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

如何在超声速巡航时减小表面摩擦阻力，实现10以上的高升深比，是下一代超声速运输机面临的技术挑战。为此，有人提出对机翼边界层进行层流控制。为了实现这样的LFC，我们必须对前缘区域生长的超声速边界层的可接受性和不稳定性有许多基本的了解。这正是我们进行数值和实验研究的动机。通过对超声速接受性和不稳定性问题的研究，得到了以下结果。数值结果表明，超声速边界层（在自由流马赫数2.2时）的接受度非常低，TS波的激发幅值与不可压缩情况相比低至1/10。通过对2.5马赫自由气流中后掠翼（后掠角60 ~ 70度）附著线附近的边界层进行扰动，数值计算表明，横向流动不稳定型流向涡的增长速度最快。对于受壁孔强周期性喷流扰动的2.5马赫平板边界层，我们从数值上了解到发夹涡的发展与低速边界层相似，表明发生了非线性亚临界过渡。我们进行了仔细的实验，以明确加速度如何抑制亚临界转变，并发现亚临界转变的临界雷诺数（基于动量厚度）从平板流的120增加到低速边界层的300，该边界层具有无因次有利压力梯度，-4.5x10^<-3>，其与动压力和动量厚度成比例。文中还着重指出，在后掠翼前缘区域除了存在横向流动不稳定外，还存在流线曲率不稳定。

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.