsuyoshi Totani, Isao Kudo Proceedings of the Combustion Institute 29 245 -250 2002年01月01日 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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...
  • 橋本望, 渡辺賢, 永田晴紀, 戸谷剛, 工藤勲 燃焼シンポジウム講演論文集 39th 511-512 2001年11月01日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
  • 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 67 934 -939 2001年04月01日 [査読無し][通常論文]
     
    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.
  • 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) 2001年03月07日 [査読無し][通常論文]
  • 永田 晴紀 日本機械学会誌 104 (990) 337 -337 2001年 [査読無し][通常論文]
  • 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 38th 511-512 2000年11月01日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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 ...
  • 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 [査読無し][通常論文]
     
    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 ...
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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...
  • 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 [査読無し][通常論文]
     
    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 ...
  • K Takahashi, H Nagata, Kudo, I JOURNAL OF SPACECRAFT AND ROCKETS 37 (5) 707 -708 2000年09月 [査読無し][通常論文]
     
    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.
  • H Nagata, M Sasaki, T Arai, T Totani, Kudo, I PROCEEDINGS OF THE COMBUSTION INSTITUTE 28 713 -719 2000年 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    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 [査読無し][通常論文]
     
    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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 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 2666 -2671 1999年12月01日 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
  • 新井 隆影, 三浦 淳二, 咲間 文順, 永田 晴紀 年会一般講演 18 (0) 113 -114 1999年07月29日 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    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.
  • 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 [査読無し][通常論文]
     
    © 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.
  • 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 64 793 -799 1998年12月01日 [査読無し][通常論文]
     
    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.
  • 新井 隆景, 森田 修至, 福添 英夫, 永田 晴紀 年会一般講演 17 (0) 239 -240 1998年07月 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
     
    © 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.
  • Takakage Arai, Akira Endo, Harunori Nagata, Hiromu Sugiyama, Shuji Morita Nippon Kikai Gakkai Ronbunshu, B Hen/Transactions of the Japan Society of Mechanical Engineers, Part B 63 3318 -3324 1997年10月01日 [査読無し][通常論文]
     
    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.
  • 新井 隆景, 永田 晴紀, 杉山 弘, 森田 修至, 細川 博 年会一般講演 16 (0) 263 -264 1997年07月 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • 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 [査読無し][通常論文]
  • H. Nagata, H. M. Kim, J. Sato, M. Kono Symposium (International) on Combustion 25 1719 -1725 1994年01月01日 [査読無し][通常論文]
     
    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.
  • 永田晴紀 第31回燃焼シンポジウム講演集 319 -321 1993年 [査読無し][通常論文]

特許

受賞

  • 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 
    受賞者: 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
     
    受賞者: 永田 晴紀
  • 2018年04月 日本機械学会 フェロー会員
     
    受賞者: 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 
    受賞者: 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
     
    受賞者: 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 
    受賞者: 永田 晴紀
  • 2006年 SI2006優秀講演賞
  • 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 
    受賞者: 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 
    受賞者: 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 
    受賞者: 永田 晴紀

共同研究・競争的資金等の研究課題

  • ハイブリッドロケットノズル浸食の機構解明
    日本学術振興会:科学研究費助成事業 基盤研究(B)
    研究期間 : 2019年04月 -2022年03月 
    代表者 : 永田 晴紀, 脇田 督司
  • 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
    日本学術振興会:科学研究費助成事業 基盤研究(C)
    研究期間 : 2019年04月 -2022年03月 
    代表者 : 脇田 督司, 永田 晴紀
  • 端面燃焼式ハイ ブリッドロケットの実用化研究
    日本学術振興会:科学研究費助成事業 基盤研究(C)
    研究期間 : 2017年04月 -2020年03月 
    代表者 : 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」についての実用化を目指している.本目標を達成するために,平成30年度は下記の3つについて研究を進めてきた. 1.造形燃料の高精度大型化及び量産化: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.応答特性の解明:燃焼に必要な酸化剤供給ラインにバルブ4つを並列に接続し,バルブのON/OFF制御によって酸化剤流量を増加/減少させて燃焼実験を行った.急激に酸化剤流量を増減させたとき,燃焼室圧力の立ち上がりに大幅な時間(応答時間)を要することが実験によって明らかとなった.また応答時間を詳細に検討するために,ニードルバルブ上にステッピングモータを設置しモータ開度によって酸化剤流量を細かく変化させて燃焼実験を行った.酸化剤流量の変化量が大きいときに応答時間も大きくなる関係を実験によって示した. 3.燃焼機構の解明:燃焼速度は火炎から固体燃料への局所熱伝達率で決まり,局所熱伝達率は火炎温度と消炎距離により決まると考えられる.消炎距離は化学反応速度に支配されている可能性が高いが,これを実験的に確認するために,これまでの純酸素に加え,亜酸化窒素を酸化剤とした単ポートでの燃焼実験を行った.燃料後退速度は純酸素と比較すると亜酸化窒素を酸化剤としたとき,1/10程度減少することが分かった.
  • 超小型火星探査機用Ne計測装置の基礎開発
    日本学術振興会:科学研究費助成事業 基盤研究(A)
    研究期間 : 2017年04月 -2020年03月 
    代表者 : 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。
  • 文部科学省:基盤研究(C)
    研究期間 : 2017年04月 -2020年03月 
    代表者 : 添田 建太郎
  • 日本学術振興会:科学研究費助成事業 基盤研究(B)
    研究期間 : 2015年04月 -2018年03月 
    代表者 : 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:科学研究費助成事業 基盤研究(A)
    研究期間 : 2013年05月 -2018年03月 
    代表者 : 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:科学研究費助成事業 挑戦的萌芽研究
    研究期間 : 2015年04月 -2017年03月 
    代表者 : 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:科学研究費助成事業 基盤研究(A)
    研究期間 : 2012年04月 -2016年03月 
    代表者 : 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:科学研究費助成事業 基盤研究(B)
    研究期間 : 2009年 -2011年 
    代表者 : 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:科学研究費助成事業 基盤研究(B)
    研究期間 : 2006年 -2008年 
    代表者 : 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:科学研究費助成事業 基盤研究(C)
    研究期間 : 2004年 -2004年 
    代表者 : 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:科学研究費助成事業 基盤研究(C)
    研究期間 : 2001年 -2002年 
    代表者 : 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:科学研究費助成事業 基盤研究(B)
    研究期間 : 2000年 -2002年 
    代表者 : 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:科学研究費助成事業 基盤研究(C)
    研究期間 : 1999年 -2000年 
    代表者 : 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:科学研究費助成事業 基盤研究(B)
    研究期間 : 1998年 -2000年 
    代表者 : 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:科学研究費助成事業 奨励研究(A)
    研究期間 : 1998年 -1999年 
    代表者 : 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:科学研究費助成事業 基盤研究(C)
    研究期間 : 1997年 -1998年 
    代表者 : 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。
  • 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
  • Development of Advanced Hybrid Rockets Development of Pulse Detnation Engine

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