Patent ID: 6461108
Filing Date: 2002-10-08
Classification: C23C,F01D,Y02T,Y10T

Abstract:
A cooling system for cooling of the squealer tip region of a high pressure turbine airfoil used in a gas turbine engine comprising:a superalloy tip cap; a superalloy squealer tip extending outward in an engine radial direction from the superalloy tip cap into a hot gas stream of the engine; at least one channel having a first and second end, the first end terminating in an exit orifice located on a surface of the tip cap, the second end connecting to a cooling circuit located within a substrate, wherein the at least one channel has a diameter to permit an effective flow of cooling fluid; a bond coat having a thickness of about 0.0005â€³ to about 0.010â€³ applied to the tip cap surface, wherein the bond coat partially fills the exit orifice of the at least one channel; a layer of porous thermal barrier coating (TBC) having a first amount of predetermined porosity applied over the bond coat such that the porous TBC substantially covers the remainder of the exit orifice of the at least one channel; and, a layer of dense ceramic TBC applied to the porous layer of TBC, wherein the dense TBC layer has an amount of predetermined porosity so that it is less porous than the porous TBC layer.