Abstract: An improved cooling design and method for cooling airfoils within a gas turbine engine is provided which includes a plenum longitudinally located within the leading edge of the airfoils. Within the plenum are positioned a plurality of turbulence promoters to provide enhance heat transfer within the leading edge. Also, the cooling design includes a plurality of inlets to receive cooling air from an internal cavity of the airfoil as well as a plurality of outlets located within a trench on the exterior surface of the leading edge through which the cooling air exits to film cool leading edge.

Abstract: A cooling circuit disposed within a first wall portion and a second wall portion of a wall for use in a gas turbine engine is provided. The cooling circuit includes a plurality of inlet apertures disposed in the first wall portion to provide a flow path of cooling air into the cooling circuit, a first exit slot disposed in the second wall portion to provide a cooling air flow path out of the cooling circuit, a plurality of pedestals arranged in rows extending between the first wall portion and the second wall portion and a first elongated pedestal disposed radially and extending between the first and second wall portions. The first elongated pedestal is positioned between the pedestals and the first exit slot.

Abstract: A turbine blade is provided comprising an internal cavity into which cooling air is flowable, an external wall and a first, second and third cooling circuits embedded with the wall. The cooling circuits include inlets that connect each respective cooling circuit with a cavity to provide a cooling air flow path into the respective cooling circuit. The cooling circuits also include an exit aperture that provides a cooling air flow path out of the respective cooling circuits. The cooling circuits are configured to increase the temperature of the cooling air as it travels from the inlet to the exit aperture. In the exemplary embodiment, the inlets are aligned with the direction of counter-rotating flow circulations experienced by the inner surface of the wall caused by Coriolis flow effects on the cooling air flowing inside the cavity.

Abstract: A turbine element airfoil has a cooling passageway network with a slot extending from a trailing passageway toward the trailing edge. A number of discrete posts span the slot between pressure and suction sidewall portions.

Abstract: An improved cooling design and method for cooling airfoils within a gas turbine engine is provided which includes a plenum longitudinally located within the leading edge of the airfoils. Within the plenum are positioned a plurality of turbulence promoters to provide enhanced heat transfer within the leading edge. Also, the cooling design includes a plurality of inlets to receive cooling air from an internal cavity of the airfoil as well as a plurality of outlets located within a trench on the exterior surface of the leading edge through which the cooling air exits to film cool the leading edge.

Abstract: An improved cooling design and method for cooling airfoils within a gas turbine engine is provided which includes an embedded microcircuit that traverses a tip between a suction sidewall and a pressure sidewall of the airfoil. The microcircuit includes at least on inlet disposed proximate to the tip and one of the sidewalls for receiving cooling air from an internal cooling cavity of the airfoil and at least outlet disposed proximate to the tip through which the cooling air ejects into a region outside the airfoil.

Abstract: A turbine airfoil section having an internal cavity and a plurality of indentations on the inner surface of the internal cavity is described. The indentations provide enhanced heat transfer for cooling the internal cavity of an airfoil thereby improving the life of the airfoil and optimizing the efficiency of the engine by minimizing the amount of compressor bleed air required. Advantageously, this cooling scheme also does not restrict the cooling flow within the internal cavity. The indentations may have varying patterns and alternative geometric configurations.

Abstract: A turbine airfoil section having an internal cavity and a plurality of indentations on the inner surface of the internal cavity is described. The indentations provide enhanced heat transfer for cooling the internal cavity of an airfoil thereby improving the life of the airfoil and optimizing the efficiency of the engine by minimizing the amount of compressor bleed air required. Advantageously, this cooling scheme also does not restrict the cooling flow within the internal cavity. The indentations may have varying patterns and alternative geometric configurations.

Abstract: A turbine vane assembly which includes an outer endcap having a plurality of generally straight passages and passage segments therethrough, an inner endcap having a plurality of passages and passage segments therethrough, and a vane assembly having an outer shroud, an airfoil body, and an inner shroud. The outer shroud, airfoil body and inner shroud each have a plurality of generally straight passages and passage segments therethrough as well. The outer endcap is coupled to the outer shroud so that outer endcap passages and said outer shroud passages form a fluid circuit. The inner endcap is coupled to the inner shroud so that the inner end cap passages and the inner shroud passages from a fluid circuit. Passages in the vane casting are in fluid communication with both the outer shroud passages and the inner shroud passages. Passages in the outer endcap may be coupled to a cooling system that supplies a coolant and takes away the heated exhaust.

Abstract: A design for a vane segment having a closed-loop steam cooling system is provided. The vane segment comprises an outer shroud, an inner shroud and an airfoil, each component having a target surface on the inside surface of its walls. A plurality of rectangular waffle structures are provided on the target surface to enhance heat transfer between each component and cooling steam. Channel systems are provided in the shrouds to improve the flow of steam through the shrouds. Insert legs located in cavities in the airfoil are also provided. Each insert leg comprises outer channels located on a perimeter of the leg, each outer channel having an outer wall and impingement holes on the outer wall for producing impingement jets of cooling steam to contact the airfoil's target surface. Each insert leg further comprises a plurality of substantially rectangular-shaped ribs located on the outer wall and a plurality of openings located between outer channels of the leg to minimize cross flow degradation.