研究者データベース

脇田 督司(ワキタ マサシ)
工学研究院 機械・宇宙航空工学部門 宇宙航空システム
助教

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

  • 助教

学位

  • 博士(工学)(北海道大学)

ホームページURL

J-Global ID

研究キーワード

  • 燃焼工学   熱工学   衝撃波   航空宇宙工学   デトネーション   パルスデトネーションエンジン   超音速燃焼   ハイブリッドロケット   燃焼   

研究分野

  • フロンティア(航空・船舶) / 航空宇宙工学

職歴

  • 2008年10月 - 現在 北海道大学 工学(系)研究科(研究院) 助教
  • 2007年10月 - 2008年09月 九州工業大学 工学部 博士研究員
  • 2006年08月 - 2007年09月 北海道大学 大学院工学研究科 博士研究員

所属学協会

  • アメリカ航空宇宙学会   日本燃焼学会   日本航空宇宙学会   日本機械学会   

研究活動情報

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.
  • 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.
  • 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.
  • 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.
  • 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.
  • 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.
  • Heat Storage and Release Tests of Heat Storage Material with Crystal Transformation
    脇田督司
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 14 ists30 1 - 6 2016年 [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]
  • Estimation of Hybrid Rocket Nozzle Throat Erosion History
    脇田督司
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 14 ists30 145 - 151 2016年 [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
    脇田督司
    Mechanical Engineering Research Journal 9 79 - 85 2015年 [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.
  • Accuracy and applicable range of a reconstruction technique for hybrid rockets
    脇田督司
    Advances in Aircraft and Spacecraft Science 1 3 273 - 289 2014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 12 ists29 Ta_1 - Ta_4 2014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 12 ists29 Pa_1 - Pa_7 2014年 [査読有り][通常論文]
  • Tsuyoshi TOTANI, Ryota INOUE, Hiroto OGAWA, Tilok Kumar DAS, Masashi WAKITA, Harunori NAGATA
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 12 ists29 Pf_11 - Pf_20 2014年 [査読有り][通常論文]
  • Ken Terakawa, Tatsuya Saito, Yuji Nakamura, Tsuneyoshi Matsuoka, Harunori Nagata, Tsuyoshi Totani, Masashi Wakita
    JOURNAL OF THERMAL SCIENCE AND TECHNOLOGY 9 2 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.
  • Tsuyoshi TOTANI, Toshifumi SATOH, Masashi WAKITA, Harunori NAGATA
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 12 ists29 Po_4_1 - Po_4_5 2014年 [査読有り][通常論文]
  • 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.
  • 脇田督司
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 11 ists28 71 - 78 2013年 [査読有り][通常論文]
  • Tsuyoshi TOTANI, Takuhiro TAKEKOSHI, Masashi WAKITA, Harunori NAGATA
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 10 ists28 Pf_1 - Pf_8 2012年 [査読有り][通常論文]
  • Masashi WAKITA, Masayoshi TAMURA, Akihiro TERASAKA, Kazuya SAJIKI, Tsuyoshi TOTANI, Harunori NAGATA
    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]
  • 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.
  • 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.
  • 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 THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 2010年06月 [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...
  • The Effect of Fuel Grain Size on the Combustion Characteristics in the Primary Combustion Chamber of Staged Combustion Hybrid Rocket
    Harunori Nagata, Kenta Hashiba, Hiroya Sakai, Tsuyoshi Totani, Masashi Wakita
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 8 ists27 Pa_7 - Pa_11 2010年 [査読有り][通常論文]
  • Preliminary Thermal Design of UNITEC-1
    Tsuyoshi Totani, Haruaki Ii, Masashi Wakita, Harunori Nagata
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 8 ists27 Pf_1 - Pf_6 2010年 [査読有り][通常論文]
  • Effect of Temporal Variations of Internal Ballistics on Fuel Regression Rate in the CAMUI Hybrid Rocket
    Yudai Kaneko, Kouichi Kishida, Nobuyuki Oshima, Takuji Nakashima, Masashi Wakita, Tsuyoshi Totani, Harunori Nagata
    Journal of Space Engineering 3 1 52 - 65 2010年 [査読有り][通常論文]
  • Yudai Kaneko, Mitsunori Itoh, Akihito Kakikura, Kazuhiro Mori, Kenta Uejima, Takuji Nakashima, Masashi Wakita, Tsuyoshi Totani, Nobuyuki Oshima, Harunori Nagata
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, SPACE TECHNOLOGY JAPAN 7 ists26 Pa_77 - Pa_80 2009年 [査読有り][通常論文]
  • Masashi Wakita, Koichi Yonemoto, Tomoki Akiyama, Shigeru Aso, Yuji Kohsetsu, Harunori Nagata
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, SPACE TECHNOLOGY JAPAN 7 ists26 Tg_21 - Tg_26 2009年 [査読有り][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 THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 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.
  • 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]
  • 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 THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 2005年09月 [査読有り][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...

