Numerical and Experimental Study of Film Cooling Effectiveness and Total Pressure Loss for a Realistic Cascade with an Annular Endwall

Abstract

This is a numerical and experimental study on the film cooling effectiveness and pressure loss of cylindrical-shaped film cooling on an annular cascade. Film cooling holes will be employed on the inner diameter endwall, just upstream of the blade stagnation region of GE-E3 first stage high pressure turbine in a subsonic commercial aircraft. For a deeper understanding of the interaction behavior between film cooling jets and secondary flows inside the cascade passage, five (four and two halves) 3D profile airfoils are used to obtain a periodic flow in the middle passage. The experimental test rig has an inner diameter of 970 mm and outer diameter of 1098 mm (scaled 3X actual size). Total pressure loss and loss coefficient after the trailing edge surface are measured using a Five Hole Probe with 0.5 o and 10mm of circumferential and radial increment respectively. Film cooling effectiveness will be reported using a TSP method. Experimental data is validated by an in-house RANS model and by the existing literature. Results show that the pressure loss and film cooling effectiveness found in the test rig coincide with existing literature as well as CFD analysis.