Curriculum Vitaes

Toru Shimada

  (嶋田 徹)

Profile Information

Affiliation
Professor Emeritus (Professor Emeritus), Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency
Specially Appointed Professor, Faculty of Science and Engineering, Department of Aerospace Engineering, Nihon University
Degree
Doctor of Engineering(Mar, 1985, The University of Tokyo)
工学修士(Mar, 1982, 東京大学)
工学学士(Mar, 1980, 京都大学)

J-GLOBAL ID
200901053726642200
researchmap Member ID
1000304541

External link

嶋田 徹(しまだ とおる)
宇宙航空研究開発機構 名誉教授

日本大学理工学部航空宇宙学科特任教授
1985年 東京大学大学院工学系研究科航空学専門課程修了・工学博士取得。1985年~2000年まで日産自動車(株)宇宙航空事業部にてロケットの設計解析に従事。2000年 旧文部省宇宙科学研究所(現:宇宙航空研究開発機構)助教授。2007年より同教授。2003年~2007年までM-Vロケットプロジェクト・ファンクションマネージャ。同ロケットの開発と打ち上げに従事。その間、北海道大学、総合研究大学院大学、東京大学で客員助教授を経て、2007年より東京大学大学院 客員教授。専門は宇宙推進流体工学、固体/ハイブリッドロケット内部の燃焼流の研究。低コストで安全なロケットの実現を目指し、2008年 よりハイブリッドロケット研究WGを主宰。2020年 宇宙飛翔工学研究系研究主幹。2021年3月 定年退職。2021年4月 再雇用(専任教授)を経て 2023年3月 退職。2023年4月 宇宙航空研究開発機構 名誉教授。2023年6月 34th International Symposium on Space Technology and Science 組織委員長。2024年4月 日本大学理工学部特任教授。


Major Papers

 18
  • 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  Peer-reviewed
    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.
  • Masaki Adachi, Toru Shimada
    AIAA JOURNAL, 53(6) 1578-1589, Jun, 2015  Peer-reviewed
    Numerical analysis on the instability of liquid/dense fluid films under supercritical operating conditions is performed on methane fuel. A numerical code for compressible fluid flows, accommodated for the van der Waals equation of state, is developed in order to deal with supercritical fluid and dense fluid layers and has shown good convergence, even at a very low-Reynolds-number flow typically seen in actual hybrid rocket engines. A linear instability analysis is conducted and shows that an amplification rate has a peak at a certain wave number of initial perturbations. The perturbation becomes unstable as the Reynolds number and chamber pressure increase, and the instability region of the wave number is enlarged when an acceleration body force in the streamwise direction is imposed. A limit cycle of the amplitude of perturbations is observed at low-Reynolds-number flows, and the instability of dense fluid layers leads to the entrainment phenomena at high-Reynolds-number flows. It is deduced that the perturbation with the peak value of the amplification rate dominates in an actual hybrid rocket engine.
  • Toru SHIMADA, Kazushige KATO, Nobuhiro SEKINO, Nobuyuki TSUBOI, Yoshio SEIKE, Mihoko FUKUNAGA, Yu DAIMON, Hiroshi HASEGAWA, Hiroya ASAKAWA
    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN, 8(ists27) Pa_29-Pa_37-Pa_37, 2010  Peer-reviewed
    In this paper, described is the development of a numerical simulation system, what we call "Advanced Computer Science on SRM Internal Ballistics (ACSSIB)", for the purpose of improvement of performance and reliability of solid rocket motors (SRM). The ACSSIB system is consisting of a casting simulation code of solid propellant slurry, correlation database of local burning-rate of cured propellant in terms of local slurry flow characteristics, and a numerical code for the internal ballistics of SRM, as well as relevant hardware. This paper describes mainly the objectives, the contents of this R&D, and the output of the fiscal year of 2008.
  • Jean-Francois Guery, I-Shih Chang, Toru Shimada, Marilyn Glick, Didier Boury, Eric Robert, John Napior, Robert Wardle, Christian Perut, Max Calabro, Robert Glick, Hiroto Habu, Nobuhiro Sekino, Gilles Vigier, Bruno d'Andrea
    ACTA ASTRONAUTICA, 66(1-2) 201-219, Jan, 2010  Peer-reviewed
    For the last 50 years solid propulsion has successfully created a multitude of small launchers and many first stages or boosters for heavy launchers with low risk, high performance. competitive cost, superb storability, and "instant" readiness in many countries. Technical Support for these successes arose from simple designs, very high thrust levels, and low development and operation costs/risks. The first solid propulsion roadmap based on these foundations and rational projections was published in 2000 [A. Davenas, D. Boury, M. Calabro, B. D'Andrea, A. McDonald, Solid propulsion for space applications: a roadmap, in: 51st International Astronautical Congress, paper IAA-00-IAA.3.3.02, October 2000]. Moreover, subsequent information Supports its enabling technologies (high strength composite cases. energetic material processing based on continuous mixing, low density insulation, reduced actuator energy requirements, and advanced detailed simulations) and applications (first stages, strap-on, add-ons, small launchers, and niche space applications). Missions currently devoted to solid propulsion and plans for present and future launchers and exploration mission developments in the USA, Japan, and Europe are sketched and targeted improvements, and potential breakthroughs are discussed. (C) 2009 Elsevier Ltd. All rights reserved.
  • Toru Shimada, Nobuhiro Sekino, Mihoko Fukunaga
    JOURNAL OF PROPULSION AND POWER, 25(6) 1300-1310, Nov, 2009  Peer-reviewed
    To understand the mechanism of the generation of large roll torque in an operating solid rocket motor with axially slotted propellant grain and a narrow nozzle-submergence region, fully three-dimensional Navier-Stokes numerical simulations are conducted. Several grain configurations are computed, and it is found that there are at least two groups of quasi-steady-state solutions: one shows large roll torque, and the other shows small roll torque. From the current simulation results, it is observed that large roll torque is generated as a result of the interaction of the circling flow around the nozzle inlet with the slot jet exhausting from the slot end into the aft-end cavity. Although the roll torque evaluated from the computation is one order higher than that observed in real fight, the simulations provide an insight into the qualitative mechanism of real roll-torque generation.
  • Toru Shimada, Hiroshi Hasegawa
    International Journal of Energetic Materials and Chemical Propulsion, 8(2) 147-158, 2009  Peer-reviewed
