Review of Platform Cooling Technology for High Pressure Turbine Blades

Abstract

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.

Conjugate heat transfer modeling of a turbine vane endwall with thermal barrier coatings

“…Heat transfer on a turbine endwall behaves completely differently compared to that on a vane surface due to the presence of strong, complex and three-dimensional secondary flows near the endwall, including horseshoe vortex, passage vortex and other small but strong corner vortex. The vortices in the vane passage affect the endwall through distorting the endwall-nearby flows, locally enhancing the external heat transfer levels and redirecting the film cooling injection, as well discussed by Simon and Piggush (6) and Wright et al (7) .The first experimental study of platform conjugate heat transfer was conducted by Mensch and Thole (8) in a low-pressure turbine blade. They presented overall cooling effectiveness on the blade platform constructed to match Biot number with an engine and compared the conjugate heat transfer results from film cooling only, internal impingement cooling only and combined impingement and film cooling.…”

Numerical investigation of heat transfer and flow characteristics of a double-wall cooling structure: Reverse circular jet impingement

“…Jet impingement has been used in many applications as a result of its high convective heat transfer rates, especially near the stagnation zone. Therefore, the jet impingement technique is widely used in a variety of applications and industries (i.e., gas turbine blades and combustion chamber cooling, aircraft wing de-icing, electronic component cooling, food processing, glass and metal tempering, paper drying, and textiles) [1], [2].…”

Influence of slashface leakage coupled with quasi-labyrinth seal technique on gas turbine endwall aerothermal performance and blade suction side surface phantom cooling

“…In specific, high-pressure coolant gas is extracted from compressor, and ejects out from turbine surfaces to form a thin protection layer. 10 Effects of film cooling hole arrangement pattern, 11,12 hole shape, 13 and flow parameters 14,15 on the surface cooling performance have been investigated on flat plate, and also in the turbine passage. Kneer et al 16 utilized superposition principle of film cooling to experimentally measure the locally film-cooled endwall.…”

“…Coolant needs to be ejected out from the slot to prevent mainstream ingestion, which can also play a role in film cooling. 10 Knost et al 19 have experimentally measured the adiabatic endwall cooling effectiveness with upstream slot and discrete film cooling holes in a stationary cascade. The results show that it is hard for the upstream slot coolant injection to protect the pressure side endwall junction and leading edge.…”