Abstract: A seal assembly between a hot gas path and a disc cavity in a turbine engine includes a non-rotatable vane assembly including a row of vanes and an inner shroud, a rotatable blade assembly axially adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, and an annular wing member located radially between the hot gas path and the disc cavity. The wing member extends generally axially from the blade assembly toward the vane assembly and includes a plurality of circumferentially spaced apart flow passages extending therethrough from a radially inner surface thereof to a radially outer surface thereof. The flow passages each include a portion that is curved as the passage extends radially outwardly to effect a scooping of cooling fluid from the disc cavity into the flow passages and toward the hot gas path.

Abstract: A seal assembly between a disc cavity and a hot gas path in a gas turbine engine includes a stationary vane assembly and a rotating blade assembly axially upstream from the vane assembly. A platform of the blade assembly has a radially outwardly facing first surface, an axially downstream facing second surface defining an aft plane, and a plurality of grooves extending into the second surface such that the grooves are recessed from the aft plane The grooves are arranged such that a circumferential space is defined between adjacent grooves During operation of the engine, the grooves impart a circumferential velocity component to purge air flowing out of a disc cavity through the grooves to guide the purge air toward a hot gas path such that the purge air flows in a desired direction with reference to a direction of hot gas flow through the hot gas path.

Abstract: A gas turbine having rotor discs (9), a disc cavity (13) and a stator stage (25) extending to the disc cavity (13). Seal housing flanges (43, 44) extend from a seal housing (29) of the stator stage (25). Rotor flanges (41i, 41o) extend from a rotor disk (9-1). An inner rotor flange (41i) and first seal housing flange (43) are inward from a second seal housing flange (44). One rotor flange (41o) is outward from the second seal housing flange (44). The inner rotor flange (41i) and first seal housing flange (43) extend toward one another to limit movement of main gas flow (17). An inlet (47) injects air (50) between the outward rotor flange (41o) and second seal housing flange (44).

Abstract: An airfoil in a gas turbine engine includes an outer wall, a cooling fluid cavity, and a cooling system. The outer wall has a leading edge, a trailing edge, a pressure side, a suction side, and radially inner and outer ends. The cooling fluid cavity is defined in the outer wall, extends generally radially between the inner and outer ends of the outer wall, and receives cooling fluid for cooling the outer wall. The cooling system receives cooling fluid from the cooling fluid cavity for cooling the trailing edge portion of the outer wall and includes a cooling fluid chamber defined by opposing first and second sidewalls that include respective alternating angled sections that provide the cooling fluid chamber with a zigzag shape.

Abstract: A seal assembly between a hot gas path and a disc cavity in a turbine engine includes a non-rotatable vane assembly including a row of vanes and an inner shroud, a rotatable blade assembly axially adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, and an annular wing member located radially between the hot gas path and the disc cavity. The wing member extends generally axially from the blade assembly toward the vane assembly and includes a plurality of circumferentially spaced apart flow passages extending therethrough from a radially inner surface thereof to a radially outer surface thereof. The flow passages each include a portion that is curved as the passage extends radially outwardly to effect a scooping of cooling fluid from the disc cavity into the flow passages and toward the hot gas path.

Abstract: A seal assembly between a hot gas path and a disc cavity in a turbine engine includes a non-rotatable vane assembly including a row of vanes and an inner shroud, a rotatable blade assembly adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, and an annular wing member located radially between the hot gas path and the disc cavity. The wing member extends generally axially from the blade assembly toward the vane assembly and includes a plurality of circumferentially spaced apart flow passages extending therethrough from a radially inner surface thereof to a radially outer surface thereof. The flow passages effect a pumping of cooling fluid from the disc cavity toward the hot gas path during operation of the engine.

Abstract: An airfoil in a gas turbine engine includes an outer wall, a cooling fluid cavity, and a cooling system. The outer wall has a leading edge, a trailing edge, a pressure side, a suction side, and radially inner and outer ends. The cooling fluid cavity is defined in the outer wall, extends generally radially between the inner and outer ends of the outer wall, and receives cooling fluid for cooling the outer wall. The cooling system receives cooling fluid from the cooling fluid cavity for cooling the trailing edge portion of the outer wall and includes a cooling fluid chamber defined by opposing first and second sidewalls that include respective alternating angled sections that provide the cooling fluid chamber with a zigzag shape.

Abstract: A seal assembly between a hot gas path and a disc cavity in a turbine engine includes a non-rotatable vane assembly including a row of vanes and an inner shroud, a rotatable blade assembly adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, and an annular wing member located radially between the hot gas path and the disc cavity. The wing member extends generally axially from the blade assembly toward the vane assembly and includes a plurality of circumferentially spaced apart flow passages extending therethrough from a radially inner surface thereof to a radially outer surface thereof. The flow passages effect a pumping of cooling fluid from the disc cavity toward the hot gas path during operation of the engine.

Abstract: A seal assembly between a disc cavity and a hot gas path in a gas turbine engine includes a stationary vane assembly and a rotating blade assembly axially upstream from the vane assembly. A platform of the blade assembly has a radially outwardly facing first surface, an axially downstream facing second surface defining an aft plane, and a plurality of grooves extending into the second surface such that the grooves are recessed from the aft plane The grooves are arranged such that a circumferential space is defined between adjacent grooves During operation of the engine, the grooves impart a circumferential velocity component to purge air flowing out of a disc cavity through the grooves to guide the purge air toward a hot gas path such that the purge air flows in a desired direction with reference to a direction of hot gas flow through the hot gas path.

Abstract: A gas turbine having rotor discs (9), a disc cavity (13) and a stator stage (25) extending to the disc cavity (13). Seal housing flanges (43, 44) extend from a seal housing (29) of the stator stage (25). Rotor flanges (41i, 41o) extend from a rotor disk (9-1). An inner rotor flange (41i) and first seal housing flange (43) are inward from a second seal housing flange (44). One rotor flange (41o) is outward from the second seal housing flange (44). The inner rotor flange (41i) and first seal housing flange (43) extend toward one another to limit movement of main gas flow (17). An inlet (47) injects air (50) between the outward rotor flange (41o) and second seal housing flange (44).

Abstract: A sealing system for a rotor assembly in a gas turbine engine is disclosed. The sealing system may include a seal formed from a side block and an upper seal that seals a gap between a radially outward extending first rotor supply channel in a rotor assembly terminating at an inlet of an axially extending second rotor supply channel that is in fluid communication with an internal blade cooling system of a turbine blade. The seal may include components that enhance the flow of cooling fluids over conventional configurations. In another embodiment, the sealing system may include an integrated sealing block configured to seal a gap between adjacent turbine blades at an intersection between the first and second rotor supply channels. The integrated sealing block may be formed from a radially inward extending leg and central body.