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

Kohei OZAWA Toru SHIMADA

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.13, no.4, pp.JFST0031, 2018 (Released:2018-11-09)

参考文献数

31

被引用文献数

7

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The characteristics of several O/F control methods for hybrid rocket propulsion have been discussed and theoretically analyzed from the physical properties of propellants and fuel regression behavior. In this research, comparisons have been made among different oxidizer injection methods of Altering-intensity Swirling Oxidizer Flow Type (A-SOFT), Aft-chamber Oxidizer Injection Method (AOIM), and Swirling-AOIM for the throttle range with a constant O/F, design restrictions of the fuel grain, penalties on the adoption of the methods, and suitable scales of the engine. Theoretical analysis on regression rates has revealed that A-SOFT has upper and lower limits of throttle while maintaining a constant O/F whereas AOIM does not have any lower limit, and Swirling-AOIM covers both the throttle ranges. The designing restriction of the fuel grain derived from the regression rate behavior has indicated that A-SOFT using paraffin and oxygen has a potential to maintain 50-100% throttle range over a burn. The penalties for the adoption of these O/F control methods have also been discussed from the aspects of the increase in the complexity of the system, structural mass, and pressure drop at the injector for the methods using gaseous injection. The pressure drop has quantitatively been evaluated by relating the available swirl strength with the cross-sectional area and gaseous oxidizer mass flux at the injector. This analysis has revealed 5 times difference in the available swirl strength between the gaseous oxygen and the decomposed gas of 90% hydrogen peroxide. The sizing of the 1st stage of the satellite launcher has revealed that A-SOFT and Swirling-AOIM are suitable for small-scale engines with a propellant mass of 100-102 [ton] using paraffin and liquid oxygen whereas AOIM and Swirling-AOIM are suitable for engines with paraffin and 90% hydrogen peroxide.

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.