    In the case of center-perforated composite solid propellant grains, the radial linear burning rate often depends on web location. In many cases, the burning rate of the propellant in the middle of the web is highest along the radial direction. This distribution of the linear burning rate along the radial direction is called a midweb anomaly or hump effect. This phenomenon was researched in the 1980s in depth with many studies disclosed the mechanisms and causes. Recently, the spatial burning rate variation was measured directly with an ultrasonic device. Many studies have explained that oxidizer ammonium perchlorate (AP) particle orientation affects the magnitude of the linear burning rate. In addition, some studies showed that the burning rate anomaly depends on the burning direction. This phenomenon is practically important for the prediction of pressure-time history of a rocket motor with high accuracy. In this study, the midweb anomaly on a small center-perforated motor was investigated. The formulations of the sample propellants were similar to practical propellants. As a result of the motor firing test, pressure hump effect was measured. The burning rate anomaly along the web was estimated by the pressure hump effect and was dependent on the slurry casting process. In order to determine the directivity of the burning rate, it was measured along the motor.
  • Toru Shimada, Masahisa Hanzawa, Takakazu Morita, Takashi Kato, Takashi Yoshikawa, Yasuhiko Wada
    AIAA JOURNAL, 46(4) 947-957, Apr, 2008  Peer-reviewed
    The acoustic combustion instability of a solid rocket motor is investigated by computational fluid dynamics and compared with theoretical results. The quasi-one-dimensional Enter equations for the unsteady flow inside the combustion chamber and the equation for the thermal conduction inside the solid propellant are simultaneously solved with a quasi-steady flame model near the burning surface. The Runge-Kutta discontinuous Galerkin method is used as the platform for the flow simulation, and a numerical accuracy study is carried out. The conventional second-order finite volume method is verified to give accurate results by comparison with the third-order Runge-Kutta discontinuous Galerkin method. The growth rate versus the nozzle entrance Mach number for the attenuation case shows good agreement with the linear theory. For the growing case, it is shown that agreement is good for small Mach numbers. The results of the stability limit show good agreement with the theory for low Mach numbers. For higher Mach numbers, the stability-limit curve of the present simulation shows a dependency on the imaginary part of the response function. Extension to the axisymmetric problem is straightforward, and preliminary results are obtained.
  • Toru Shimada, Hiroto Habu, Yoshio Seike, Seiji Ooya, Hideo Miyachi, Masaaki Ishikawa
    FLOW MEASUREMENT AND INSTRUMENTATION, 18(5-6) 235-240, Oct, 2007  Peer-reviewed
    Simulated solid propellant slurry containing lead sphere tracers is experimentally cast into a double-circular cylindrical container. During the casting, the temperature and the pressure environment has been mimicked to an actual composite solid propellant casting of solid rocket motors. X-rays are projected on to the slurry flow from two directions and penetration images are recorded by a flat-panel detector and an X-ray image intensifier. By suppressing the mutual interference of the X-rays, the two-directional X-ray photography has been successfully conducted. Using the time series data of the X-ray images from the two directions, the identification of each tracer particle in space and time has been done and their three-dimensional paths have been computed. From these procedures, the flow field or the velocity field inside the slurry flow, which is usually invisible, has been estimated. (c) 2007 Elsevier Ltd. All fights reserved.
  • Toru Shimada, Masurni Sekiguchi, Nobuhiro Sekino
    AIAA JOURNAL, 45(6) 1324-1332, Jun, 2007  Peer-reviewed
    Three-dimensional, single-phase (equilibrium two-phase) flows inside a solid rocket motor at three burn-back grain configurations are studied by computational fluid dynamics analyses of the Reynolds-averaged Navier-Stokes equations. The major concern is the relationship between th flowfield and the circumferentially periodic erosion pattern arising in the inlet region of the nozzle, which will be of help for better understanding of the surface recession mechanism. Obtained results for the first two cases show that, because the mass flux of the slot phase is notably large compared with that of the fin phase, a remarkable interphase gap in the amount of convection heating appears either in the throat or the exit cone. The peak heating rate appears, commonly to all cases, azimuthally in the slot phase and axially at the expansion ratio of about 0.9 upstream of the throat. The flow which comes out of a slot into a fin base region spreads toward the fin central part under the influence of the pressure gradient in the circumferential direction and forms a vortical flow tube of opposite rotation mutually with the flow which swirls out of the next slot. At the fin phase, because the proportionality relation is accepted between the total mass recession per unit area and the total convective heat mass transfer per unit area, there is little mechanical erosion, and corrosion is considered to be dominant. On the other band, in the slot phase, surface recession which cannot be explained only by corrosion in a nozzle inlet nose exists. This surface recession has a very high possibility of having occurred by abrasion by the aluminum/alumina particles contained in the How which comes out of the axial slot of grain and collides with the thermal protection system surface. It is expected that the periodic erosion pattern which synchronized with axial slots observed after the static-firing test is the result of such a mechanism ruling. In both the throat and the exit cone, it is thought irrespective of a phase that the effect of mechanical erosion is very small and corrosion or a so-called "chemical attack" is the dominant mechanism of surface recession.

