Cooling of turbine parts in a gas turbine engine is necessary for operation as the temperature of combustion gases is higher than the melting temperature of the turbine materials. The gap between rotating turbine blades and the stationary shroud provides an unintended flow path for hot gases. Gases that flow through the tip region cause pressure losses in the turbine section and high heat loads to the blade tip. This thesis studies the heat transfer on an innovative tip geometry intended to help reduce aerodynamic losses. The blade tip has a depression (shelf) on the tip surface along much of the pressure side of the blade and film-cooling holes along the depression. This research experimentally measured the effect of the shelf, coolant flow and tip gap on heat transfer on the blade tip.

Stationary experiments were performed in a low speed wind tunnel on a linear cascade with two different tip gaps and multiple coolant flow rates through the film-cooling holes. Tests showed that baseline Nusselt numbers on the tip surface were reduced with the shelf tip compared with a flat tip. Measurements indicated that film-cooling was more effective with a small tip gap than with a large tip gap. Experimental and computational results demonstrated a lack of coolant spreading that was detrimental to regions between the film-cooling holes. While the coolant was effective on the blade tip, the leading and trailing edge regions were found to have high heat transfer coefficients with little available cooling.