研究者詳細 - 北川　幸樹

論文 - 北川　幸樹

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Static burning tests on a bread board model of altering-intenisty swirling-oxidizer-flow-type hybrid rocket engine 査読有り

Ozawa K., Usuki T., Mishima G., Kitagawa K., Yamashita M., Mizuchi M., Katakami K., Maji Y., Aso S., Tani Y., Wada Y., Shimada T.

52nd AIAA/SAE/ASEE Joint Propulsion Conference, 2016 2016年01月

記述言語：英語掲載種別：研究論文（国際会議プロシーディングス）

A bread board model of Altering-number Swirling Flow Type (A-SOFT) hybrid rocket engine has been newly developed and the static burning tests have been conducted. A-SOFT hybrid rocket engines have controllability of both thrust and O/F without any replacement of components in the engine. This advantage is acquired by controlling oxidizer mass flow rate and effective swirl number. The purpose of this set of experiments is to confirm the continuity, monotonousness and predictability of the performances of A-SOFTs. The A-SOFT BBM showed a favorable fuel regression behavior. The fuel regression averaged along spatial direction fit the shape of regression rate function proposed before the experiments within ±3.5% errors, and the fuel regression rates are continuous and monotonous along swirl number and oxidizer mass flux. The O/F and thrust data also respectively fit the prediction formulas within ±4.2% and ±4.7% errors. c* efficiency is evaluated with the Isp efficiency - nozzle efficiency ratio ηIsp/ ηCF in order to compensate the pressure sensing errors caused by the centrifugal forces of swirling flows. Though ηIsp/ ηCF in the cases of weak swirl injection was clearly larger than in the cases of axial injection, its dependence on effective geometric swirl number was not clear in strong swirl conditions.

Scopus

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Effect of oxidizer particle orientation on burning rates of composite propellants 査読有り

Hasegawa H., Fukunaga M., Kitagawa K., Shimada T.

International Journal of Energetic Materials and Chemical Propulsion 15 ( 4 ) 285 - 304 2016年01月

記述言語：英語掲載種別：研究論文（学術雑誌）

Regarding center-perforated composite solid propellant grains, the radial linear burning rate of the propellant often depends on its location in the web. In most cases, the burning rate at the middle of the web is highest along the radial direction. This deviation of the linear burning rate along the radial direction is called the midweb anomaly, the hump effect, and so on. The phenomenon is the result of multiple factors (e.g., orientation of oxidizer particles, existence of binder layers, and the effect of those factors on combustion and heat conduction in the solid phase). The physical cause of this phenomenon has not been understood in sufficient detail. In this study, the effect of oxidizer particle orientation on the burning rate is studied. The particle orientation in inert solid propellant grains is observed by micro-focus X-ray computerized tomography. To simplify the indefinite shape of actual ammonium perchlorate particles, simple cylindrical inert particles were mixed into inert propellant slurry. As a result the particle orientations were observed in center-perforated and solid cylindrical inert grains. The orientation of the particles seemed to be along the isochronous surface, which is formed during the casting process. One physical correlation between the particle orientation and the mean or local burning rate is investigated by numerical simulation from the point of view of the combustion surface configuration, which is formed by the different burning rates of the propellant ingredients. The results of the simulation suggest that the orientation of oxidizer particles in line with the burning direction increase the mean burning rate by forming and developing uneven surfaces.

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2016014195

Scopus

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Development of test facilities for 5 kn-thrust hybrid rocket engines and a swirling-oxidizer-flow-type hybrid rocket engine for technology demonstration 査読有り

Kitagawa K., Yuasa S., Sakurai T., Hatagaki S., Shiraishi N., Ando H., Yagishita T., Suzuki N., Takayama A., Yui R., Shimada T.

International Journal of Energetic Materials and Chemical Propulsion 15 ( 6 ) 435 - 451 2016年01月

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（学術雑誌）

At the Hybrid Rocket research Working Group, performing demonstrations of current technologies is pursued. The target thrust was set as 5 kN and a hybrid rocket test engine program (HTE-5-1) was initiated. To demonstrate the hybrid rocket engine technology, a hybrid rocket test facility was designed at the Akiruno facility, which belongs to the Japan Aerospace Exploration Agency. A low-flow rate gaseous oxygen (GOX) feed system that includes a purge system, test stand, measurement system, and control system was manufactured. The operability and functionality, such as the GOX mass flow range and natural frequency of the test stand, were examined. It was confirmed that combustion tests under a condition of up to 0.5 kg/s GOX mass flow rate could be conducted. A 5-kN-thrust swirling-oxidizer-flow-type hybrid rocket engine using a GOX/polypropylene propellant was designed and manufactured as a demonstrator. Two combustion tests were performed with the demonstration engine under conditions below the design point. Reliable ignition and stable combustion for long durations were achieved, and accurate experimental data were obtained. This demonstrated the effectiveness of the combustion test and validated the suitability of our engine design method. It was confirmed that the combustion test could be performed safely and properly using the new test facility and demonstrator.