Major Misc.

 254
  • Toru Shimada, Carmine Carmicino, Arif Karabeyoglu
    Aerospace, 9(5), May, 2022  
  • 嶋田徹
    宇宙科学技術連合講演会講演集(CD-ROM), 62nd ROMBUNNO.2S08, 2018  
  • 嶋田徹, 高野忠
    宇宙科学技術連合講演会講演集(CD-ROM), 61st ROMBUNNO.2C17, 2017  
  • 嶋田徹, 北川幸樹, 本江幹朗
    宇宙航空研究開発機構特別資料 JAXA-SP-(Web), (16-003) 113‐114 (WEB ONLY), Sep 30, 2016  
  • Toru Shimada, Tomoaki Usuki
    Proceedings of the International Astronautical Congress, IAC, 2016  
    Results of conceptual study on technology demonstration in flight of a newly proposed hybrid rocket (HR) being enabled mixture-ratio-controlled throttling (MRCT) are described in this paper. The proposed system, named Altering-intensity Swirling-Oxidizer-Flow-Type (A-SOFT) hybrid rocket[1], is essentially-non-explosive and equipped with an MRCT technology. By performing a multi-objective optimization of A-SOFT HR, it has been shown that MRCT is remarkably effective for expanding mission applicability of a sounding rocket[2]. The A-SOFT is realized by independently modulating axial and tangential oxidizer mass flow rates so that both thrust and mixture ratio (O/F) are simultaneously controlled. In most cases, during throttling of a hybrid rocket, O/F varies in accordance with the (1-n)-th power of the oxidizer mass flow rate, where n is usually in the range of 0.5-0.8. So, the propulsion performance deteriorates remarkably in throttling down at lower-than-optimum O/F, or in throttling up at larger-than-optimum O/F, since the specific impulse is usually an upward-convex function of O/F[3]. From launch-system-wise viewpoints, one of the most serious problems caused by O/F shift is the resulting propellant residue[4]. So, MRCT is one of the most-important key technologies for the achievement of high-energy mission, such as a satellite launch, of hybrid rockets in space transportation. Mission requirements for the technology demonstration of MRCT of a hybrid rocket in flight, are to demonstrate 1) capability of designing a compact thrust chamber employing a method of high fuel regression rate, 2) capability of lowering propellant residual and of wide-range thrust control with MRCT technology, and 3) capability of re-ignition in space. During the flight demonstration, for a feedback control of both two quantities being assured, real-time on-board measurements of the fuel web-thickness and of the combustion pressure have to be done.
  • Mikiro Motoe, Toru Shimada
    49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 1 PartF, Sep 16, 2013  
    It has been clarified by experimental investigations that the regression rate can be improved by swirling injection of gaseous oxidizers. Because it has not been enough that the analysis considering characteristics of Swirling-Oxidizer-Flow-Type Hybrid Rocket, it has been hard to mention that internal state of the rocket has been completely cleared. Therefore, in this study, combustion simulation using LES is performed in order to clarify internal state of Swirling-Oxidizer-Flow-Type Hybrid Rocket.
  • P. Tadini, U. Tancredi, M. Grassi, L. Anselmo, C. Pardini, F. Branz, A. Francesconi, F. Maggi, M. Lavagna, L. T. De Luca, N. Viola, S. Chiesa, V. Trushlyakov, T. Shimada
    Proceedings of the International Astronautical Congress, IAC, 3 2319-2328, Jan 1, 2013  
    During the last 40 years, the mass of the artificial objects in orbit increased quite steadily at the rate of about 145 metric tons annually, leading to about 7000 metric tons. Most of the cross-sectional area and mass (97% in low Earth orbit) is concentrated in about 4500 intact abandoned objects plus a further 1000 operational spacecraft. Analyses have shown that the most effective mitigation strategy should focus on the disposal of objects with larger cross-sectional area and mass from densely populated orbits. Recent NASA results have shown that the worldwide adoption of mitigation measures in conjunction with active yearly removal of approximately 0.2-0.5% of the abandoned objects would stabilize the debris population. Targets would have typical masses between 500 and 1000 kg in the case of spacecraft, and of more than 1000 kg for rocket upper stages. In the case of Cosmos-3M second stages, more than one object is located nearly in the same orbital plane. This provides the opportunity of multi-removal missions, more suitable for yearly removal rate and cost reduction needs. This paper deals with the feasibility study of a mission for the active removal of large abandoned objects in low Earth orbit. In particular, a mission is studied in which the removal of two Cosmos-3M second stages, that are numerous in low Earth orbit, is considered. The removal system relies on a Chaser spacecraft which performs rendezvous maneuvers with the two targets. The first Cosmos-3M stage is captured and an autonomous de-orbiting kit, carried by the Chaser, is attached to it. The de-orbiting kit consists of a Hybrid Propulsion Module, which is ignited to perform stage disposal and controlled reentry after Chaser separation. Then, the second Cosmos-3M stage is captured and, in this case, the primary propulsion system of the Chaser is used for the disposal of the mated configuration. Critical mission aspects and related technologies are investigated at a preliminary level. In particular, an innovative electro-adhesive system for target capture, mechanical systems for the hard docking with the target and a hybrid propulsion technology suitable for rendezvous, de-orbiting and controlled reentry operations are analyzed. This is performed on the basis of a preliminary mission profile, in which suitable rendezvous and disposal strategies have been considered and investigated by numerical analysis. A preliminary system mass budget is also performed, showing that the Chaser overall mass is about 1350 kg, including a primary propulsion system of about 300 kg, and a de-orbiting kit with a mass of about 200 kg. The system designed results suitable to be launched with VEGA, actually the cheapest European space launcher.
  • Shota Yamanaka, Toru Shimada
    Proceedings of the International Astronautical Congress, IAC, 9 6967-6988, Jan 1, 2013  
