Abstract: A coolable gas turbine engine airfoil for a gas turbine engine suitable for a blade or vane includes an outer airfoil wall with pressure and suction sides extending chordwise between leading and trailing edges of the airfoil, a leading edge cooling plenum formed between a forward most span rib and the outer wall along the leading edge of the airfoil, and a cooling air channel within the airfoil bounded in part by the forward most rib. A slotted cooling air impingement element is disposed in the span rib for impinging cooling air from the channel on an interior surface of the outer airfoil wall along the leading edge of the airfoil. One embodiment of the slotted cooling air impingement element is a single longitudinally extending slot extending along almost an entire length of the forward most rib and the longitudinally slot preferably includes longitudinally spaced apart rounded ends.

Abstract: A curved component such as a turbine airfoil, shroud, or combustor centerbody is prepared with a platinum or a platinum-aluminide protective coating over only a portion of the surface thereof. The coating may serve as an environmental coating, or as a bond coat of a thermal barrier coating system. The partial coverage is achieved by depositing platinum only over a portion of the surface of the component, typically including the concave portion in the case of an airfoil, optionally depositing an aluminum layer, and optionally interdiffusing the platinum and aluminum layers.

Abstract: The present invention discloses a system for cooling turbine rotor disk posts in a gas turbine engine with a preferred embodiment illustrating an application for a high pressure turbine stage 1 rotor disk. The disk includes a plurality of circumferentially alternating posts and slots disposed about the periphery of the disk with each slot receiving a dovetail of a radially extending blade. The cooling system comprises a plurality of seal bodies with each seal body including a relatively low volume thermal isolation chamber positioned over the top of a corresponding disk post; a plurality of axially extending blade cooling plenums supplied with cooling air with one of the plenums positioned radially inward of each of the blades; a plurality of shallow, radially extending slots formed in a relatively low stressed axially facing surface of each of the blade dovetails for diverting cooling air from the blade cooling plenums to the thermal isolation chambers.

Abstract: A turbine blade squealer tip has a plurality of holes defined through the end cap of the blade tip portion and located contiguous with an interior surface of the squealer tip wall along at least one of the turbine blade pressure and suction sides of the tip cap for providing effective cooling air on the squealer tip wall. Each tip hole has an inner portion which communicates with the interior of the blade and an outer portion which communicates with a cavity defined by the blade end cap and the peripheral wall of the squealer tip. The cross-section of the inner hole portion is circular and the inner hole portion overall defines a right cylinder. The cross-section of the outer hole portion is rectangular and the outer hole portion overall defines a rectangular trapezoid.