DOI： 10.1615/IntJEnergeticMaterialsChemProp.2017011994

Scopus

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Development of solid propulsion system for enhanced epsilon launch vehicle and M-35 static firing test 査読有り

Kitagawa K., Tokudome S., Hori K., Tanno H., Nakano N.

Proceedings of the International Astronautical Congress, IAC 2016年01月

担当区分：筆頭著者記述言語：英語掲載種別：研究論文（国際会議プロシーディングス）

The Epsilon launch vehicle, the newest version of Japan's solid propulsion rocket, made its maiden ight in September of 2013. The purpose of the Epsilon rocket is to provide small satellites with responsive launching with low-cost, user-friendly and efficient launch system. Now that the rst ight was successfully nished, JAXA has been conducting intensive researches on a next generation Epsilon to launch a more powerful and lower cost version of Epsilon (Evolved Epsilon). In order to minimize technical risks and to keep up with demand of future payloads, JAXA plans to take a step-by-step approach toward Evolved Epsilon. As the first upgrade toward Evolved Epsilon, JAXA has started the development of Enhanced Epsilon. The Enhanced Epsilon is required to enhance launch capability by ERG satellite mission which have decided to change an orbit to be put into to farther and also to improve on-board capability in size and in weight by ASNARO2 satellite mission. The development of Enhanced Epsilon is mainly the renewal of the second stage, and also includes the each subsystem's improvement. The main change of the solid propulsion system is exposure of the second motor M-35. Currently, the design has been finished. The outside diameter of the motor case is expanded into approximately 2.5 m in order to increase the amount of the solid propellant and the outer shell of the motor case is used as the outer shell of the launch vehicle. Solid propellant which can the high-performance equal to a conventional upper-stage motor developed newly, reducing the cost. A general front ignition system is adopted instead of the rear ignition system of the throw-away type which was adopted for the previous motor. A new development material is applied to the case lining. An expansible nozzle is not adopted because compatibility of high-performance and cost reduction. In order to verify the design, the static firing test of the second motor M-35 on condition of vacuum has been conducted. This paper describes overview of development of the solid propulsion system for Enhanced Epsilon and the results of the M-35 static firing test.

Scopus

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Ascendancy of extinction-reignition on single-stage hybrid sounding rocket in view of fuels 査読有り

Chiba K., Yoda H., Ito S., Kanazaki M., Watanabe S., Kitagawa K., Shimada T.

57th AIAA/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference 2016年01月

記述言語：英語掲載種別：研究論文（国際会議プロシーディングス）

A single-stage launch vehicle with hybrid rocket engine has been conceptually designed by using design informatics, which has three points of view, i.e., problem definition, optimization, and data mining. The primary objective of the present design is that the downrange and the duration time in the lower thermosphere are sufficiently secured for the aurora scientific observation, whereas the initial gross weight is held down to the extent possible. The multidisciplinary design optimization was performed by using a hybrid evolutionary computation. Data mining was also implemented by using a scatter plot matrix. Polypropylene and liquid oxygen with swirling flow are adopted as solid fuel and liquid oxidizer, respectively. The condition of two-time ignitions is assumed in fight sequence on the equation of motion for the three degree of freedom rigid body. Consequently, the design information regarding the tradeoffs, the behaviors of the design variables in the design space to become the nondominated solutions, and the implication of the design variables for the objective functions have been obtained quantitatively. The structurization and visualization of the design space has been implemented in order to observe the effectiveness of the local regions of each design variable. The advantage of extinction-reignition has been indicated.

DOI： 10.2514/6.2016-0583

Scopus

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Design Optimization of Single-Stage Launch Vehicle Using Hybrid Rocket Engine,” International Journal of Aerospace System Engineering 査読有り国際誌

Kanazaki, M., Ariyarit, A., Yoda, H., Ito, S., Chiba, K., Kitagawa, K., Shimada, T.

