流体力学与飞行力学

激波风洞边界层转捩测量技术及应用

  • 李强 ,
  • 江涛 ,
  • 陈苏宇 ,
  • 常雨 ,
  • 赵磊 ,
  • 张扣立
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  • 中国空气动力研究与发展中心 超高速空气动力研究所, 绵阳 621000

收稿日期: 2018-10-19

  修回日期: 2018-11-11

  网络出版日期: 2019-02-15

基金资助

国家重点研发计划(2016YFA0401201)

Measurement technique and application of boundary layer transition in shock tunnel

  • LI Qiang ,
  • JIANG Tao ,
  • CHEN Suyu ,
  • CHANG Yu ,
  • ZHAO Lei ,
  • ZHANG Kouli
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  • Hypervelocity Aerodynamics Institute, China Aerodynamics Research and Development Center, Mianyang 621000, China

Received date: 2018-10-19

  Revised date: 2018-11-11

  Online published: 2019-02-15

Supported by

National Key Research and Development Program of China(2016YFA0401201)

摘要

高超声速边界层转捩对摩阻、传热等有重要影响。在高超声速飞行器研制中,迫切希望能精确预测和控制边界层转捩。激波风洞作为高超声速气动热环境试验的主要地面模拟设备,是研究高超声速边界层转捩的重要设备。但激波风洞原有测量技术适用于工程型号试验,需要依据高超声速边界层转捩特点进行适应性改造和升级。依据高超声速边界层转捩过程中的热流、压力、密度等物理参数变化,发展了薄膜热流传感器测热技术、温敏热图测量技术、高频脉动压力测量技术、高清晰度纹影显示技术等适用于激波风洞的边界层转捩测量技术。并针对头部钝度0.05 mm的半锥角7°尖锥模型,在中国空气动力研究与发展中心Ø2 m激波风洞(FD-14A)马赫数10、单位雷诺数1.2×107/m的流场条件下开展了边界层转捩试验。采用多种转捩测量技术同时进行测量,获得尖锥模型表面边界层转捩情况、边界层脉动压力频谱特征、边界层内清晰的第2模态波和湍流斑纹影图像,不同测量技术获取的试验结果可相互印证,线性稳定性理论分析结果与试验结果相吻合。

本文引用格式

李强 , 江涛 , 陈苏宇 , 常雨 , 赵磊 , 张扣立 . 激波风洞边界层转捩测量技术及应用[J]. 航空学报, 2019 , 40(8) : 122740 -122740 . DOI: 10.7527/S1000-6893.2019.22740

Abstract

Hypersonic boundary layer transition has an important influence on friction drag and heat transfer. The development of hypersonic vehicles is expected to accurately predict and control boundary layer transition. As the main ground equipment for hypersonic aerodynamic thermal environment test, the shock tunnel is an important equipment for studying hypersonic boundary layer transition. However, the original measurement technology of the tunnel is designed for engineering experiments, and it needs to be adapted and upgraded according to the characteristics of hypersonic boundary layer transition in shock tunnel. Therefore, according to the changes in physical parameters such as heat flux pressure density during the hypersonic boundary layer transition process, we develop the thin film heat flux sensor technique, the temperature-sensitive paint technique, the high frequency fluctuation pressure measurement, and the high definition schlieren visualization technique. The boundary layer transition experiment is carried out in China Aerodynamics Research and Development Center Ø2 m shock tunnel (FD-14A). The model is a 7° half-angle cone model with 0.05 mm nosetip bluntness, 10 nominal Mach number, the unit Reynolds number of 1.2×107/m. Simultaneous measurements are carried out by using all the above techniques, obtaining good experimental results of boundary layer transition condition, fluctuation pressure spectrum characteristics, clear schlieren photos of the second mode wave and turbulent spots. The results obtained by these different measurement techniques can cross reference each other, and the theoretical analyses of linear stability are consistent with the experimental results.

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