一种超声速颌下进气道变几何方案研究

针对宽范围定几何颌下进气道高马赫数下的压缩量不足问题，提出了一种喉部滑块前后移动的变几何调节方案，该方案通过滑块前后移动改变高低马赫数下的喉道尺寸，使进气道能够满足高低马赫数下的压缩量要求。本文提出了两种滑块布局方式，针对内锥侧滑块布局方式，按调节原理进行了滑块型面与进气道内流道型面的匹配设计，并将变几何颌下进气道与定几何方案进行了性能比较。数值研究表明：按Ma2.5-Ma4.0范围设计的变几何颌下进气道，在设计点，临界状态出口总压恢复系数为0.51，较公开文献中定几何方案提高8.5%；在Ma4.0，0°攻角工况下，临界状态出口总压恢复系数为0.46，提高12.2%；在Ma2.7，1°攻角工况下流量系数为0.69, 临界状态出口总压恢复系数为0.78。气动性能表明，该颌下进气道性能优越，调节方案简单可行。

Variable Geometry Scheme of Supersonic Chin Inlet

Regarding the insufficient compression at high Mach number for a fixed geometry chin inlet with a wide working range, a variable geometry chin inlet scheme with a throat slider moving forth and back is presented. The throat size can be changed by moving the slider so that the inlet can meet the compression requirements at high and low Mach numbers. Two kinds of slider layout are presented in this paper. As for the internal cone side layout, the matching design between slider and internal channel profile is conducted according to the adjustment principle. The performance of variable geometry chin inlet is compared with that of fixed geometry scheme. The numerical simulation results show that the exit critical total pressure recovery coefficient of the variable geometry chin inlet designed with the working range from Ma2.5 to Ma4.0 is 0.51 and 0.46 at the design point and Ma4.0, 0 degree angle of attack respectively, which is 8.5% and 12.2% higher than that of the fixed geometry scheme in the literature respectively. The flow coefficient is 0.69 and the critical total pressure recovery coefficient is 0.78 at Ma2.7 with 1 degree angle of attack. The aerodynamic result shows that the adjustment scheme of chin inlet is simple and feasible with good performance.