一种鼻锥钝化高超声速轴对称进气道流动特性实验

doi:10.7527/S1000-6893.2014.0259

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一种鼻锥钝化高超声速轴对称进气道流动特性实验

高文智, 李祝飞, 杨基明

中国科学技术大学近代力学系, 合肥 230027

收稿日期:2014-07-31 修回日期:2014-09-15 出版日期:2015-01-15 发布日期:2014-09-19
通讯作者: 杨基明,Tel.: 0551-63603390 E-mail: jmyang@ustc.edu.cn E-mail:jmyang@ustc.edu.cn
作者简介:高文智男, 博士研究生。主要研究方向: 高超声速进气道。 Tel: 0551-63603390 E-mail: ensureme@mail.ustc.edu.cn;李祝飞男, 博士。主要研究方向: 高超声速进气道。 Tel: 0551-63601242 E-mail: lizhufei@ustc.edu.cn;杨基明男, 博士, 教授, 博士生导师。主要研究方向: 高超声速空气动力学, 实验流体力学。 Tel: 0551-63603390 E-mail: jmyang@ustc.edu.cn
基金资助:
国家自然科学基金 (11132010)

Flow characteristics experiments of a hypersonic axisymmetric inlet with nose bluntness

GAO Wenzhi, LI Zhufei, YANG Jiming

Department of Modern Mechanics, University of Science and Technology of China, Hefei 230027, China

Received:2014-07-31 Revised:2014-09-15 Online:2015-01-15 Published:2014-09-19
Supported by:
National Natural Science Foundation of China (11132010)

摘要/Abstract

摘要：

前缘钝化尺度是高超声速进气道设计中的关键参数。针对一种前体锥加弯曲压缩面的高超声速轴对称进气道,选取最大尺度为3.2 mm (5%唇缘半径)的几种典型鼻锥钝化半径,在马赫数Ma=6来流,及模型安装攻角为0°、4°、7°的条件下开展鼻锥钝化尺度对进气道流动性能影响的实验研究。采用纹影拍摄及压力测量记录各来流条件下进气道前体流场结构及壁面压强分布,并在无攻角来流条件下利用微型扰流器进行边界层强制转捩研究。结果表明,对无攻角来流而言,即使是尺度高达3.2 mm的钝化半径对进气道前体流场结构及壁面静压分布也基本没有影响。此来流条件下,几种不同鼻锥钝化半径的前体压缩面均出现小范围流动分离,而添加扰流器后该分离区均消失。钝化尺度的影响随着攻角的增加而显现,尽管不同鼻锥钝化尺度下迎风面流场及壁面压强分布几乎没有差别,但背风面随钝化尺度增大表现为边界层明显增厚、流动趋于不稳定。其中最大钝化尺度R=3.2 mm的构型在4°攻角来流时背风面即出现明显的分离区,而7°攻角来流时背风面更是出现大范围流动分离、进气道背风侧不起动,并导致进气道内部壁面压强显著下降。

关键词: 高超声速流, 轴对称进气道, 鼻锥钝化, 攻角来流, 流动分离

Abstract:

Blunt scale of leading edge is a key parameter in the design of hypersonic inlet. Flow characteristics of a hypersonic axisymmetric inlet, of which forebody compression surfaces consisted of a cone and curved surfaces, are studied experimentally at Ma=6 with nose blunt scales up to 3.2 mm (5% cowl lip radius). High speed schlieren imaging of external flow field and centerbody pressure distribution are recorded during experiments, with the model installing angles of attack of 0°, 4° and 7°. Forced transition tests are also explored with trips at angle of attack of 0°. It is shown that the variations of the forebody flowfield and pressure distribution are negligible within 3.2 mm nose radius for the horizontal freestream, while small separation regions exist around the inlet entrance. The tripped cases show obvious suppression of the flow separation, validating successful transition dominated by trips. Obvious discrepancies of nose effects have been found between windward and leeward sides of the axisymmetric inlet under the freestream angle of attack. Variations of windward flowfield and pressure distribution can be hardly noticed for current runs, while slip lines of leeward side move outward and leeward flowfields turn to be unstable with increasing nose scale. For the largest 3.2 mm nose radius, evident separation appears on the leeward side at angle of attack of 4°, corresponding to surface pressure rise in the separation region. The separation becomes more severe when the angle of attack increases and the leeward side of the inlet turns to be unstart at angle of attack of 7°, which results in remarkable pressure drop.

Key words: hypersonic flow, axisymmetric inlet, nose bluntness, freestream angle of attack, flow separation

中图分类号:

高文智, 李祝飞, 杨基明. 一种鼻锥钝化高超声速轴对称进气道流动特性实验[J]. 航空学报, 2015, 36(1): 302-310.

GAO Wenzhi, LI Zhufei, YANG Jiming. Flow characteristics experiments of a hypersonic axisymmetric inlet with nose bluntness[J]. ACTA AERONAUTICAET ASTRONAUTICA SINICA, 2015, 36(1): 302-310.

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E-mail：hkxb@buaa.edu.cn

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一种鼻锥钝化高超声速轴对称进气道流动特性实验

Flow characteristics experiments of a hypersonic axisymmetric inlet with nose bluntness

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