Study of Combustion Characteristics of Hybrid Rocket Engine with Swirling Liquid Oxygen Flow and its Application to Small Sounding Rocket Engine
| Project/Area Number |
14350513
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| Research Category |
Grant-in-Aid for Scientific Research (B)
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| Allocation Type | Single-year Grants |
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| Section | 一般 |
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| Research Field |
Aerospace engineering
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| Research Institution | Tokyo Metropolitan Institute of Technology |
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Principal Investigator |
YUASA Saburo Tokyo Metropolitan Institute of Technology, Faculty of Engineering, Professor, 工学部, 教授 (60123147)
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| Co-Investigator(Kenkyū-buntansha) |
YUASA Saburo Tokyo Metropolitan Institute of Technology, Faculty of Engineering, Professor (60123147)
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| Project Period (FY) |
2002 – 2004
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| Project Status |
Completed (Fiscal Year 2004)
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| Budget Amount *help |
¥14,400,000 (Direct Cost: ¥14,400,000)
Fiscal Year 2004: ¥5,200,000 (Direct Cost: ¥5,200,000)
Fiscal Year 2003: ¥2,100,000 (Direct Cost: ¥2,100,000)
Fiscal Year 2002: ¥7,100,000 (Direct Cost: ¥7,100,000)
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| Keywords | Hybrid Rocket / Liquid Oxygen / Swirling Flow / PMMA / Regenerative-cooling Nozzle / エンジン性能 / 旋回型液体酸素インジェクタ |
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| Research Abstract |
By applying swirl to an oxidizer flow, our hybrid rocket engine with GO_2 and PMMA could obtain the high performance. In 2001, we launched the first hybrid rocket of 1.8m in length, which recorded the first successful flight of a hybrid rocket, in Japan. As the next stage, a hybrid rocket engine system with a swirling LOX flow was constructed. In order to investigate combustion processes of directly exposed PMMA to LOX, tests of a small hybrid combustor with a single port filled with LOX or GOX without swirling were carried out. It was found that a LOX layer appeared on the wall and combustion processes of LOX/PMMA are determined by evaporation of LOX.
Experiments of combustion in a hybrid rocket engine with a swirling LOX flow showed that fuel regression rate, C^* efficiency and Isp of the hybrid rocket engine with the swirling LOX flow were smaller than those with the swirling GOX flow, resulting in lower performance of the hybrid rocket engine with LOX than that with GOX. The existen
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ce of a LOX layer on the combustion chamber wall decreased the combustion characteristics due to the loss of the angular momentum, the decrease of the net Go and the longer combustion time. Combustion oscillation called a "Chugging" mode also occurred.
To improve the poor combustion performance, a regenerative-cooling nozzle of the hybrid rocket engine using LOX was adopted. Thermal analysis of the nozzle was performed, and a longitudinal channel-wall type regenerative-cooling nozzle with LOX was made. Vaporization experiments with LOX as the coolant, in which the mass flow rates of the oxygen used for combustion of the hybrid rocket engine and LOX passing through the nozzle were independently supplied to examine vaporization conditions of LOX, were conducted. It was confirmed that LOX was safely vaporized through the nozzle.
A 1500-N thrust hybrid rocket engine was made. The burning test using GO2 was carried out and showed good performance. The future issue to achieve this project purpose is to match the regenerative-cooling LOX nozzle with the swirling oxygen type hybrid rocket engine for a sounding rocket. Less
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Report
(4 results)
Research Products
(40 results)