    In this paper, we develop a systematical method for the reduction of chemistry model of hydro-carbon oxygen/air reaction in order to compute the ignition process of boundary layer combustion with a proposed dynamic load balance strategy for the parallel computation of unsteady non-equilibrium chemically reacting flows. Firstly, the reduction method is achieved by omitting Zhu's chemical species determination process, which makes it possible to perform the reduction systematically. By the proposed method the necessary times for chemical reactions of propylene/oxygen and methane/air are reduced to 1/50 and 1/5 of the original each other. Secondly, it is found that, by the dynamic load balance strategy, we can compute the problem 18 times faster than simple load allocation, conventional, approach. Finally, an ignition process of boundary layer combustion of methane/air is calculated by applying the model and computational schemes. We set the initial flow field by using the converged cold-flow solution of air over a methane-injecting porous wall. Injecting high temperature methane gas from a part of the porous wall sets out the ignition simulation. As a result, the first hot spot has appeared at t=0.12 sec near the line of stoichiometry in the boundary layer. Propagation of flame is seen from the hot spot along the line of stoichiometry. The burning speeds are evaluated as 25 and 38 cm/s for the forward moving one and the backward moving one, respectively. They are very close to experimental data (45±5 cm/s). Simulation results also show that the phenomenon occurs under almost constant pressure and enthalpy conditions, and furthermore, the reaction is promoted mainly by the diffusion of radical species. Copyright © 2013 by the International Astronautical Federation. All rights reserved.
  • 船見祐揮, 嶋田徹
    宇宙科学技術連合講演会講演集(CD-ROM), 57th ROMBUNNO.3A12, 2013  
  • 本江幹朗, 嶋田徹
    宇宙科学技術連合講演会講演集(CD-ROM), 57th ROMBUNNO.3A11, 2013  
  • 嶋田徹
    火薬学会年会講演要旨集, 2011 107-110, May 26, 2011  
  • 嶋田徹, 石向桂一, 本江幹朗, 船見祐揮, NOVOZHILOV Vasily
    宇宙科学技術連合講演会講演集(CD-ROM), 55th ROMBUNNO.3B04, 2011  
  • T. Shimada, K. Kitagawa, H. Hasegawa, M. Fukunaga, H. Asakawa
    61st International Astronautical Congress 2010, IAC 2010, 3 2123-2133, Dec 1, 2010  
    This paper describes the development of a numerical simulation system, "Advanced Computer Science on Solid-Rocket-Motor (SRM) Internal Ballistics (ACSSIB)". The objectives of this technology development consist of development of composite-propellant slurry casting-flow simulation, development of local burning-rate correlation with the slurry flow field characteristics, and development of the internal ballistics, i.e., combustion pressure time history, prediction. The ACSSIB have proved itself a promising technology for improvement of SRM reliability and drawn the following conclusions. (1) Hump effect of solid rocket motor combustion is verified by small-scaled motor firing tests and strand burner measurements. (2) Form microscopic observation by microfocus X-ray CT and data deduction by image processing, it is verified that there is a significant correlation between the orientation of coarse AP particles and the burning rate. (3) Development of propellant slurry casting simulation has been successfully conducted. From the casting simulations, it is verified that there is a significant correlation between the angle of the burning direction against the isochrone surface tangent (in plane with the normal) and the burning rate. (4) Development of simulation technique for internal ballistics has been successfully conducted. Simulation results are in good agreement with static firing test results of real motors. Finally, several future technical challenges are identified. Copyright ©2010 by the International Astronautical Federation. All rights reserved.
  • 嶋田徹, 福永美保子, 長谷川宏, 北川幸樹, 淺川弘也, 佐藤航
    宇宙科学技術連合講演会講演集(CD-ROM), 54th ROMBUNNO.1F09, 2010  
  • 嶋田徹
    宇宙科学技術連合講演会講演集(CD-ROM), 54th ROMBUNNO.2B16, 2010  
  • 嶋田徹, 坪井伸幸, 大門優, 関野展弘, 福永美保子, 淺川弘也, 加藤一成, 清家誉志男, 長谷川宏
    航空原動機・宇宙推進講演会講演論文集(CD-ROM), 49th B07, 2009  
  • 嶋田徹
    航空原動機・宇宙推進講演会講演論文集(CD-ROM), 49th B06, 2009  
  • Toru Shimada
    International Astronautical Federation - 59th International Astronautical Congress 2008, IAC 2008, 10 6261-6274, Dec 1, 2008  
    Discussions are made on the localized erosion of charring ablators used in the expanding part of solid rocket motor nozzles. Such erosion pattern is sometimes seen over liner surface downstream the throat inserts after static firing tests. The major characteristic of the localized erosion is that its shape is groove-like, its erosion amount is very large compared to surrounding region, and its location of occurrence is not simply related to the upstream configuration, such as axial slots or fins of the solid propellant grain. The objective here is to consider the growth mechanism of the localized erosion by reviewing facts reported in the literature on the charring ablators, ablation patterns, and vortical three-dimensional flows in nozzles.
  • Toru Shimada, Toru Shimada, Nobuhiro Sekino, Nobuhiro Sekino, Mihoko Fukunaga, Mihoko Fukunaga
    44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Dec 1, 2008  
    In order to understand the mechanism of the large roll-torque generation in the operating solid rocket motor with axially-slotted propellant grain and the narrow nozzle submergence region, fully three-dimensional Navier-Stokes numerical simulations have been conducted. The several grain configurations are computed and it is found that there are at least two groups of quasi-steady state solutions, one shows large roll torque, and the other shows small one. From the present simulation results, it is observed that the large roll torque is generated due to the interaction of the circling flow around the nozzle inlet with the slot jet exhausting out from the slot end into the aft-end cavity. Although the roll torque evaluated from the computation is one-order higher than that observed in the real fight, the present simulation serves the insight into the qualitative mechanism of the real roll torque generation.
  • SHIMADA TORU
    宇宙航空研究開発機構特別資料 JAXA-SP-, (07-023) 5-10, Feb 29, 2008  
  • 嶋田徹, 安田誠一, 鈴木直洋, 冨澤利夫, 二宮一芳, 菊地原清澄, 矢島卓, 尾澤剛
    宇宙科学技術連合講演会講演集(CD-ROM), 52nd ROMBUNNO.1J12, 2008  
  • 嶋田徹
    宇宙科学技術連合講演会講演集(CD-ROM), 52nd ROMBUNNO.2D13, 2008  
  • Toru Shimada, Masahisa Hanzawa, Takakazu Morita, Takashi Kato, Takashi Kato, Takashi Yoshikawa, Takashi Yoshikawa, Yasuhiko Wada, Yasuhiko Wada
    13th AIAA/CEAS Aeroacoustics Conference (28th AIAA Aeroacoustics Conference), Dec 1, 2007  
