激波风洞7°尖锥边界层转捩实验研究

高超声速边界层转捩对摩阻、传热等有重要影响，飞行器的研制迫切希望能精确预测和控制边界层转捩。在JF8A激波风洞中开展了7°半锥角的高超声速尖锥边界层转捩实验研究，利用响应频率达到1 MHz量级的高频压力传感器对尖锥壁面脉动压力进行了测量，并结合热流测量结果，研究了高超声速尖锥边界层中扰动波的发展过程。实验结果表明:JF8A激波风洞在雷诺数为6.4×106/m状态下核心流的自由流噪声为2.8%；高频脉动压力测量技术能清晰地捕捉转捩过程中的第二模态波及其发展历程，试验状态下模型的第二模态波频率范围为165~206 kHz。当前研究结果能够为高超声速数值方法验证提供数据支撑。 

Experimental study of boundary-layer transition on a 7° sharp cone in shock tunnel

Hypersonic boundary layer transition has an important influence on friction drag and heat transfer, and thus accurate prediction and control of boundary layer transition are critical to the development of hypersonic vehicles. In this paper, experimental study on a 7° half-angle sharp cone was conducted in the JF8A shock tunnel to investigate the hypersonic boundary layer transition. The wall fluctuation pressure was measured by the pressure transducer with the response frequency as high as 1 MHz, and the development process of the disturbance wave in hypersonic sharp cone boundary layer was also investigated together with the results of heat flux measurement. The experimental results show that the pressure fluctuation in the free flow is 2.8% under the test condition of Reynolds number equals to 6.4×106/m. Structures and development of the second mode waves in the process of transition can be obtained by the high-frequency fluctuation pressure measurement technique. The characteristic frequency of the second mode wave changes from 165 kHz to 206 kHz under the present test condition. The current research results can provide data support for hypersonic numerical method validation.