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对高超声速飞行器表面凸起附近的气流流动和气动加热开展了实验研究和分析。实验在高超声速炮风洞中进行,来流马赫数为8.2、单位雷诺数为9.35×106 m-1。利用薄膜传热测量方法进行了凸起几何形状和边界层状态对干扰流动加热的影响评估。利用流油图谱和纹影摄像法得到了凸起周围的流动特征:若凸起上游边界层未分离,最大峰值热流发生在凸起侧方附近处;若凸起上游边界层完全分离,最大峰值热流通常发生在凸起的上游表面。实验发现最大峰值热流和来流边界层状态关系不大,原因是流动干扰区表现出较强的三维扰动特性,使得来流层流边界层在干扰区内会转变成过渡甚至完全湍流状态。
The experimental study and analysis of the airflow and aerodynamic heating near the surface bump of hypersonic vehicles are carried out. Experiments were performed in a hypersonic gun tunnel with a Mach-flow of 8.2 and a unit Reynolds number of 9.35 × 106 m-1. The influence of the convex geometry and the boundary layer state on the disturbance flow heating was evaluated by the thin film heat transfer measurement method. The flow characteristics around the bump are obtained by using the flow chart and the schlieren method: if the upper boundary layer is not separated, the maximum peak heat flow occurs near the side of the bump; if the upper boundary layer is completely separated, the maximum peak Heat flow usually occurs on the raised upstream surface. The experimental results show that the relationship between the maximum peak heat flux and the boundary layer state is insignificant. The reason is that the three-dimensional disturbances show strong three-dimensional perturbation characteristics, which makes the transitional boundary layer transition to even complete turbulence in the disturbance zone.