Method and system to facilitate enhanced local cooling of turbine engines

1. A method of assembling a gas turbine engine, said method comprising:

coupling at least one turbine nozzle segment within the gas turbine engine, wherein the at least one turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band that includes an aft flange and a radial inner surface;

coupling at least one turbine shroud segment downstream from the at least one turbine nozzle segment, wherein the at least one turbine shroud segment includes a leading edge and a radial inner surface;

coupling a cooling fluid source in flow communication with the at least one turbine nozzle segment such that cooling fluid channeled to each turbine nozzle outer band aft flange is directed at an oblique discharge angle towards the leading edge of the at least one turbine shroud segment; and

channeling the cooling fluid through at least a first group of cooling openings having a first aggregate cross-sectional area and a second group of cooling openings having a second aggregate cross-sectional area, the first aggregate cross-sectional area is larger than the second cross-sectional area, to facilitate preferential cooling of the turbine shroud.

2. A method in accordance with claim 1 wherein said coupling a cooling fluid source in flow communication with the at least one turbine nozzle segment further comprises coupling the cooling fluid source to the turbine nozzle segment such that cooling fluid is channeled into a gap defined between the aft flange and the leading edge.

3. A method in accordance with claim 1 wherein channeling the cooling fluid further comprises channeling cooling fluid through the first group of openings to facilitate enhanced cooling fluid flow to hotter portions of the shroud segment inner surface.

4. A method in accordance with claim 1 wherein channeling the cooling fluid further comprises channeling cooling fluid through the second group of openings to facilitate reducing cooling fluid flow to cooler portions of the shroud segment inner surface.

5. A method in accordance with claim 1 further comprising forming the first group of cooling openings from a plurality of cooling openings having a first diameter and forming the second group of cooling openings from a plurality of cooling openings having a second diameter, the first diameter is larger than the second diameter.

6. A method in accordance with claim 1 further comprising forming the first group of cooling openings from a first density of cooling openings and forming the second group of cooling openings from a second density of cooling openings, the first density of cooling openings is greater than the second density of cooling openings.

7. An engine assembly comprising:

a nozzle assembly comprising:

an inner band;

an outer band comprising an aft flange and a radial inner surface, said aft flange comprising a first group of cooling openings having a first aggregate cross-sectional area and a second group of cooling openings having a second aggregate cross-sectional area, said first aggregate cross-sectional area is larger than said second aggregate cross-sectional area, the cooling openings configured to direct cooling fluid therefrom at an oblique discharge angle; and

at least one airfoil vane extending between said inner band and said outer band.

8. An engine assembly in accordance with claim 7 further comprising a shroud assembly, said shroud assembly comprising a shroud inner surface and a shroud assembly leading edge, said shroud assembly leading edge comprising a first shroud group of cooling openings having a third aggregate cross-sectional area and a second shroud group of cooling openings having a fourth aggregate cross-sectional area, said third aggregate cross-sectional area is larger than said fourth aggregate cross-sectional area, said cooling openings configured to discharge cooling fluid therefrom at an oblique angle with respect to a centerline of a gap defined between said outer band aft flange and said shroud assembly leading edge.

9. An engine assembly in accordance with claim 8 wherein said first group of cooling openings and said first shroud group of cooling openings include a plurality of cooling openings having a first diameter.

10. An engine assembly in accordance with claim 9 wherein said second group of cooling openings and said second shroud group of cooling openings include a plurality of cooling openings having a second diameter, said first diameter is larger than said second diameter.

11. An engine assembly in accordance with claim 8 wherein said first group of openings and said first shroud group of openings include a plurality of densely oriented cooling holes.

12. An engine assembly in accordance with claim 11 wherein said second group of openings and said second shroud group of openings include a plurality of cooling holes less densely oriented than said first group of openings and said first shroud group of openings.

13. An engine assembly in accordance with claim 8 wherein said first group of cooling openings and said second group of cooling openings facilitate reducing hot gas ingestion into a gap defined between said outer band aft flange and said shroud assembly leading edge.

14. An engine assembly in accordance with claim 8 wherein said first group of cooling openings and said first shroud group of cooling openings facilitate film cooling of said shroud inner surface.

15. A gas turbine engine comprising a nozzle assembly comprising an inner band, an outer band, and at least one airfoil vane extending between said inner band and said outer band, said outer band comprising an aft flange and a radial inner surface, said aft flange comprising a first group of cooling openings having a first aggregate cross-sectional area and a second group of cooling openings having a second aggregate cross-sectional area, said first aggregate cross-sectional area is larger than said second aggregate cross-sectional area, said cooling openings configured to direct cooling fluid therefrom at an oblique discharge angle.

16. A gas turbine engine in accordance with claim 15 further comprising a shroud assembly, said shroud assembly comprising a shroud inner surface and a shroud assembly leading edge, said shroud assembly leading edge comprising a first shroud group of cooling openings having a third aggregate cross-sectional area and a second shroud group of cooling openings having a fourth aggregate cross-sectional area, said third aggregate cross-sectional area is larger than said fourth aggregate cross-sectional area, said cooling openings configured to discharge cooling fluid therefrom at an oblique angle with respect to a centerline of a gap defined between said outer band aft flange and said shroud assembly leading edge.

17. A gas turbine engine in accordance with claim 16 wherein said first group of cooling openings and said first shroud group of cooling openings include a plurality of cooling openings having a first diameter.

18. A gas turbine engine in accordance with claim 17 wherein said second group of cooling openings and said second shroud group of cooling openings include a plurality of cooling openings having a second diameter, said first diameter is larger than said second diameter.

19. A gas turbine engine in accordance with claim 16 wherein said first group of cooling openings facilitates reducing hot gas ingestion into a gap defined between said outer band aft flange and said shroud assembly leading edge.

20. A gas turbine engine in accordance with claim 16 wherein said first group of cooling openings and said first shroud group of cooling openings facilitate film cooling of said shroud inner surface.