その他活動・業績

共同研究・競争的資金等の研究課題

  • 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
    文部科学省:科学研究費補助金(基盤研究(C))
    研究期間 : 2019年04月 -2022年03月 
    代表者 : 脇田督司
  • 文部科学省:科学研究費補助金(若手研究(B))
    研究期間 : 2013年 -2014年 
    代表者 : 脇田 督司
  • 高レイノルズ数域におけるCAMUI型ハイブリッドロケットの燃料後退機構の解明
    文部科学省:科学研究費補助金(基盤研究(A))
    研究期間 : 2012年 -2014年 
    代表者 : 永田 晴紀
  • 文部科学省:科学研究費補助金(若手研究(B))
    研究期間 : 2009年 -2011年 
    代表者 : 脇田 督司
     
    次世代航空宇宙推進用機関の1つとしてパルスデトネーションエンジン(Pulse Detonation Engine, PDE)が注目されている.実用化に向けた課題の1つにデトネーションの開始がある.この課題に対し,我々は『円錐形状反射板を用いたパルスデトネーションエンジン(PDE)イニシエータ』を提案している.昨年度は爆轟波管出口に円柱と円錐を組み合わせた形状の中子を設置することによって反射板以降の流路形状を環状に変更し,円錐部の角度を変更することによって,デトネーション波の伝播・消炎のメカニズムを観察したが,本年度はさらに円柱部の半径を変更し,流路幅5mmおよび10mmに加えて15mmおよび20mmの伝播限界および伝播形態を詳細に調査した.入射する円環デトネーション波は流路幅に関係なく,円錐角度が15°の場合では流路に沿って消えることなく伝播していたが,30°以上の角度では環状流路拡大部分でいったん消炎し円錐部分で再開始することによって円環デトネーションから平面デトネーションに遷移した.また,円環デトネーション波から平面デトネーション波へ消炎することなく遷移する場合の環状流路に存在するセルの個数と,円錐部分の角度αの関係を求めると,αが15°の場合は2.2個,30°の場合は3.2個,45°以上の場合は4.2個以上流路幅に対しセルの個数が存在すれば消炎することなく伝播することが...
  • 文部科学省:科学研究費補助金(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
     
    (1)燃料後退速度式中の係数αの取得各燃焼面の燃料後退速度式における係数αの値(局所O/Fの関数になる)にスケール効果モデルを組み合わせて、大型モータ用グレイン形状の最適化設計を行うための基礎式を構築した。(2)酸化剤噴流が衝突する燃焼流れ場の解明酸化剤噴流が衝突する燃焼流れ場を異なるスケールで構築し、燃焼場のスケールと燃料後退特性との関係を調べた。レイノルズ数および噴流の速度をスケール間で一致させることにより燃料後退速度がスケール比で相似関係になることが、本実験系においても確認された。また、レイノルズ数を一定のまま圧力を下げ、流量を増加させると、燃料後退速度が増大することが確認された。本結果は実モータにおいても観察されているものである。(3)最適グレイン形状を探索するアルゴリズムの開発CAMUIロケットの燃料グレインにおいては、縦に並べられた6~10個の燃料ブロックが同時に燃え進む。燃料グレインの燃え残りをより少なく、O/Fシフトによる比推力損失をより小さくするような、最適グレイン形状を探索するアルゴリズムを開発する。昨年度は、最適グレイン形状を得るための探索アルゴリズムとして遺伝的アルゴリズムを適用し、良好な探索速度を得た。本年度は、最適解近傍までは遺伝的アルゴリズムを適用し、ある程度良好な解が得られた後は燃え残り燃料を他のブロックに再配分する手法により最適解に至る手法を...
  • 文部科学省:科学研究費補助金(若手研究(スタートアップ))
    研究期間 : 2007年 -2008年 
    代表者 : 脇田 督司
     
    本研究は, パルスデトネーションエンジンにおける, デトネーション開始機構に関する研究である.従来のプリデトネータを用いたデトネーションイニシエータに対し, 本研究ではプリデトネータ出口に円盤状あるいは円錐形状反射板を設置し, 且つドライバーガスをプリデトネータから過供給するという2つの方法を同時に用いる開始機構を提案した.実験と数値解析により, 本開始機構を用いることにより, ドライバーガス使用量を全体の0.7~1.0 %程度に抑えることができる可能であることを実証した

大学運営

委員歴

  • 2016年02月 - 現在   第30回宇宙技術および科学の国際シンポジウム組織委員会 プログラム小委員会(Chemical Propulsion and Air-breathing Engines)委員
  • 2015年 - 現在   日本航空宇宙学会   宇宙利用部門 幹事


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