International Journal of Aerospace System Engineering 2 ( 2 ) 29 - 33 2015年12月

Evolutionary algorithm applied to ballistic launch vehicle design using hybrid rocket engine evaluated by enhanced flight simulation 査読有り

Yoda H., Ito S., Kanazaki M., Chiba K., Kitagawa K., Shimada T.

2015 IEEE Congress on Evolutionary Computation, CEC 2015 - Proceedings 618 - 625 2015年09月

記述言語：英語掲載種別：研究論文（国際会議プロシーディングス）

A multi-objective genetic algorithm (MOGA) has been applied to the multidisciplinary design optimization (MDO) of a launch vehicle (LV) with a hybrid rocket engine (HRE) to investigate the ability of an HRE to serve as a sounding rocket from various perspectives. In this study, the flight evaluation was enhanced to 3-degree-of-freedom (3DoF) in order to consider the equations of motion for horizontal and vertical motion and rotation of the LV. In the consideration of the rotation of the LV, the time variation of the center of gravity due to the fuel burn was estimated. The non-dominated sorting genetic algorithm-II (NSGA-II) was used to solve multi-objective problems (MoPs). Four design problems were examined in order to understand the practical physics of hybrid rocket. As a result, tradeoff information was observed for all design problems. The results for the present four design problems indicate that economical performance of LV is limited with the HRE in terms of the maximum altitude and maximum downrange distances achievable. The hypervolume, which was used as the metric to evaluate the difficulty of the design problems, reveals that the convergence of the solutions for not only altitude maximization in the case of a vertical launch but also the maximization of downrange at higher target altitudes was affected by the severe limitation. To observe the dependence of the design problems on the constraint, the design problems were visualized using a colored parallel coordinates plot (PCP), and the LV geometries determined from the nondominated solutions were successfully examined.

DOI： 10.1109/CEC.2015.7256948

Scopus

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Multidisciplinary design exploration for sounding launch vehicle using hybrid rocket engine in view of ballistic performance 査読有り

Chiba K., Kanazaki M., Ariyarit A., Yoda H., Ito S., Kitagawa K., Shimada T.

International Journal of Turbo and Jet Engines 32 ( 3 ) 299 - 304 2015年09月

記述言語：英語掲載種別：研究論文（国際会議プロシーディングス）

Conceptual design of a launch vehicle with a hybrid rocket engine (HRE) has been implemented using design informatics approach in order to investigate the feasibility of a single-stage hybrid rocket. Two test design problems were formulated by using two objective functions: maximization of downrange and minimization of initial gross weight, seven design variables which describe geometry and initial conditions, and one constraint relative to target altitude. The optimization result reveals the economical performance of hybrid rocket is limited with HRE in terms of the maximum downrange achievable. Moreover, the data-mining result indicates the mechanism of design-variable behavior.

DOI： 10.1515/tjj-2015-0038

Scopus

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Numerical analysis of multi-parallelized swirling flow inside a circular pipe 査読有り

Takayama A., Kitagawa K., Shimada T.

Journal of Mechanical Science and Technology 29 ( 3 ) 951 - 962 2015年03月

担当区分：責任著者記述言語：英語掲載種別：研究論文（学術雑誌）

The flow field of multi-parallelized swirling flow inside a circular pipe was investigated numerically. Two types of swirling flow configuration are considered. One type is the co-rotating type. Four co-rotating swirls are arranged at the vertex position of square in this type. The other type is the counter-rotating type which consists of two pairs of swirls having opposite swirl rotations. Each pair is arranged diagonally at the vertex position of a square. By coupling the discrete vortex method and boundary element method, unsteady flow simulation is performed. Swirl modeling with vortex elements is used in this simulation and its validity is confirmed. From the simulation results, in the co-rotating type, the four swirls interact and their shape is deformed. Each vortex motion vanishes rapidly in the downstream region. Finally, they are turned into a single swirling flow. In counter-rotating type, each vortex motion is maintained a little bit longer than co-rotating type, and their shape is not so deformed. However, the flow patterns are changed completely in the downstream region. The swirling velocity of each swirl mostly vanishes. Finally, they are turned into an axial flow. For the investigation of the mixing promoting effect due to parallelizing swirls, particle tracking simulations are performed in the co-rotating type and the counter-rotating type. As a comparison, the simulation for single swirl flow is also performed. In these simulations, the particles are introduced in the vicinity of pipe inner wall. In addition, the assumption that particles follow the flow motion absolutely is used. From the results, the motion of particles in these three cases is completely different. For the co-rotating and counter-rotating type, the particle entrainment into the main axial flow is clearly observed. This indicates the mixing is improved compared to single swirl flow. The difference of particle entrainment motion between co-rotating and counter-rotating type is slight.

