著者

Masahiro KANAZAKI Kai TOMISAWA Koji FUJITA Akira OYAMA Hiroki NAGAI

出版者

The Japan Society of Mechanical Engineers

雑誌

Journal of Fluid Science and Technology (ISSN:18805558)

巻号頁・発行日

vol.14, no.3, pp.JFST0017, 2019 (Released:2019-12-19)

参考文献数

11

被引用文献数

2

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We redesigned the Mars Airplane Balloon Experiment Two (MABE-2) based on MABE-1 to improve the vehicle’s stability and controllability. Following the redesign, the MABE-2 vehicle had a larger horizontal tail volume than that of MABE-1 for improved stability performance. In addition, to further improve the stability and control characteristics, a rectangular planform was employed for the horizontal tail wing; in contrast, MABE-1 had a tapered planform. The vertical tail position of MABE-2 was moved to the end of the horizontal tail wing, because the vertical tail of MABE-1, which was positioned at the mid span of the horizontal tail wing, showed aerodynamic interaction with the horizontal tail wing. In this paper, we discussed the aerodynamic performance of a control surface based on computational fluid dynamics with variation in the deflection angle between the control surface and the horizontal tail (elevator), and we examined the effects of this redesign on longitudinal control characteristics. Numerical investigations confirmed the linear variation in the pitching moment and the aerodynamic force with the changing elevator deflection angle in MABE-2. Surface pressure observations indicated that MABE-2 shows a smooth variation in the pressure distribution with changing elevator deflection angle, while MABE-1 does not. These results demonstrate that the aerodynamic control characteristics of MABE-2 were improved in comparison to those of MABE-1.

著者

Masahiro KANAZAKI Shoma ITO Fumio KANAMORI Masaki NAKAMIYA Koki KITAGAWA Toru SHIMADA

出版者

一般社団法人 日本機械学会

雑誌

Journal of Fluid Science and Technology (ISSN:18805558)

巻号頁・発行日

vol.11, no.1, pp.JFST0003-JFST0003, 2016 (Released:2016-02-17)

参考文献数

8

被引用文献数

1

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This paper reports on the conceptual design of a three-stage launch vehicle (LV) with a clustered hybrid rocket engine (HRE) through multi-disciplinary design optimization. This LV is a space transportation concept that can deliver micro-satellites to sun-synchronous orbits (SSOs). To design a high-performance LV with HRE, the optimum size of each component, such as an oxidizer tank containing liquid oxidizer, a combustion chamber containing solid fuel, a pressurizing tank, and a nozzle, should be determined. In this study, paraffin (FT-0070) is used as a propellant for the HRE, and three cases are compared: In the first case, HREs are optimized for each stage. In the second case, HREs are optimized together for the first and second stages but separately for the third stage. In the third case, HREs are optimized together for each stage. The optimization results show that the performance of the design case that uses the same HREs in all stages is 40% reduced compared with the design case that uses optimized HREs for each stage.

著者

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

出版者

一般社団法人 日本機械学会

雑誌

Journal of Fluid Science and Technology (ISSN:18805558)

巻号頁・発行日

vol.9, no.5, pp.JFST0071-JFST0071, 2014 (Released:2014-11-28)

参考文献数

6

被引用文献数

2

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The subject of this paper is to improve on parameterization for conceptual design method of three stage hybrid rocket. Multi-Objective Genetic Algorithm (MOGA) is employed to solve multi-disciplinary design exploration of a three-stage launch vehicle concept using a hybrid rocket engine. MOGA which is used as the optimization methods for multi-objective problems utilizes real-number cording and the Pareto ranking method. According to our previous study, the propulsive performance of MOGA's solution was as low as the lower limit of design space. The design space of a conceptual three-stage launch vehicle hybrid rocket engine was reconsidered based on the results of multi-disciplinary design optimization. The design variables of the nozzles were reconsidered by exploring the design space. Specifically, the nozzle expansion ratio was considered as the ratio of the nozzle exit radius to the body radius. In this way, there are no solutions which violate the design constraints about the geometric condition of the nozzle exit. Consequently, the new conceptual design method can effectively explore solutions which have higher propulsive performance than previous method. As the result, the combustion chamber pressure is increased in the first stage. In the second stage, the solutions which are explored, modified parameterization are shown larger thrust level than previously.