| 马赫数2.5超/亚声流流向涡掺混超燃室冷态实验 |
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| 引用本文： | 介克,杨占宇,赵吕顺,刘建,王洪铭,黄勇,单鹏.马赫数2.5超/亚声流流向涡掺混超燃室冷态实验[J].航空动力学报,2007,22(12):2012-2017. |
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| 作者姓名： | 介克 杨占宇 赵吕顺 刘建 王洪铭 黄勇 单鹏 |
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| 作者单位： | 北京航空航天大学,能源与动力工程学院,北京,100083 |
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| 基金项目： | 国家自然科学基金
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国家自然科学基金重大研究计划子课题 |
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| 摘 要： | 结合对应的数值计算,对一种带亚声速预燃室和流向涡掺混器的超声速燃烧模型燃烧室实验台,在其进口马赫数为2.5的来流条件下,进行了冷态流场的实验研究.实验测得其壁面静压分布和激波系结构与流场的CFD计算结果基本一致.实验结果表明,模型燃烧室全流场超声速,达到设计状态.马赫数2.5下的冷态实验数据和CFD计算数据为进行点火实验提供了依据.
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| 关 键 词： | 航空、航天推进系统 超声速燃烧 预燃室 波瓣掺混器 流向涡掺混 |
| 文章编号： | 1000-8055(2007)12-2012-06 |
| 收稿时间： | 2006/11/7 0:00:00 |
| 修稿时间： | 2006年11月7日 |
Cold flowfield experiment study of an inlet Mach number 2.5 scramjet combustor with a stream-wise vortices mixer between supersonic/subsonic-pilot-combustion flow |
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| JIE Ke,YANG Zhan-yu,ZHAO Lü-shun,LIU Jian,WANG Hong-ming,HUANG Yong,SHAN Peng.Cold flowfield experiment study of an inlet Mach number 2.5 scramjet combustor with a stream-wise vortices mixer between supersonic/subsonic-pilot-combustion flow[J].Journal of Aerospace Power,2007,22(12):2012-2017. |
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| Authors: | JIE Ke YANG Zhan-yu ZHAO Lü-shun LIU Jian WANG Hong-ming HUANG Yong SHAN Peng |
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| Institution: | School of Jet Propulsion, Beijing University of Aeronautics and Astronautics, Beijing 100083, China;School of Jet Propulsion, Beijing University of Aeronautics and Astronautics, Beijing 100083, China;School of Jet Propulsion, Beijing University of Aeronautics and Astronautics, Beijing 100083, China;School of Jet Propulsion, Beijing University of Aeronautics and Astronautics, Beijing 100083, China;School of Jet Propulsion, Beijing University of Aeronautics and Astronautics, Beijing 100083, China;School of Jet Propulsion, Beijing University of Aeronautics and Astronautics, Beijing 100083, China;School of Jet Propulsion, Beijing University of Aeronautics and Astronautics, Beijing 100083, China |
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| Abstract: | An experimental study of cold flow field was conducted under entrance Mach number 2.5 on a supersonic combustor experiment rig with subsonic pilot combustor and stream-wise vortices mixer.The experiment result suggests that the up-wall static pressure distribution of wall surface and shock wave structure coincide with the CFD calculation results,and the full flow field supersonic combustor reaches the design state.The cold test data and CFD calculation data under 2.5Mach provide a basis for ignition test. |
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| Keywords: | aerospace propulsion system supersonic combustion pilot-combustor lobed-mixer stream-wise vortices mixing |
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