Methods and system for recuperated circumferential cooling of 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 an inner surface and a leading edge that is substantially perpendicular to the inner surface;

coupling a turbine nozzle to the turbine shroud segment such that a gap is defined between an aft edge of an outer band of the turbine nozzle and the leading edge;

directing cooling air into the gap such that the cooling air is substantially circumferentially mixed in a plenum defined within the leading edge; and

directing the cooling air in the gap through at least one cooling hole formed between the plenum and the inner surface.

2. A method in accordance with claim 1 wherein directing the cooling air in the gap through at least one cooling hole formed between the plenum and the inner surface further comprises directing the cooling air toward the inner surface to facilitate forming a recuperated film cooling layer across the inner surface.

3. 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 at least one forwardly directed cooling opening defined in the plenum and positioned radially outward from the at least one cooling hole extending between the plenum and the inner surface.

4. A method in accordance with claim 1 wherein directing cooling air into the gap further comprises directing cooling air into the gap such that the air is substantially uniformly distributed throughout the gap from the plenum.

5. A method in accordance with claim 1 wherein a lip is formed on the aft edge and extends downstream from the aft edge, said directing cooling air into the gap further comprises:

metering spent turbine nozzle cooling air into the gap through at least one discharge opening extending through the aft edge; and

directing the spent turbine nozzle cooling air in the gap against the lip and towards the shroud segment to facilitate cooling the shroud segment.

6. A method in accordance with claim 5 wherein the shroud segment includes a rounded corner portion extending between the leading edge and the inner surface, said method further comprises directing the spent turbine nozzle cooling air along the rounded corner portion to facilitate forming a film cooling layer across the inner surface.

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

8. 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 and an inner surface that is substantially perpendicular to said leading edge; and

a turbine nozzle comprising an outer band comprising an aft edge, said turbine nozzle is upstream from said shroud segment and is coupled with said shroud segment such that a gap is defined between said aft edge and said leading edge, said shroud segment comprises:

a circumferential plenum defined within said leading edge, said plenum is configured to substantially uniformly distribute cooling air circumferentially throughout the gap; and

at least one cooling hole formed between said plenum and said inner surface, said at least one cooling hole is configured to direct the cooling air in said gap toward the hot gas flow path.

9. A turbine nozzle and shroud assembly in accordance with claim 8 wherein said at least one cooling hole is further configured to direct cooling air in the gap towards said inner surface to facilitate forming a recuperated film cooling layer across said inner surface.

10. A turbine nozzle and shroud assembly in accordance with claim 8 further comprising at least one forwardly directed cooling opening defined in said plenum, said at least one forwardly directed cooling opening is radially outward from said at least one cooling hole formed between said plenum and said inner surface and is configured to meter cooling air into said gap.

11. A turbine nozzle and shroud assembly in accordance with claim 8 further comprising a duct at least partially defined by said turbine nozzle and said turbine shroud segment, said duct configured to direct cooling air into said gap.

12. A turbine nozzle and shroud assembly in accordance with claim 8 further comprising:

a lip formed on said aft edge, said lip extends downstream from said gap; and

at least one discharge opening extending through said aft edge, said at least one discharge opening configured to meter spent turbine nozzle cooling air into said gap, such that said lip directs the spent turbine nozzle cooling air from said gap towards said shroud segment to facilitate cooling said shroud segment.

13. A turbine nozzle and shroud assembly in accordance with claim 12 further comprising a rounded corner portion extending between said leading edge and said inner surface, said rounded corner portion configured to direct the spent turbine nozzle cooling air towards said inner surface to facilitate forming a film cooling layer against said inner surface.

14. A turbine nozzle and shroud assembly in accordance with claim 8 wherein said gap is pressurized to facilitate minimizing a turbine nozzle wake effect within said gap.

15. A cooling system for a gas turbine engine that includes a shroud segment having a leading edge and an inner surface that is substantially perpendicular to the leading edge, and a turbine nozzle that is upstream from the shroud segment and that includes an outer band having an aft edge, said cooling system comprising:

a plenum defined within the leading edge, said plenum is configured to substantially uniformly distribute cooling air through a gap defined between the aft edge and the leading edge; and

at least one cooling hole formed between said plenum and the inner surface, said at least one cooling hole is configured to direct the cooling air from the gap towards a hot gas flow path defined within the gas turbine engine.

16. A cooling system in accordance with claim 15 further configured to direct cooling air from the gap through said at least one cooling hole extending between said plenum and the inner surface to facilitate forming a recuperated film cooling layer across the inner surface of the shroud segment.

17. A cooling system in accordance with claim 15 further comprising at least one forwardly directed cooling opening defined in said plenum, said at least one forwardly directed cooling opening is radially outward from said at least one cooling hole extending between said plenum and the inner surface, said at least one forwardly directed cooling opening is configured to meter cooling air into the gap.

18. A cooling system in accordance with claim 15 further comprising a duct at least partially defined by the turbine nozzle and the turbine shroud segment, said duct configured to direct cooling air into the gap.

19. A cooling system in accordance with claim 15 further comprising:

a lip formed on the outer band aft edge, said lip extends downstream from the gap; and

at least one discharge opening extending through the aft edge, said at least one discharge opening is configured to meter spent turbine nozzle cooling air into the gap, said lip is configured to direct the spent turbine nozzle cooling air to facilitate film cooling the shroud segment.

20. A cooling system in accordance with claim 19 further comprising a rounded corner portion extends between the leading edge and the inner surface, said rounded corner portions configured to direct spent turbine nozzle cooling air towards the inner surface of the shroud segment to facilitate forming a film cooling layer across the inner surface.