Global Aeroheating Measurements of Shock-Shock Interactions on Swept Cylinder

摘要

The effects of the fin leading-edge radius and sweep angle on peak heating rates due to shock-shock interactions were investigated in the NASA Langley Research Center 20-Inch Mach 6 Air Tunnel. The cylindrical leading-edge models, with radii varied from 0.25 to 0.75in., represent wings or struts on hypersonic vehicles. A planar oblique shock at 16.7deg to the flow intersected the fin bow shock, producing a shock-shock interaction that impinged on the fin leading edge. Three fin sweep angles were tested: 0, -15, and -25deg (swept forward). Global temperature data were obtained from the surface of the fused silica fins using phosphor thermography. Metal oil-flow models were used to visualize the streamline patterns for each angle of attack. High-speed zoom-schlieren videos were recorded to show the features of the shock-shock interactions. The temperature data were analyzed using one-dimensional semi-infinite and one- and two-dimensional finite-volume methods. These results were compared to determine the proper heat transfer analysis approach to minimize errors from lateral heat conduction. The dimensional peak heat transfer coefficient augmentation increased with decreasing leading-edge radius. The dimensional peak heat transfer output from the two-dimensional code was about 20% higher than the value from a standard, semi-infinite one-dimensional method.