DOI： 10.1007/s12206-015-0209-8

Scopus

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Conceptual design of single-stage launch vehicle with hybrid rocket engine using design informatics 査読有り

Chiba K., Kanazaki M., Kitagawa K., Shimada T.

Computational Methods in Applied Sciences 36 369 - 384 2015年01月

記述言語：英語掲載種別：研究論文（国際会議プロシーディングス）

A single-stage launch vehicle with hybrid rocket engine has been conceptually designed by using design informatics, which has three points of view as problem definition, optimization, and data mining. The primary objective of the design in the present study is that the sufficient down range and the duration time in the lower thermosphere are achieved for aurora scientific observation whereas the initial gross weight is held down. Multidisciplinary design optimization and data mining were performed by using evolutionary hybrid computation under the conditions that polypropylene as solid fuel and liquid oxygen as liquid oxidizer were adopted and that single-time ignition is implemented in sequence. Consequently, the design information regarding the tradeoffs and the behaviors of the design variables in the design space was obtained in order to quantitatively differentiate the advantage of hybrid rocket engine.

DOI： 10.1007/978-3-319-11541-2_24

Scopus

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多目的進化計算による多数回燃焼を行うハイブリッドロケットの性能評価査読有り

金崎雅博, 千葉一永, 北川幸樹, 嶋田徹

進化計算学会論文誌 ( 進化計算学会 ) 6 ( 3 ) 137 - 145 2015年01月

記述言語：日本語掲載種別：研究論文（学術雑誌）

With the multi-combustion technology, the combustion in a hybrid rocket engine (HRE) can be temporarily stopped via oxidizer throttling. In this paper, two types of HREs, one with multi-combustion technology and the other without, are compared to investigate the effects of multi-combustion on the flight performance of launch vehicles (LVs). Non-dominated Sorting Genetic Algorithm-II (NSGA-II) which was a multi-objective evolutionary algorithm (MOEA) was applied to solve the design problems using real-number coding and the Pareto ranking method. To investigate the effects of the multi-combustion on flight performance of LV with HRE, three design problems were considered. The first case was the maximization of the flight altitude and the minimization of the gross weight. The second case was the minimization of the maximum acceleration and the minimization of the gross weight. The final case was the maximization of the flight downrange and the minimization of the gross weight. Many non-dominated solutions were obtained by NSGA-II, and a trade-off was observed between the two objective functions in each case. MOEA results were visualized using a parallel coordinate plot. According to the exploration result, it was found that the multi-combustion of HRE was effective to reduce the maximum acceleration. Such ability could be expected to reduce the shock load to payloads carried by the LV with HRE.

DOI： 10.11394/tjpnsec.6.137

CiNii Article

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Performance calculations and burning tests on altering-intensity swirling oxidizer flow type hybrid rocket engines 査読有り

Ozawa K., Kitagawa K., Shimada T.

Proceedings of the International Astronautical Congress, IAC 9 7332 - 7345 2015年01月

記述言語：英語掲載種別：研究論文（国際会議プロシーディングス）

As a technique to eliminate O/F shifts, Altering-intensity Swirling Oxidizer Flow Type (A-SOFT) hybrid rocket is proposed using the combination axial and tangential oxidizer injection. In this paper, the possible geometric design point of the motor is considered about 2-tons class single stage A-SOFT hybrid rockets, and the nominal performance increase by eliminating O/F shifts is evaluated. The highest altitude of the A-SOFT is 450 [km] and the averaged ISP is 284[s] with 100[kg] payload, and this performance is several percent higher than the ones of S-520 with 95[kg] payload, Japanese 2-tons class solid sounding rocket. Though A-SOFTs increase the nominal flight performance by 1% from Swirling Oxidizer Flow Type hybrid rockets, which cannot O/F shift, their feedback control prevents the large performance losses caused by fuel regression errors and residuals, and this function is not found in the conventional hybrid rockets. In this paper, a planning of steady state burning tests of A-SOFT bread board model is also explained. The purpose of the burning tests are to demonstrate the concept of A-SOFTs. The test motor is 250[N] scale and uses polypropylene and GOX and its burning duration is 5[s]. In more than 10 times steady state tests, 50% throttling conditions with effective geometric swirl intensity between 0 and 37.3 are inculded.

