Review of Platform Cooling Technology for High Pressure Turbine Blades
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Copyright © 2014 by ASME. With the relatively large surface area of the platform of the gas turbine blades being exposed directly to the hot, mainstream gas, it is vital to efficiently cool this region of the blades. This region is particularly difficult to protect due to the strong secondary flows developed at the airfoil junction (formation of the leading edge horseshoe vortex) and circumferentially across the blade passage (strengthening passage vortex moving from the pressure side to the suction side of the passage). Over the past decade, researchers and engine designers have attempted to combat the enhanced heat transfer to the blade platform by implementing both frontside and backside novel cooling techniques. This paper presents a review of platform cooling technology ranging from frontside film cooling via stator-rotor purge flow, mid-passage purge flow, and discrete film holes to backside cooling achieved via impinging jet arrays or cooling channels. To gain a full understanding of state-of-the-art cooling technology, recent patents, journal articles, and conference proceedings are included in this review.
author list (cited authors)
Wright, L. M., Malak, M. F., Crites, D. C., Morris, M. C., Yelavkar, V., & Bilwani, R.