超音速细长梯形翼背风面流型

给出了前缘后掠６５°、双弧形剖面的细长梯形翼背风面流动显示结果。实验Ｍａｃｈ数为１．１０，１．５３，２．５３，３．０１和４．０１，攻角范围为５°～２５°。应用蒸汽屏、纹影和油流技术拍摄了空间和表面流型照片。蒸汽屏显示表明：在机翼背风面三角形区域的空间流型随法向攻角αＮ（在垂直于前缘的平面内流速与弦线间的夹角）和法向Ｍａｃｈ数ＭａＮ（来流Ｍａｃｈ数在垂直于前缘平面内的分量）变化，并可在αＮ和ＭａＮ为坐标的平面上划分出７种流型存在的区域。侧缘区有侧缘分离涡形成；后缘有尾涡拖出。从纹影照片与横截面上的蒸汽屏照片对照可获得机翼锥面激波位置随Ｍａｃｈ数的变化；以及激波－诱导分离线位置随Ｍａｃｈ数和攻角变化曲线。机翼表面油流谱显示出了主再附线、二次分离线、二次再附线和侧缘涡区。显示出的流型与其他有关实验和数值计算结果比较符合得很好

TYPES OF FLOW OVER LEESIDE OF A CLIPPED DELTA WING AT SUPERSONIC SPEEDS

This paper describes a flow visualization experiment over the leeside of a slender tapered wing with a leading edge sweepback angle of 65° and a biconvex section. By using laser vapor screen, schlieren and oil flow techniques, the test was carried out at Mach numbers of 1.10, 1.53, 2.53, 3.01 and 4.01 for angles of attack from 5° through 25°. And photos of flow off body and on lee surface have been taken. The vapor screen photos show seven distinct types of flow developed over the delta region of the wing. These types of flow are displayed in a plane of Mach number and angle of attack normal to the leading edge. In the region of the side edge there are side vortices formed and in the region of the trailing edge trailing vortices shed. The variations of both the bow shock positions at different Mach numbers and the positions of separating lines induced by shock waves with Mach numbers and angles of attack are obtained based upon the schlieren plus the vapor screen photos in the sections. The oil flow visualization shows clearly primary reattachment, secondary separation, secondary reattachment lines and side edge vortices on the lee surface. The results of this test are in good agreement with experiments and numerical simulations of others.