The clearance gap between the stationary outer air seal and blade tips of an axial turbine allows a clearance gap leakage flow to be driven through the gap by the pressure-to-suction side pressure difference. The presence of strong secondary flows on the pressure side of the airfoil tends to deliver air from the hottest regions of the mainstream to the clearance gap. The blade tip region, particularly near the trailing edge, is very difficult to cool adequately with blade internal coolant flow. In this case, film cooling injection directly onto the blade tip region can be used in an attempt to directly reduce the heat transfer rates from the hot gases in the clearance gap to the blade tip. The present paper is intended as a memorial tribute to the late Professor Darryl E. Metzger who has made significant contributions in this particular area over the past decade. A summary of this work is made to present the results of his more recent experimental work that has been performed to investigate the effects of film coolant injection on convection heat transfer to the turbine blade tip for a variety of tip shapes and coolant injection configurations. Experiments are conducted with blade tip models that are stationary relative to the simulated outer air seal based on the result of earlier works that found the leakage flow to be mainly a pressure-driven flow which is related strongly to the airfoil pressure loading distribution and only weakly, if at all, to the relative motion between blade tip and shroud. Both heat transfer and film effectiveness are measured locally over the test surface using a transient thermal liquid crystal test technique with a computer vision data acquisition and reduction system for various combinations of clearance heights, clearance flow Reynolds numbers, and film flow rates with different coolant injection configurations. The present results reveal a strong dependency of film cooling performance on the choice of the coolant supply hole shapes and injection locations for a given tip geometry.

Described in this paper is an experimental study of heat transfer over a trailing edge configuration preceded with an internal cooling channel of pedestal array. The pedestal array consists of both circular pedestals and oblong shaped blocks. Downstream to the pedestal array, the trailing edge features pressure side cutback partitioned by the oblong shaped blocks. The local heat transfer coefficient over the entire wetted surface in the internal cooling chamber has been determined by using a “hybrid” measurement technique based on transient liquid crystal imaging. The hybrid technique employs the transient conduction model in a semi-infinite solid for resolving the heat transfer coefficient on the endwall surface uncovered by the pedestals. The heat transfer coefficient over a pedestal can be resolved by the lumped capacitance method with an assumption of low Biot number. The overall heat transfer for both the pedestals and endwalls combined shows a significant enhancement compared to the case with thermally developed smooth channel. Near the downstream most section of the suction side, the land, due to pressure side cutback, is exposed to the stream mixed with hot gas and discharged coolant. Both the adiabatic effectiveness and heat transfer coefficient on the land section are characterized by using the transient liquid crystal technique.

Pin fins are commonly used as heat transfer augmentors for internal cooling of turbine airfoils. These pins may extend from one wall to the other or may be segmented to meet specific requirements of removing the airfoil’s varying heat load. Three configurations of the partial pins were tested in a 25:1 aspect ratio channel and the results are compared with those for the full pins. The array average heat transfer rate decreases linearly with increasing gap and is bounded by the value for full pins at one end and that for the smooth channel at the other. However, the local distribution of the Nusselt number and the average for each of the two walls depends on the configuration of the partial pins. The friction factor was lower for partial pins than for the full pins and also decreased with increasing gap. For the configuration with all partial pins on one wall, the friction factor was found to be the lowest with no change in the corresponding heat transfer rate from a wall with pins.

Pin fins are commonly used as heat transfer augmentors for internal cooling of turbine airfoils. These pins may extend from one wall to the other or may be segmented to meet specific requirements of removing the airfoil’s varying heat load. Three configurations of the partial pins were tested in a channel with aspect ratio 25:1 and the results are compared with those for the full pins. The array average heat transfer rate decreases linearly with increasing gap and is bounded by the value for full pins at one end and that for the smooth channel at the other. However, the local distribution of the Nusselt number and the average for each of the two walls depends on the configuration of the partial pins. The friction factor was lower for partial pins than for the full pins and also decreased with increasing gap. For the configuration with all partial pins on one wall, the friction factor was found to be the lowest, with no change in the corresponding heat transfer rate from a wall with pins.

Aircraft propulsion engines, land-based power generation, and industrial machines have, as a primary component, the turbine as means to produce thrust or generate power. In the turbine section of the engine, airfoil components are subjected to extremely complex and damaging environments. The combination of high gas temperatures and pressures, strong gradients, abrupt geometry changes, viscous forces, rotational forces, and unsteady turbine vane/blade interactions, all combine to offer a formidable challenge in terms of turbine durability. Nevertheless, the ultimate goal is to maintain or even improve the highest level of turbine performance and simultaneously reduce the amount of cooling flow needed to achieve this end. As such, coolant flow is a penalty to the cycle and thermal efficiency. Cooling strategies are developed and presented to determine ways for coolant flow management. The main variables include film cooling configurations, and convective efficiency schemes to balance turbine airfoil thermal loads for target overall cooling effectiveness. The desired targets are determined by the turbine airfoil durability requirements of oxidation and fatigue on a local scale and for creep on the bulk scale. Emphasis is provided to the general modes of cooling including film cooling, impingement cooling, and convective cooling for different parts of the airfoil such as leading edge, mid-body, trailing edge, tip and endwalls. Convective cooling is presented in terms of fundamental cooling enhancements, such as turbulating trip strips and pedestals. Recent literature dealing with these topics is listed.