Thesis (Ph.D.) - Cairo University - Faculty of Engineering - Department of Aerospace Engineering

Over the past fifty years, aircraft and power generation gas turbine designers have focused on increasing the temperature at the combustor exit to improve the engine performance, to increase the thrust and to reduce the fuel consumption. Unfortunately, these high temperatures have a negative effect on the integrity of the high-pressure turbine components and specifically the turbine blades. Therefore, there is a need to have an efficient cooling system engineered in a way such that the maximum blade surface temperature during the engine operation does not exceed the maximum allowable temperature of the blade material. In addition, the tip leakage flow across the turbine blade reduces its aerodynamic performance and causes blade - tip overheating due to this high speed flow and the existence of a thin boundary layer. Film cooling is vital to gas turbine blades to protect them from high temperatures and hence high thermal stresses. In the current study, film cooling on a flat plate is studied numerically to select the different parameters of the compound - hole shape that could enhance the cooling effectiveness

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