Methods and system for shielding cooling air to facilitate cooling integral turbine nozzle and shroud assemblies

1. 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;

impinging cooling air exiting the gap against the lip;

directing post impingement cooling air at the forward face to facilitate forming a film cooling layer on the shroud segment; and

shielding the film cooling layer from combustion gases flowing through the gas turbine engine.

2. A method in accordance with claim 1 further comprising pressurizing the gap to facilitate minimizing a turbine nozzle wake effect within the gap.

3. A method in accordance with claim 1 wherein shielding the film cooling layer from combustion gases further comprises metering a flow of cooling air through a plurality of discharge openings defined in the outer band.

4. A method in accordance with claim 3 wherein directing cooling air into the gap further comprises:

directing cooling air through a plurality of forwardly directed cooling openings defined in the leading edge; and

metering a flow of cooling air into the gap through the plurality of cooling openings.

5. A method in accordance with claim 1 wherein shielding the film cooling layer from combustion gases further comprises metering a flow of cooling air through a plurality of discharge openings extending through the lip and positioned radially inwardly with respect to the film cooling layer.

6. A method in accordance with claim 5 wherein metering a flow of cooling air through a plurality of discharge openings extending through the lip further comprises discharging cooling air through the plurality of discharge openings substantially parallel to the combustion gases flow path.

7. A method in accordance with claim 1 wherein directing cooling air into the gap further comprises metering a flow of cooling air into the gap through a plurality of forwardly directed cooling openings defined in the leading edge.

8. A method in accordance with claim 1 wherein directing cooling air into the gap further comprises directing leakage air into the gap from a duct that is at least partially defined by the turbine nozzle and the turbine shroud assembly.

9. A turbine nozzle and shroud assembly for a gas turbine engine, the gas turbine engine defining a central axis, said turbine nozzle and shroud assembly comprising:

a shroud segment comprising a leading edge defining a forward face of said shroud segment;

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 combustion gases flowing through the gas turbine engine, said aft face comprises a lip formed therein radially inward with respect to said gap and extending axially downstream from said gap; and

a plurality of discharge openings defined in said lip, said plurality of discharge openings configured to meter a flow of cooling air with respect to said gap.

10. A turbine nozzle and shroud assembly in accordance with claim 9 wherein said plurality of discharge openings are configured to facilitate shielding a film cooling layer formed on an inner surface of said shroud segment from the hot gas flow path.

11. A turbine nozzle and shroud assembly in accordance with claim 9 wherein said plurality of discharge openings are positioned radially inward with respect to a film cooling layer formed on an inner surface of said shroud segment.

12. A turbine nozzle and shroud assembly in accordance with claim 9 wherein said plurality of discharge openings extend substantially parallel to a central axis of the gas turbine engine, said plurality of discharge openings meter a flow of spent cooling air parallel to the hot gas flow path.

13. A turbine nozzle and shroud assembly in accordance with claim 9 further comprising a rounded corner portion formed on said leading edge, said rounded corner portion partially defines said gap and is configured to facilitate forming a film cooling layer on an inner surface of said shroud segment.

14. A turbine nozzle and shroud assembly in accordance with claim 9 further comprising a plurality of forwardly directed cooling openings defined in said leading edge, said plurality of cooling openings configured to meter a flow of edge cooling air into said gap.

15. A turbine nozzle and shroud assembly in accordance with claim 14 wherein at least one of said plurality of cooling openings is substantially parallel to at least one of said plurality of discharge openings defined in said lip.

16. 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 upstream of the shroud segment and coupled with the shroud segment such that a gap is defined between the aft face and the forward face, the aft face forming a lip radially inward with respect to the gap and extending substantially axially downstream from the gap, said cooling system configured to:

direct cooling air through the gap and radially inward towards combustion gases flowing through the gas turbine engine;

impinge cooling air exiting the gap against the lip to facilitate forming a film cooling layer along the shroud segment; and

shield the film cooling layer from combustion gases.

17. A cooling system in accordance with claim 16 further configured to meter cooling air through a plurality of discharge openings defined in the outer band radially inward with respect to the film cooling layer.

18. A cooling system in accordance with claim 16 further configured to meter cooling air through a plurality of discharge openings extending through the lip radially inward with respect to the film cooling layer.

19. A cooling system in accordance with claim 18 wherein the plurality of discharge openings are configured to meter spent cooling air parallel to a combustion gases flow path.

20. A cooling system in accordance with claim 16 further configured to pressurize the gap to facilitate minimizing a turbine nozzle wake effect within the gap.