Methods and system for cooling integral turbine nozzle and shroud assemblies
AssigneeGeneral Electric Company
InventorsChing-Pang Lee, Eric Alan Estill, James Harvey Laflen, Daniel Verner Jones
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Abstract
A method for cooling a shroud segment of a gas turbine engine includes providing a turbine shroud assembly including a shroud segment having a leading edge defining a forward face. A turbine nozzle is coupled to the turbine shroud assembly such that a gap is defined between an aft face of an outer band of the turbine nozzle and the forward face, wherein a lip formed on the aft face is positioned radially inwardly with respect to the gap and extends substantially axially downstream from the gap. Cooling air is directed into the gap. Cooling air exiting the gap impinges against the lip to facilitate film cooling the shroud segment.
Claims20
A method for cooling a shroud segment of a gas turbine engine, said method comprising:
- providing a turbine shroud assembly including a shroud segment having a leading edge defining a forward face;
- coupling a turbine nozzle to the turbine shroud assembly such that a gap is defined between an aft face of an outer band of the turbine nozzle and the forward face, wherein a lip formed on the aft face is positioned radially inwardly with respect to the gap and extends substantially axially downstream from the gap;
- directing cooling air into the gap;
- and
- impinging cooling air exiting the gap against the lip to facilitate film cooling the shroud segment.
A turbine nozzle and shroud assembly for a gas turbine engine, said turbine nozzle and shroud assembly comprising:
- a shroud segment comprising a leading edge defining a forward face of said shroud segment;
- and
- a turbine nozzle comprising an outer band having a trailing edge defining an aft face of said outer band, said turbine nozzle upstream from said shroud segment and coupled with said shroud segment such that a gap is defined between said aft face and said forward face, said gap is configured to direct cooling air towards a hot gas flow path flowing through the gas turbine engine, said aft face comprises a lip formed thereon radially inward with respect to said gap and extending axially downstream of said gap.
A cooling system for a gas turbine engine, the gas turbine engine comprising a shroud segment having a leading edge defining a forward face, and a turbine nozzle comprising an outer band having a trailing edge defining an aft face, the turbine nozzle positioned upstream of the shroud segment and coupled with a turbine shroud assembly such that a gap is defined between the aft face and the forward face, wherein a lip formed on the aft face is positioned radially inwardly with respect to the gap and extends substantially axially downstream from the gap, said cooling system configured to:
- direct cooling air through the gap and towards a hot gas flow path flowing through the gas turbine engine substantially parallel to a central axis of the gas turbine engine;
- and
- impinge cooling air exiting the gap against the lip to facilitate film cooling the shroud segment.
Dependencies are inferred from the claim text.