STUDIES OF RECEPTIVITY AND LAMINAR-FLOW-CONTROL PROBLEMS FOR THE LEADING-EDGE-REGION BOUNDARY LAYER OF THREE-DIMENSIONAL WING

如何在超音速巡航时实现L/D比大于10的高摩阻比，是下一代超音速运输机面临的技术挑战。为此，有人建议对机翼边界层采用层流控制(LFC)。为了实现这种LFC，我们必须对在前缘区域增长的超音速边界层的感受性和不稳定性有更多的基础知识。这正是我们进行数值和实验研究的动机。通过对超音速感受性和不稳定性问题的研究，我们得到了以下结果。数值计算表明，超音速边界层(自由流马赫数为2.2)的感受性很弱，TS波的激发振幅比相应的不可压缩情况低1/10。通过对马赫数为2.5马赫的后掠翼附着线附近附面层的扰动，数值计算表明，横流不稳定型流向涡的增长速度最快。对于马赫数为2.5的平板边界层，受来自壁孔的强周期射流的扰动，数值结果表明发夹涡在低速边界层中的发展类似于低速边界层，表明存在非线性亚临界转变。我们进行了仔细的实验，以阐明加速度如何抑制亚临界转变，发现亚临界转变的临界雷诺数(基于动量厚度)从平板流动的120增加到低速边界层的300，该边界层具有无量纲有利压力梯度-4.5×10^&lt；-3；指出后掠翼前缘区除了横流不稳定外，还存在流线曲率引起的不稳定。

It is a technological challenge posed by the next generation supersonic transport to reduce the skin friction drag and to realize a high L/D ratio above 10 at supersonic cruising. For this purpose it has been proposed to apply laminar flow control (LFC) for the wing boundary layer. To achieve such LFC we have to develop much fundamental knowledge on the receptivity and instability of supersonic boundary layer growing in the leading edge region. This is exactly the motivation of our numerical and experimental studies. Through investigating the supersonic receptivity and instability problems we obtained the following results. Numerically the receptivity of the supersonic boundary layer (at a free-stream Mach number 2.2) is found to be very week, with the excited amplitude of TS wave being as low as 1/10 compared with the corresponding incompressible case. By disturbing the boundary layer near the attachment-line of a swept wing (with swept angle 60 to 70 degrees) in Mach 2.5 free-stream it is numerically shown that streamwise vortices of cross-flow instability type appear with the highest growth. For the case of a Mach 2.5 flat-plate boundary layer disturbed by strong periodic jets from wall orifice we learned numerically that hairpin vortices develop like in low-speed boundary layers, suggesting the occurrence of the non-linear subcritical transition. We made careful experiments to clarify how the acceleration can suppress the subcritical transition and found that the critical Reynolds number (based on momentum thickness) for the subcritical transition increases from 120 for the flatplate flow to 300 for a low-speed boundary layer with a non-dimensional favorable pressure gradient, -4.5x10^<-3>, which being scaled with the dynamic pressure and momentum thickness. It is also stressed that we found the instability due to streamline curvature to exist in the leading edge region of a swept wing in addition to the cross flow instability.