“蚌式”进气道附面层扫除特性风洞试验

Boundary layer blowing characteristics of bump inlet by wind-tunnel test

以某低、高速风洞为平台,设计搭建了"蚌式"进气道附面层扫除特性测量试验系统,进行了不同流量系数和来流马赫数下进气道鼓包表面附面层扫除特性的风洞试验,通过对试验数据的整理、计算和对比分析同型号的飞行试验结果,研究了"蚌式"进气道鼓包表面附面层扫除特性。研究结果表明:在相同的来流马赫数下,随着流量系数的增大,鼓包表面附面层的扫除能力逐渐减弱;在亚声速工况的绝大多数流量范围内,鼓包表面压力系数沿鼓包中心线对称分布、压力梯度变化明显,且在不同截面沿主流方向具有增大的特征,鼓包构型对附面层扫除效果较强;超声速工况下具有明显附面层扫除能力的流量范围明显小于亚声速工况,进气道唇口形成的弓形激波是影响鼓包表面不同位置压力梯度变化的主要因素,进而决定着附面层扫除特性。在接近来流马赫数1.8及以上飞行工况下,附面层的扫除能力减弱,附面层分离加强,进而会造成较大的进气压力损失和畸变。

Based on the low and high speed wind tunnel,a test system for measuring the boundary layer blowing characteristics of the bump inlet at a wind tunnel was designed and set up. The boundary layer blowing characteristics were studied by test at different incoming Mach numbers and flow ratios. Through analysis of the test data, calculation and comparison of the flight test results of the same type, the boundary layer blowing characteristics of the bump inlet were studied. Results showed that the boundary layer blowing ability of the bump inlet gradually decreased with the increase of the flow ratio at the same incoming Mach number.Within the wide range of the flow ratio at subsonic, the pressure coefficient of the bump inlet surface was symmetric along the center line with high pressure gradient, and it increased along the mainstream. The bulge structure showed a relatively strong blowing ability of the boundary layer. The flow ratio range with effective blowing ability at supersonic was significantly smaller than that at subsonic. At supersonic the bow-shaped shock wave of the inlet lip was the main factor affecting the pressure gradient change at different positions on the bump surface and the boundary layer blowing ability. Under the flight conditions near Mach number 1.8 and above, the boundary layer blowing ability was decreased, and the separation of the boundary layer flowing was strengthened, causing large loss and distortion of the inlet pressure.