    The acoustic combustion instability of solid rocket motor (SRM) is investigated by computational fluid dynamics and compared with theoretical results. The quasi-one-dimensional (Q1D) Euler equations for the unsteady flow inside the combustion chamber and the equation for the thermal conduction inside the solid propellant are simultaneously solved with a quasi-steady flame model near the burning surface. The Runge-Kutta Discontinuous Galerkin (RKDG) method is used as the platform for the flow simulation and the numerical accuracy study is carried out. The conventional second-order Finite Volume Method is verified to give accurate results by the comparison with the third-order RKDG method. The growth rate versus the nozzle entrance Mach number for the attenuation case shows good agreement with the linear theory. For the growing case, it is shown that agreement is good for small Mach numbers. The results of the stability limit show good agreement with the theory for low Mach number. For higher Mach numbers, the stability-limit curve of the present simulation show the dependency on the imaginary part of the response function. Extension to the axisymmetric problem is straightforward and preliminary results have been obtained. © 2007 by the authors.
  • Toru Shimada, Nobuhiro Sekino
    International Astronautical Federation - 58th International Astronautical Congress 2007, 9 5712-5720, Dec 1, 2007  
    This paper describes our experiment and computation of roll torque caused by the internal flow of star-perforated solid rocket motor. The roll torque induced by motor internal flow is known from the early days but is not sufficiently understood among rocket scientists in academia and industry. In the background, there is complexity of a three-dimensional vortical flow in combustion chambers. The roll torque occurring in the launch of the Mu-V rocket was reported by the author in the previous paper (Shimada, IAC-06-C4.3.02, Oct.2006), in which the relation with the internal three-dimensional flow was considered. The roll torque was observed in every seven launches during the early operation period of M-14 motor and it was one-order high compared with that of the aerodynamic and/or of thrust misalignment. The cause of the roll torque was discussed on the possibility of Type-I of Knauber's classification, namely the combustion instability, but it was concluded that the possibility of Type-I was small because the mass efflux from the burning surface was relatively large in M-14 and at the same time, no strong sign of combustion instability existed. In this paper, first, the result of a static firing test of a small motor (diameter of 500mm, burning period of 30 seconds, combustion pressure of about 5 MPa, the maximum thrust of about 50 kN, AP/HTPB/Al+MgAl propellant) is described. In this experiment, the swirling component of exhaust plume and the roll torque acting on the motor have been measured. The swirling flow is measured by the lift force acting on the vane which is installed right downstream the nozzle exit. The result shows the swirling has increased for several seconds after the ignition and attenuated gradually after that. On the other hand, roll torque has been evaluated from the balances of the force and the moment among the gravitational force, the suspension force from the test stand, and the two peripheral loads measured at diametrically either side (right and left) of the motor. The results show that the maximum torque has been about 28 N-m at around several seconds after the ignition in the opposite direction of the swirling flow. The evaluated dimensionless torque coefficient is rather a big value of 1.1 × 10-3. Next, discussion is made on whether the roll torque of M-14 is caused by Type-II, i.e., the internal swirling flow due to the grain shape. The M-14 has seven axial slots in each two grain segments. Because the mass efflux from the slots is larger than the remaining parts of the circumference of the cross section, a jet will flow out from each slot into the central port region. At least two possibilities can be considered; one is symmetric and the other is asymmetric secondary flow field in the cross section. It is only the symmetric case that no torque is generated; in which seven pairs of longitudinal vortices should steadily exist. On the other hand, if the symmetric flow is unstable, these jets might merge into one swirling flow which is supposed to be stabler than the symmetric flow. In this paper verification is sought concerning this supposition employing computational fluid dynamics simulations of the three-dimensional internal flow.
  • 嶋田徹, 羽生宏人, 清家誉志男, 大矢清司, 宮地英生, 石川正明
    宇宙航空研究開発機構研究開発報告 JAXA-RR-, 6(06-021) 11P-9, Mar 30, 2007  
    By employing X-ray photography and image analyses, internal three-dimensional flow field of a simulated solid propellant slurry containing lead sphere tracers is visualized in a double circular cylinder container. X-rays are projected on to the slurry flow from two directions perpendicular to each other and penetration images are recorded by a flat panel detector and an X-ray image intensifier. By suppressing the mutual interference of the X-rays, the two-directional X-ray photography has been successfully conducted. Using the time series data of the X-ray images from the two directions, the identification of each tracer particle in space and time has been done and their three-dimensional paths have been computed. From these procedures, the flow field or the velocity field inside the opaque slurry flow has been estimated.
  • Toru Shimada, Hiroto Habu, Yoshio Seike, Seiji Ooya, Hideo Miyachi, Masaaki Ishikawa
    MULTIPHASE FLOW: THE ULTIMATE MEASUREMENT CHALLENGE, PROCEEDINGS, 914 863-+, 2007  
    Simulated solid propellant slurry containing lead sphere tracers is experimentally cast into a double circular cylinder container. During the casting, the temperature and the pressure environment has been mimicked to an actual composite solid propellant casting of solid rocket motors. X-rays are projected on to the slurry flow from two directions and penetration images are recorded by a flat panel detector and an X-ray image intensifier. By suppressing the mutual interference of the X-rays, the two-directional X-ray photography has been successfully conducted. Using the time series data of the X-ray images from the two directions, the identification of each tracer particle in space and time has been done and their three-dimensional paths have been computed. From these procedures, the flow field or the velocity field inside the slurry flow, which is invisible usually, has been estimated.
  • Toru Shimada, Toru Shimada, Toru Shimada, Masumi Sekiguchi, Masumi Sekiguchi, Masumi Sekiguchi, Nobuhiro Sekino, Nobuhiro Sekino, Nobuhiro Sekino
    Collection of Technical Papers - 36th AIAA Fluid Dynamics Conference, 4 2500-2512, Dec 11, 2006  
