An improved numerical method is presented for calculating the load distribution on oscillating rectangular wings in subsonic flow. The wing is divided into many chordwise strips. In the strips containing control point, the kernel function is treated in the form of expansion series and spanwise integration is performed first to get a chordwise logarithmic singularity. Through accurate evaluation the singularity, present method gives quick convergence and excellent computational efficiency to other current methods.
This paper presents an experimental study of shock wave/turbulent boundary layer interaction induced by a semicone placed on the floor of the wind tunnel. The experiments were carried out in a 8×10cmcm2 supersonic wind tunnel at free stream Mach number of 1.98. The Reynolds number per metre in the test section was 4.01×107. The semicone models with half angles of 20°, 30°, 40°, 50° and 60° were used in this study. The ratio of the height to the boundary layer thickness was greater than 2. Surface static pressure measurements, oil flow studies and Schlieren photographs of the flow field were made. It was found that, the separation shock angle is insensitive to cone angle. The length of the separation point measured along the centerline from the apex of the semicone was found to increase linearly with the increase of cone angle and the shape of the separation lines fall into a single curve irrespective of cone angle. Furthermore, the pressure distribution of the flat plate became qualitatively sfmilar to the hemicylindrical fin until the plateau region in the spanwise extent of y/d=0.5.
The breakdown structure of longitudinal vortices in a boundary layer along a concave wall of 1m in radius of curvature is studied experimentally under main air flow velocity of 2.5 m/s. The flow patterns are observed by lightoil vapour with stroboscope, and the mean velocity and the turbulence intensity are measured by a hot-wire anemometer. In the upstream reagion, every couple of the vortices forms periodic high-frequency horseshoe-type vortices, which bring about dominant behaviour in the outer layer. In the downstream reagion, it is observed that high-intensity irregular fluctuations are generated beneath the horseshoe-type vortices in the inner layer. Therefore, the longitudinal vortices are broken down, and the boundary layer transition is also hastened downstream.
The experimental study was conducted to determine the effects of the pressure distortion on the operation of a turbofan engine. The turbofan engine was tested with three kinds of inlet distortions, i.e., tip radial, circumferential and combined pressure distortions. The each distortion was generated with screens placed in the inlet ducting upstream of the fan face. Their intensity was varied by the screens with different porosity. This paper presents an investigating result of the evaluation method of the stability for the distortions and the experimental results of the effect on the two characteristics, pressure ratio and efficiency, of two stage axial flow fan and five stage axial flow compressor. The new method is to indicate two distortion indices on the orthogonal coordinate, the circumferential distortion index is related to the radial distortion index. This method led to the estimation that the examination of three kinds of the distortion seems to be enough to cover the total patterns of the distortion encountering inflight. The pressure ratio and the efficiency of a fan with distorted inflow decreased as compared with those of the clean inlet flow, however the operation line and the efficiency of a compressor hardly suffered from influence of the distortion.