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A COMPARATIVE STUDY OF THE FILM COOLING HOLE CONFIGURATION EFFECTS ON THE LEADING EDGE OF ASYMMETRICAL TURBINE BLADE

DOI: 10.1615/ICHMT.2009.HeatTransfGasTurbSyst.200
16 pages

Mustapha Benabed
Aeronautical Laboratory and Propulsive Systems, Faculte de Genie-Mecanique, Universite des Sciences et de la Technologie d'Oran, B.P. 1505 El-Mnouar, Oran, Algeria

Abbes Azzi
Laboratory of Naval Aero-Hydrodynamic, Faculty of Mechanical Engineering, Oran University of Sciences and Technology, PO Box 1505, El-Mnaouar Oran, Algeria

B. A. Jubran
Department of Aerospace Engineering, Ryerson University, 350 Victoria Street, Toronto, Ontario, Canada

Abstract

The focus of this comparative-numerical study is to investigate the effects of advanced cooling hole geometries on film cooling effectiveness. Computational results are presented for a row of coolant injection holes on each side of an asymmetrical turbine blade model near the leading edge. Six film cooling configurations are considered in the present study, namely: (1) a cylindrical film hole, (2) a shaped film hole, (3) a uniform film slot, (4) a convergent film slot, (5) a crescent film hole, and (6) a trenched film hole. All simulations are conducted for the same density ratio of 1.0 and the same inlet plenum pressure. Results show that, at the suction side, the five configurations provide an increase of film effectiveness with the convergent, shaped, and crescent slot exits providing the highest increase over the baseline case by as much as 43%. At the pressure side, the other configurations also produce an increase of film effectiveness of up to 33%.

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