Patent ID: 7186091

Claim:
A method of fabricating a gas turbine engine component, said method comprising forming a plurality of pores in a wall of the component, wherein the pores extend substantially perpendicularly through the wall, wherein the wall includes a first surface and an opposite second surface, wherein the pores each include a first diameter defined by the wall first surface and a second diameter defined by the opposite wall second surface; forming a plurality of film cooling holes in the wall, wherein the holes extend substantially perpendicularly through the wall; coating the first wall surface of the wall of the component with a thermal barrier coating (TBC) such that the TBC extends over and seals a first end of the pores, wherein at least one of the plurality of pores has the first diameter at the first wall surface that is smaller than the second diameter at the opposite wall second surface therein; and coupling the component in flow communication to a cooling fluid source, such that during operation cooling fluid may be channeled through the pores for back side cooling an inner surface of the thermal barrier coating, and such that cooling fluid may be channeled through the holes for film cooling an outer surface of the thermal barrier coating.