Investigation of Leading Edge Film Cooling Jets in a Cross Flow

Abstract

The thermal effect of the turbine blade film cooling and the penetration area of jets issuing at an angle into cross mainstream flow have been investigated numerically and experimentally. Experimental and numerical simulations have been introduced to simulate a discrete circle hole film cooling flow over a symmetrical airfoil representing turbine guide vanes surface. Several cases have been studied in the experimental work by using three-velocity ratio (0.5, 0.9, 1.3), and three different jet issuing angles, longitudinal injection angle (37.50 and 900) both with lateral injection angle (stagger angle = 00 and 450). Airfoil angle of attack has been changed during the experimental program throughout (00, 50, 100, and 150). Experimental investigations gave qualitative information about the penetration area and flow structure of the mixing flow at all cases and the results were used to verify the computation method and to select the best velocity ratios for the flow penetration, flow structure, and the thermal effect of cool jets.