Scopus

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Conceptual Design of Single-stage Rocket Using Hybrid Rocket by Means of Genetic Algorithm 査読有り国際誌

Kanazaki M., Ariyairt A., Chiba K., Kitagawa K., Shimada T.

Procedia Engineering ( Elsevier ) 99 198 - 207 2015年01月

記述言語：英語掲載種別：研究論文（国際会議プロシーディングス）

In this study, a multi-objective genetic algorithm (MOGA) was applied to the multidisciplinary design optimization (MDO) of a hybrid rocket. A swirling-oxidizer-type hybrid rocket engine (HRE) with a single cylindrical grain port was designed. It was considered that this HRE could temporarily stop combustion via oxidizer throttling; this feature is called multi-combustion. The MOGA was applied to solve the multi-objective problem using real-number coding and the Pareto ranking method. In this study, three design problems were considered. First problem was the maximization of the flight altitude and minimization of the gross weight. Second problem was the minimization of the maximum acceleration and minimization of the gross weight. Third problem was the maximization of the duration time over the target flight altitude and minimization of the gross weight. Each objective function was empirically estimated. In addition, this study compared two types of HREs to investigate the emects of the multi-combustion: one type was able to carry out the multi-combustion, and the other was not. Many non-dominated solutions were obtained using the MOGA, and a trade-off was observed between the two objective functions. To understand the design problem, the MOGA results were visualized using a parallel coordinate plot (PCP).

DOI： 10.1016/j.proeng.2014.12.526

Scopus

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Conceptual design: Dependence of parameterization on design performance of three-stage hybrid rocket 査読有り国際誌

Masahiro KANAZAKI, Fumio KANAMORI, Yosuke KITAGAWA, Masaki NAKAMIYA, Koki KITAGAWA, Toru SHIMADA

Journal of Fluid Science and Technology ( The Japan Society of Mechanical Engineers ) 9 ( 5 ) JFST0071 2014年11月

Structurization of Design Space for Launch Vehicle with Hybrid Rocket Engine Using Stratum-Type Association Analysis 国際誌

Chiba K., Kanazaki M., Watanabe S., Kitagawa K., Shimada T.

Adaptation, Learning and Optimization ( Springer ) 1 509 - 521 2014年11月

設計情報学を用いたハイブリッドロケットエンジン搭載単段式宇宙輸送機の概念設計査読有り

千葉一永, 渡邉真也, 金崎雅博, 北川幸樹, 嶋田徹

日本機械学会論文集 ( 日本機械学会 ) 80 ( 818 ) TRANS0287 2014年10月

Diversity of Design Knowledge for Launch Vehicle in View of Fuels on Hybrid Rocket Engine 査読有り国際誌

Kazuhisa CHIBA, Masahiro KANAZAKI, Masaki NAKAMIYA, Koki KITAGAWA, Toru SHIMADA

Journal of Advanced Mechanical Design, Systems, and Manufacturing ( The Japan Society of Mechanical Engineers ) 8 ( 3 ) JAMDSM0023 - JAMDSM0023 2014年07月

設計情報学による燃料種を考慮した科学観測用単段式ハイブリッドロケットの概念設計査読有り

千葉一永, 金崎雅博, 中宮賢樹, 北川幸樹, 嶋田徹

航空宇宙技術 ( 日本航空宇宙学会 ) 13 41 - 50 2014年06月

Experimental Study of Fragmentation of Hybrid Rocket Fuel 査読有り国際誌

Koki KITAGAWA, Yoshio NAKAYAMA, Tomoharu MATSUMURA, Kunihiko WAKABAYASHI, Ryo TODA, Takakazu MORITA, Toru SHIMADA

TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN ( 日本航空宇宙学会 ) 12 ( ists29 ) Pa_15 - Pa_20 2014年05月

A Linear Stability Analysis of Oscillatory Combustion Induced by Combustion Time Delays of Liquid Oxidizer in Hybrid Rocket Motors 査読有り国際誌

Takakazu Morita, Koki Kitagawa, Toru Shimada, Shigeru Yamaguchi

International Journal of Energetic Materials and Chemical Propulsion ( Begell House ) 13 ( 1 ) 83 - 96 2014年01月

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