Experimental Technique Study Of Heat Transfer Measurement On Sharp Leading Edges

As the development of new hypersonic vehicles, which can make a long time and long range flight, configuration with high ratio of lift to drag becomes the key point. For the blunt wedges could not meet the needs of air performances of the new hypersonic vehicles, the present of sharp leading edge with minimal drag becomes reasonable. Sharp leading edges offer numerous advantages on air performances. Meanwhile sharp leading edges bring high level of aero heating and difficulty of thermal protection. As the new hypersonic vehicles want to maintain its shape and high ratio of lift to drag, it has to employ non ablating thermal protection, which needs high precision heat-transfer data. It is of great significance for the development of new hypersonic vehicles to make experimental technique study of heat transfer measurement of sharp leading edges, which would provide reliable data for theoretical predictions and thermal protection researches.Investigations on transient heat transfer measurement in shock tunnel have been done. For the characters of small scale and high heating level of sharp leading edges, an instrument-model integration technique was occupied to develop a kind of integration thin film gauge, which was suitable for measurement of stagnation-point heat transfer on the sharp leading edges. Studies on Fay-Riddell relationship, low Reynolds number theory and rarefied flow effects have been conducted. Through studying on the flow properties of stagnation region of nose tips and sharp leading edges, methods of prediction of aero heating on stagnation point of such configurations have been achieved.Measurement of heat transfer on sharp leading edges has been carried out in shock tunnel FD-20. In the experiments, several cylinder-wedge geometry models, which were the combinations of gauges and steel bases, were employed. Those models had a range of nose radius from 1 - 5 mm. Results of stagnation-point heat transfer of unswept leading edges were obtained for Mach 5 and 6. The repeatability error of those results was less than 15%. In the radius range of 3 - 5 mm, measurement results of heat transfer on the stagnation point were close to that of continuum theory, indicating that classical boundary layer theory could apply. As the Reynolds numbers decreased, in the radius range of 1 - 1.5 mm, the shock thickness increased, and the thin boundary layer grew out toward the strong shock wave. There exited large flow interference in the stagnation region, effect of low Reynolds numbers appeared, and experimental results were larger than that of continuum theory by 10% to 20%. The increasing trend of experiment results was in agreement with modified boundary layer theory. At last this paper presents requirements for future works.