    Three-dimensional, single-phase (equilibrium two-phase) flows inside a solid rocket motor at three burn-back grain configurations are studied by computational fluid dynamics (CFD) analyses of the Reynolds-averaged Navier-Stokes equations (RANS). The major concern is the relationship between the flow field and the circumferentially periodic erosion pattern arising in the inlet region of the nozzle, which will be of help for the better understanding of the surface recession mechanism. Obtained results for the first two cases show that, since the mass flux of slot phase is notably large compared with that of fin phase, a remarkable inter-phase gap in the amount of convective heating appears either in the throat or the exit cone. The peak heating rate appears, commonly to all cases, azimuthally in the slot phase and axially at the expansion ratio of about 0.9 upstream of the throat. The flow, which comes out of a slot into a fin base region, spreads toward the fin central part under the influence of the pressure gradient in the circumferential direction, and forms vortical flow tube of opposite rotation mutually with the flow which swirls out of the next slot. At fin phase, since proportionality relation is accepted between the total mass recession per unit area and the total convective heat mass transfer per unit area, it is considered that corrosion is dominant ablation mechanism. On the other hand, in slot phase, there exists surface recession which cannot be explained only by corrosion around a nozzle inlet nose. This surface recession has a very high possibility of having occurred by abrasion by the aluminum/alumina particles contained in the flow which comes out of axial slot of grain and collides with the TPS surface. It is expected that periodic erosion pattern which synchronized with axial slots observed after static-firing test is the result of such a mechanism ruling over. In both the throat and the exit cone, it is thought irrespective of a phase that the effect of mechanical erosion is very small and corrosion or so-called "chemical attack" is the dominant mechanism of surface recession.
  • Toru Shimada
    AIAA 57th International Astronautical Congress, IAC 2006, 9 6132-6143, Dec 1, 2006  
    There are unique flow-induced phenomena about solid rocket motors (SRM) whose mechanisms have not been fully understood. The generation of roll torque acting on SRM and peculiar ablation patterns of a nozzle liner surface are taken as examples. By reviewing the open literature, it is found that very few systematic prediction methods exist on these phenomena. Roll torque has been observed during the burning of the first-stage motor of the Mu-V rocket in all six flights since 1997. The cause of the roll torque is sought by evaluating the acoustic effect with mass efflux and combustion response, but sufficiently consistent results have not been obtained. The ablation pattern called striation and cross-hatching has been observed on many specimens in the ablation tests, on reentry , objects after recovery, and on the inner surface of SRM nozzle exit cone. The mechanism of the occurrence of these phenomena is discussed. The existence of the longitudinal vortices is essential for the striation, but for the cross-hatching, whether or not it is an indispensable condition is a pending issue.
  • 嶋田徹
    日本伝熱シンポジウム講演論文集(CD-ROM), 43rd F223, 2006  
  • 嶋田徹, 関野展弘
    航空宇宙技術研究所特別資料 SP-, (57) 154-159, Mar, 2003  
  • 嶋田徹, 関野展弘
    航空原動機・宇宙推進講演会講演集, 43rd 37-42, Jan 30, 2003  
  • 嶋田 徹
    Rocket news, 440 1-1, Apr 1, 2002  
  • 嶋田徹, 山本行光, 広瀬直喜
    航空宇宙技術研究所特別資料 SP-, (41) 123-128, Feb, 1999  
  • Toru Shimada, Yukimitsu Yamamoto, Naoki Hirose, Shuichi Ueda, Katsuhiro Itoh
    AIAA Paper 99-3493, AIAA 33rd Thermophysics Conference, 1999  
    Computations of three-dimensional thermo-chemical non-equilibrium flows around a scale model of the HYFLEX re-entry vehicle have been conducted. Major concern of the simulation is to verify the simulation code by comparison with the measurement data of the HEK shock-tunnel experiments. A modified Equilibrium Flux Method is devised to evaluate the convective terms in an aicurate and stable manner. A non-dimensional parameter is deduced from dimensional analysis to correlate the stagnation-point heating rate with parameters such as the total enthalpy and the binary-scaling parameter. Four cases of free-stream conditions are computed. Computed and measured results are compared on the stagnation-point heating and the heating rate distribution. Computed normalized heat flux distributions do not vary much among the test cases considered. As for stagnation-point heat flux, while computed results show similar tendency to Detra- Kemp-Riddell correlation, they show rather large discrepancy with the experimental data. Both experimental and computational aspects of reasons for the discrepancy have been discussed.
  • 嶋田徹, 山本行光, 広瀬直喜
    航空宇宙技術研究所特別資料 SP-, (37) 133-138, Feb, 1998  
  • Shimada Toru, Yamamoto Yukimitsu, Hirose Naoki
    Special publication of National Aerospace Laboratory : SP, 37 133-138, 1998  
  • Shimada Toru, Sekino Nobuhiro
    Special publication of National Aerospace Laboratory : SP, 34 83-88, 1997  
  • 嶋田徹, 山本行光, 広瀬直喜
    流体力学講演会講演集, 29th 193-196, 1997  
  • 嶋田徹, 関野展弘
    航空宇宙技術研究所特別資料 SP-, (34) 83-88, Jan, 1997  
  • 嶋田徹
    日本機械学会スペース・エンジニアリング・コンファレンス講演論文集, 5th 18(1)-18(6), Jul, 1996  
  • 嶋田徹
    1995年度宇宙航行の力学シンポジウム, 152-155, 1995  
  • 嶋田徹, 田村直樹, 関野展弘, 辻村直久
    日産技報論文集, 1992 50-57, Jun, 1992  
  • 嶋田徹
    流体力学講演会講演集, 24th 207-210, 1992  
  • 嶋田徹, 和田安弘, 古浦勝久
    航空宇宙技術研究所特別資料 SP-, (16) 27-32, Dec, 1991  
  • 嶋田徹
    日本機械学会全国大会講演論文集, 69th(Pt B) 96-98, Oct, 1991  
  • Shimada Toru, Wada Yasuhiro, Koura Katsuhisa
    Special publication of National Aerospace Laboratory : SP, 16 27-32, 1991  
  • 嶋田徹
    流体力学講演会講演集, 22nd 2-5, 1990  
  • 嶋田徹
    日産技報論文集, 1989 188, May, 1989  
  • 嶋田徹
    6(2) 26-31, 1989  Peer-reviewed
  • SHIMADA Toru
    ISAS Report, Institute of Space and Astronautical Science, 629 1-12, 1988  
    Transient aerodynamic characteristics of the flows around bodies of parachute-like configuration are numerically analysed from solution of the Navier-Stokes equations. The computational method is mainly based upon combination of effective and efficient techniques recently developed in the field of computational fluid mechanics. The results show that the flow behavior around a mouth plays a key role in determining the maximum peak drag acting of the parachute-like body in the starting period from the rest and also a vent is effective in controlling the starting peak of the drag.
  • 嶋田徹
    宇宙ステーション講演会講演集, 4th 103-104, 1988  
  • 嶋田徹, 安部隆士
    航空宇宙技術研究所特別資料 SP-, (8) 109-114, Nov, 1987  
  • Shimada Toru, Abe Takashi
    Special publication of National Aerospace Laboratory : SP, 8 109-114, 1987  
  • IAF-87-298,38th Congress of the International Astronautical Federation, 1987  
  • 嶋田徹, 川崎和憲
    宇宙科学技術連合講演会講演集, 30th 480-481, Oct, 1986  
  • 嶋田徹, 川崎和憲
    宇宙科学技術連合講演会講演集, 30th 478-479, Oct, 1986  
  • 嶋田徹, 小口はく郎
    流体力学講演会講演集, 16th 258-261, 1984  

