The present study evaluates an innovative approach for enhancement of surface heat transfer in a channel using concavities, rather than protruding elements. Serving as a vortex generator, a concavity is expected to promote turbulent mixing in the flow bulk and enhance the heat transfer. Using a transient liquid crystal imaging system, local heat transfer distribution on the surface roughened by an staggered array based on two different shapes of concavities, i.e. hemispheric and tear-drop shaped, have been obtained, analyzed and compared. The results reveal that both concavity configurations induce a heat transfer enhancement similar to that of continuous rib turbulators, about 2.5 times their smooth counterparts 10,000 ≤ Re ≤ 50,000. In addition, both concavity arrays reveal remarkably low pressure losses that are nearly one-half the magnitudes incurred with protruding elements. In turbine cooling applications, the concavity approach is particularly attractive in reducing system weight and ease of manufacturing.

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.

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Recent advances in thermochromic liquid crystal (TLC) thermography have improved its usefulness as a very effective temperature and heat transfer measurement technique. One of the approaches to determine the local heat transfer coefficient, known as the transient technique, is to monitor the temporal evolution of surface temperature in conjunction with the solution of a transient heat conduction model penetrating to the wall substrate. The local heat transfer coefficient resulted from such a transient test, by nature, has its reference temperature based on the inlet temperature of the test rig, rather than the local bulk mean temperature. The latter during a transient test varies with both time and streamwise location. The heat transfer coefficient based on the inlet temperature presents difficulty in data interpretation in designs of turbine cooling passages, particularly for passages with large length-to-diameter ratios. This study evaluates four different approaches and theoretical background associated for determining the local bulk mean temperature and the sensible local heat transfer coefficient. Using a test model of an internal cooling passage with delta-wing shaped vortex generators mounted on one of the passage walls, the magnitudes of the sensible heat transfer coefficient resulted from various approaches vary as much as 40%. Validated with the experimental data, two of the four methods yield superb data accuracy. Nevertheless, one of them stands out as the best choice, as it requires much less post-processing time and implementation effort.

Experimental tests were performed at the USAF Turbine Research Facility (TRF) to obtain heat transfer and aerodynamic data on a first stage vane of a modern high pressure turbine. This is a transient blowdown facility that provides data from short duration tests. Data for a matrix of test conditions were obtained to document the effect of inlet Reynolds number, the stage pressure ratio across the vane, and the gas-to-wall temperature ratio. The objectives of these tests were to assess the capability of obtaining accurate aerodynamic total pressure loss measurements and airfoil static pressure measurements as well as determine the heat transfer coefficient distributions on the vanes. Results from these tests were compared to analytical predictions and are presented. The unique contribution of the work presented herein is: 1) demonstration of circumferential traversing temperature and pressure data in a short duration facility test, and 2) heat loss closure during a short duration test using heat flux gauges and the measured temperature loss. The transient heat loss during a short duration test is a fundamental requirement to determine turbine efficiency when work extraction is determined from the temperature drop across the turbine stage. Heat transfer data were acquired from heat flux gauges that were fabricated using thin-film sputtering techniques and placed on the airfoil surfaces. The surface temperature of the gauge was measured and heat flux was determined from a closed form transient semi-infinite solution that included the resistance of the heat flux gauge and the underlying metal substrate. Circumferentially, pressure measurements were obtained on the airfoil surfaces and on traversing rakes at the inlet and exit of the vane test section. Total and differential pressure rake instrumentation was required to obtain accurate aerodynamic loss measurements over a range of gas-to-wall temperature ratios.