Abstract: A gas turbine engine hollow airfoil includes an airfoil outer wall having width wise spaced apart pressure and suction side walls joined together at chordally spaced apart leading and trailing edges of the airfoil and extending radially outward from a base to a tip. The airfoil includes a single three pass aft flowing serpentine cooling circuit and a straight through leading edge cooling channel extending radially through the airfoil and bounded in part by the leading edge. The exemplary embodiment includes a trailing edge cooling plenum and a straight through single pass trailing edge feed channel for providing cooling air to the trailing edge cooling plenum.

Abstract: A gas turbine engine hollow airfoil has an airfoil outer wall with widthwise spaced apart pressure and suction side walls joined together at chordally spaced apart leading and trailing edges of the airfoil and extending radially from a radially inner base to a radially outer airfoil tip. A cooling circuit within the airfoil has radially extending first, middle, and last channels arranged respectively in series with the first channel in fluid communication with a source of cooling air from outside the airfoil. The last channel is in fluid communication with one of the edges. A refresher passageway extends through a radially inner wall bounding a radially inner portion of the last channel and is in fluid communication with the source of cooling air. The refresher passageway is separate, spaced apart, and independent from the first channel. In one embodiment, a metering plate covers an inlet to the refresher passageway and the metering plate has a metering hole over the inlet and the metering hole is adjustable.

Abstract: A rotor blade for a gas turbine engine including a tip region that facilitates reducing operating temperatures of the rotor blade is described. The tip region includes a first tip wall and a second tip wall extending radially outward from a tip plate of an airfoil. The tip walls extend from adjacent a leading edge of the airfoil to connect at a trailing edge of the airfoil. A notch is defined between the first and second tip walls at the airfoil leading edge. At least a portion of the second tip wall is recessed to define a tip shelf.

Abstract: A rotor blade for a gas turbine engine including a tip region that facilitates reducing operating temperatures of the rotor blade is described. The tip region includes a first tip wall and a second tip wall extending radially outward from a tip plate of an airfoil. The tip walls extend from adjacent a leading edge of the airfoil to connect at a trailing edge of the airfoil. A notch is defined between the first and second tip walls at the airfoil leading edge. At least a portion of the second tip wall is recessed to define a tip shelf.

Abstract: A turbine bucket includes an airfoil extending from a platform, having high and low pressure sides; a wheel mounting portion; a hollow shank portion located radially between the platform and the wheel mounting portion, the platform having an under surface. An impingement cooling plate is located in the hollow shank portion, spaced from the under surface, and the impingement plate is formed with a plurality of impingement cooling holes therein.

Abstract: A cooling hole configuration for an air-cooled component, such as a gas turbine engine airfoil. The cooling hole is configured to have cross-sectional variations and a noncircular-shaped diffuser-type opening that significantly improve the cooling film distribution across the external surface of an airfoil, with the result that heat transfer from the surrounding environment to the airfoil is reduced. The cooling hole is configured to have its central axis at an acute angle to the exterior surface of the airfoil, and defines a noncircular-shaped opening at the airfoil surface. The cooling hole generally has a first region adjacent the airfoil surface and a second region interior to the airfoil.

Abstract: A method of forming a shaped hole in a substrate, including substrates protected by coatings, as is the case where an air-cooled component has a thermal barrier coating for protection from a hostile thermal environment, such as the turbine, combustor and augmentor sections of a gas turbine engine. In particular, cooling holes are formed in an air-cooled component after a ceramic layer has been deposited on the surface of the component without damaging or spalling the ceramic layer. Processing steps generally include forming a hole through the ceramic layer and a substrate protected by the ceramic layer, thereby defining an opening at the surface of the ceramic layer. The diameter of the hole is less than the final size required for the cooling hole.

Abstract: A seal joint including a sealing member for sealing segmented components of a gas turbine engine is spaced from the gas path to provide more effective cooling. The segments include radial extension flanges spacing the seal joint from the gas path. Various cooling arrangements are adaptable to the modified segments to provide for open or closed circuit impingement or convection cooling plus film cooling. Excessive thermal gradients are avoided as the entire corner of the segment is evenly cooled. Moreover, air usage at the joint is bounded by seal leakage with no extra dedicated cooling air.