Major Books and Other Publications

 9

Major Presentations

 211
  • Toru Shimada
    18th International Conference on Flow Dynamics, Oct 28, 2021
  • Toru Shimada
    2nd Hybrid Rocket Symposium, Jul 11, 2019  Invited
  • SHIMADA TOORU
    火薬学会年会講演要旨集, May 26, 2011
  • T. Shimada, K. Kitagawa, H. Hasegawa, M. Fukunaga, H. Asakawa
    61st International Astronautical Congress 2010, IAC 2010, Sep 27, 2010
    This paper describes the development of a numerical simulation system, "Advanced Computer Science on Solid-Rocket-Motor (SRM) Internal Ballistics (ACSSIB)". The objectives of this technology development consist of development of composite-propellant slurry casting-flow simulation, development of local burning-rate correlation with the slurry flow field characteristics, and development of the internal ballistics, i.e., combustion pressure time history, prediction. The ACSSIB have proved itself a promising technology for improvement of SRM reliability and drawn the following conclusions. (1) Hump effect of solid rocket motor combustion is verified by small-scaled motor firing tests and strand burner measurements. (2) Form microscopic observation by microfocus X-ray CT and data deduction by image processing, it is verified that there is a significant correlation between the orientation of coarse AP particles and the burning rate. (3) Development of propellant slurry casting simulation has been successfully conducted. From the casting simulations, it is verified that there is a significant correlation between the angle of the burning direction against the isochrone surface tangent (in plane with the normal) and the burning rate. (4) Development of simulation technique for internal ballistics has been successfully conducted. Simulation results are in good agreement with static firing test results of real motors. Finally, several future technical challenges are identified. Copyright ©2010 by the International Astronautical Federation. All rights reserved.
  • SHIMADA TOORU
    宇宙科学技術連合講演会講演集(CD-ROM), 2010
  • SHIMADA TOORU, FUKUNAGA MIHOKO, HASEGAWA HIROSHI, KITAGAWA KOKI, ASAKAWA HIROYA, SATO WATARU
    宇宙科学技術連合講演会講演集(CD-ROM), 2010
  • SHIMADA TOORU
    航空原動機・宇宙推進講演会講演集(CD-ROM), 2009
  • SHIMADA TOORU, TSUBOI NOBUYUKI, DAIMON YU, SEKINO NOBUHIRO, FUKUNAGA MIHOKO, ASAKAWA HIROYA, KATO KAZUSHIGE, SEIKE YOSHIO, HASEGAWA HIROSHI
    航空原動機・宇宙推進講演会講演集(CD-ROM), 2009
  • Toru Shimada
    International Astronautical Federation - 59th International Astronautical Congress 2008, IAC 2008, Sep 29, 2008
    Discussions are made on the localized erosion of charring ablators used in the expanding part of solid rocket motor nozzles. Such erosion pattern is sometimes seen over liner surface downstream the throat inserts after static firing tests. The major characteristic of the localized erosion is that its shape is groove-like, its erosion amount is very large compared to surrounding region, and its location of occurrence is not simply related to the upstream configuration, such as axial slots or fins of the solid propellant grain. The objective here is to consider the growth mechanism of the localized erosion by reviewing facts reported in the literature on the charring ablators, ablation patterns, and vortical three-dimensional flows in nozzles.
  • Toru Shimada, Toru Shimada, Nobuhiro Sekino, Nobuhiro Sekino, Mihoko Fukunaga, Mihoko Fukunaga
    44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Jul 21, 2008
    In order to understand the mechanism of the large roll-torque generation in the operating solid rocket motor with axially-slotted propellant grain and the narrow nozzle submergence region, fully three-dimensional Navier-Stokes numerical simulations have been conducted. The several grain configurations are computed and it is found that there are at least two groups of quasi-steady state solutions, one shows large roll torque, and the other shows small one. From the present simulation results, it is observed that the large roll torque is generated due to the interaction of the circling flow around the nozzle inlet with the slot jet exhausting out from the slot end into the aft-end cavity. Although the roll torque evaluated from the computation is one-order higher than that observed in the real fight, the present simulation serves the insight into the qualitative mechanism of the real roll torque generation.
  • SHIMADA TOORU
    宇宙科学技術連合講演会講演集(CD-ROM), 2008
  • SHIMADA TOORU, YASUDA SEIICHI, SUZUKI NAOHIRO, TOMIZAWA TOSHIO, NINOMIYA KAZUYOSHI, KIKUCHIHARA KIYOZUMI, YAJIMA TAKASHI, OZAWA TAKESHI
    宇宙科学技術連合講演会講演集(CD-ROM), 2008
  • Toru Shimada, Masahisa Hanzawa, Takakazu Morita, Takashi Kato, Takashi Kato, Takashi Yoshikawa, Takashi Yoshikawa, Yasuhiko Wada, Yasuhiko Wada
    13th AIAA/CEAS Aeroacoustics Conference (28th AIAA Aeroacoustics Conference), Dec 1, 2007
    The acoustic combustion instability of solid rocket motor (SRM) is investigated by computational fluid dynamics and compared with theoretical results. The quasi-one-dimensional (Q1D) Euler equations for the unsteady flow inside the combustion chamber and the equation for the thermal conduction inside the solid propellant are simultaneously solved with a quasi-steady flame model near the burning surface. The Runge-Kutta Discontinuous Galerkin (RKDG) method is used as the platform for the flow simulation and the numerical accuracy study is carried out. The conventional second-order Finite Volume Method is verified to give accurate results by the comparison with the third-order RKDG method. The growth rate versus the nozzle entrance Mach number for the attenuation case shows good agreement with the linear theory. For the growing case, it is shown that agreement is good for small Mach numbers. The results of the stability limit show good agreement with the theory for low Mach number. For higher Mach numbers, the stability-limit curve of the present simulation show the dependency on the imaginary part of the response function. Extension to the axisymmetric problem is straightforward and preliminary results have been obtained. © 2007 by the authors.
  • Toru Shimada, Nobuhiro Sekino
    International Astronautical Federation - 58th International Astronautical Congress 2007, Dec 1, 2007
    This paper describes our experiment and computation of roll torque caused by the internal flow of star-perforated solid rocket motor. The roll torque induced by motor internal flow is known from the early days but is not sufficiently understood among rocket scientists in academia and industry. In the background, there is complexity of a three-dimensional vortical flow in combustion chambers. The roll torque occurring in the launch of the Mu-V rocket was reported by the author in the previous paper (Shimada, IAC-06-C4.3.02, Oct.2006), in which the relation with the internal three-dimensional flow was considered. The roll torque was observed in every seven launches during the early operation period of M-14 motor and it was one-order high compared with that of the aerodynamic and/or of thrust misalignment. The cause of the roll torque was discussed on the possibility of Type-I of Knauber's classification, namely the combustion instability, but it was concluded that the possibility of Type-I was small because the mass efflux from the burning surface was relatively large in M-14 and at the same time, no strong sign of combustion instability existed. In this paper, first, the result of a static firing test of a small motor (diameter of 500mm, burning period of 30 seconds, combustion pressure of about 5 MPa, the maximum thrust of about 50 kN, AP/HTPB/Al+MgAl propellant) is described. In this experiment, the swirling component of exhaust plume and the roll torque acting on the motor have been measured. The swirling flow is measured by the lift force acting on the vane which is installed right downstream the nozzle exit. The result shows the swirling has increased for several seconds after the ignition and attenuated gradually after that. On the other hand, roll torque has been evaluated from the balances of the force and the moment among the gravitational force, the suspension