马赫数2.5超/亚声流流向涡掺混超燃室冷态实验

基金项目: 国家自然科学基金(50076003);国家自然科学基金重大研究计划子课题(90305010)

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

School of Jet Propulsion, Beijing University of Aeronautics and Astronautics, Beijing 100083, China

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.
- 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
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 CFDcalculation results,and the full flow field supersonic combustor reaches the design state.The cold test data and CFDcalculation data under 2.5Mach provide a basis for ignition test.
- aerospace propulsion system /
- supersonic combustion /
- pilot-combustor /
- lobed-mixer /
- stream-wise vortices mixing

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