force from the test stand, and the two peripheral loads measured at diametrically either side (right and left) of the motor. The results show that the maximum torque has been about 28 N-m at around several seconds after the ignition in the opposite direction of the swirling flow. The evaluated dimensionless torque coefficient is rather a big value of 1.1 × 10-3. Next, discussion is made on whether the roll torque of M-14 is caused by Type-II, i.e., the internal swirling flow due to the grain shape. The M-14 has seven axial slots in each two grain segments. Because the mass efflux from the slots is larger than the remaining parts of the circumference of the cross section, a jet will flow out from each slot into the central port region. At least two possibilities can be considered; one is symmetric and the other is asymmetric secondary flow field in the cross section. It is only the symmetric case that no torque is generated; in which seven pairs of longitudinal vortices should steadily exist. On the other hand, if the symmetric flow is unstable, these jets might merge into one swirling flow which is supposed to be stabler than the symmetric flow. In this paper verification is sought concerning this supposition employing computational fluid dynamics simulations of the three-dimensional internal flow.
  • Toru Shimada, Hiroto Habu, Yoshio Seike, Seiji Ooya, Hideo Miyachi, Masaaki Ishikawa
    AIP Conference Proceedings, Sep 24, 2007
    Simulated solid propellant slurry containing lead sphere tracers is experimentally cast into a double circular cylinder container. During the casting, the temperature and the pressure environment has been mimicked to an actual composite solid propellant casting of solid rocket motors. X-rays are projected on to the slurry flow from two directions and penetration images are recorded by a flat panel detector and an X-ray image intensifier. By suppressing the mutual interference of the X-rays, the two-directional X-ray photography has been successfully conducted. Using the time series data of the X-ray images from the two directions, the identification of each tracer particle in space and time has been done and their three-dimensional paths have been computed. From these procedures, the flow field or the velocity field inside the slurry flow, which is invisible usually, has been estimated. © 2007 American Institute of Physics.
  • SHIMADA TOORU, HABU HIROTO, SEIKE YOSHIO, OYA SEIJI, MIYACHI HIDEO, ISHIKAWA MASAAKI
    宇宙航空研究開発機構研究開発報告 JAXA-RR-, Mar 30, 2007
    By employing X-ray photography and image analyses, internal three-dimensional flow field of a simulated solid propellant slurry containing lead sphere tracers is visualized in a double circular cylinder container. X-rays are projected on to the slurry flow from two directions perpendicular to each other and penetration images are recorded by a flat panel detector and an X-ray image intensifier. By suppressing the mutual interference of the X-rays, the two-directional X-ray photography has been successfully conducted. Using the time series data of the X-ray images from the two directions, the identification of each tracer particle in space and time has been done and their three-dimensional paths have been computed. From these procedures, the flow field or the velocity field inside the opaque slurry flow has been estimated.
  • Toru Shimada, Toru Shimada, Toru Shimada, Masumi Sekiguchi, Masumi Sekiguchi, Masumi Sekiguchi, Nobuhiro Sekino, Nobuhiro Sekino, Nobuhiro Sekino
    Collection of Technical Papers - 36th AIAA Fluid Dynamics Conference, Dec 11, 2006
    Three-dimensional, single-phase (equilibrium two-phase) flows inside a solid rocket motor at three burn-back grain configurations are studied by computational fluid dynamics (CFD) analyses of the Reynolds-averaged Navier-Stokes equations (RANS). The major concern is the relationship between the flow field and the circumferentially periodic erosion pattern arising in the inlet region of the nozzle, which will be of help for the better understanding of the surface recession mechanism. Obtained results for the first two cases show that, since the mass flux of slot phase is notably large compared with that of fin phase, a remarkable inter-phase gap in the amount of convective heating appears either in the throat or the exit cone. The peak heating rate appears, commonly to all cases, azimuthally in the slot phase and axially at the expansion ratio of about 0.9 upstream of the throat. The flow, which comes out of a slot into a fin base region, spreads toward the fin central part under the influence of the pressure gradient in the circumferential direction, and forms vortical flow tube of opposite rotation mutually with the flow which swirls out of the next slot. At fin phase, since proportionality relation is accepted between the total mass recession per unit area and the total convective heat mass transfer per unit area, it is considered that corrosion is dominant ablation mechanism. On the other hand, in slot phase, there exists surface recession which cannot be explained only by corrosion around a nozzle inlet nose. This surface recession has a very high possibility of having occurred by abrasion by the aluminum/alumina particles contained in the flow which comes out of axial slot of grain and collides with the TPS surface. It is expected that periodic erosion pattern which synchronized with axial slots observed after static-firing test is the result of such a mechanism ruling over. In both the throat and the exit cone, it is thought irrespective of a phase that the effect of mechanical erosion is very small and corrosion or so-called "chemical attack" is the dominant mechanism of surface recession.
  • Toru Shimada
    AIAA 57th International Astronautical Congress, IAC 2006, Dec 1, 2006
    There are unique flow-induced phenomena about solid rocket motors (SRM) whose mechanisms have not been fully understood. The generation of roll torque acting on SRM and peculiar ablation patterns of a nozzle liner surface are taken as examples. By reviewing the open literature, it is found that very few systematic prediction methods exist on these phenomena. Roll torque has been observed during the burning of the first-stage motor of the Mu-V rocket in all six flights since 1997. The cause of the roll torque is sought by evaluating the acoustic effect with mass efflux and combustion response, but sufficiently consistent results have not been obtained. The ablation pattern called striation and cross-hatching has been observed on many specimens in the ablation tests, on reentry , objects after recovery, and on the inner surface of SRM nozzle exit cone. The mechanism of the occurrence of these phenomena is discussed. The existence of the longitudinal vortices is essential for the striation, but for the cross-hatching, whether or not it is an indispensable condition is a pending issue.
  • SHIMADA TOORU
    日本伝熱シンポジウム講演論文集(CD-ROM), 2006
  • SHIMADA TOORU, SEKINO NOBUHIRO
    航空宇宙技術研究所特別資料 SP-, Mar, 2003
  • SHIMADA TOORU, SEKINO NOBUHIRO
    航空原動機・宇宙推進講演会講演集, Jan 30, 2003
  • 宇宙輸送シンポジウム講演集, 宇宙科学研究所, 2001
  • 宇宙輸送シンポジウム講演集, 宇宙科学研究所, 2001
  • 第102回月例講演会, 宇宙科学研究所, 2001
  • 宇宙輸送シンポジウム講演集, 宇宙科学研究所, 2001
  • SHIMADA TOORU, YAMAMOTO YUKIMITSU, HIROSE NAOKI
    航空宇宙技術研究所特別資料 SP-, Feb, 1999
  • 使える最先端流動解析とその応用事例-デモ展示付-、日本機械学会関西支部第238回講習会教材, 1999
  • 共著
    航空宇宙数値シミュレーション技術シンポジウム’99論文集, 1999
  • SHIMADA TOORU, YAMAMOTO YUKIMITSU, HIROSE NAOKI
    航空宇宙技術研究所特別資料 SP-, Feb, 1998
  • Shimada Toru, Yamamoto Yukimitsu, Hirose Naoki
    Special publication of National Aerospace Laboratory : SP, 1998
  • SHIMADA TOORU, YAMAMOTO YUKIMITSU, HIROSE NAOKI
    流体力学講演会講演集, 1997
  • SHIMADA TOORU, SEKINO NOBUHIRO
    航空宇宙技術研究所特別資料 SP-, Jan, 1997
  • SHIMADA AKIRA, WADA YASUHIRO, KOURA KATSUHISA
    航空宇宙技術研究所特別資料 SP-, Dec, 1991
  • 日本機械学会第69期全国大会講演会講演論文集, 1991
  • SHIMADA AKIRA, KAWASAKI KAZUNORI
    宇宙科学技術連合講演会講演集, Oct, 1986
  • SHIMADA AKIRA, KAWASAKI KAZUNORI
    宇宙科学技術連合講演会講演集, Oct, 1986
  • Toru Shimada, Hakuro Oguchi
    流体力学講演会講演集, 1984

Major Research Projects

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