Cooled shroud assembly and method of cooling a shroud

1. A cooled shroud assembly for a gas turbine engine having a longitudinal centerline, said shroud assembly comprising:

at least one arcuate shroud segment surrounding a row of rotating turbine blades, said shroud segment having a forward flange, an aft flange defining an axially-facing aft edge, and an inwardly-facing flowpath surface, wherein said shroud segment lacks cooling holes for cooling said aft flange; and

at least one stationary turbine nozzle disposed axially next to said shroud segment, comprising:

an airfoil; and

an arcuate outer band disposed at a radially outer end of said airfoil and positioned axially adjacent to said shroud;

wherein a plurality of cooling holes including at least a first cooling hole and at least a second cooling hole are formed in said outer band and in fluid communication with a source of cooling air, said first cooling hole and said second cooling hole being positioned so as to impinge a flow of cooling air against said shroud segment and such that said first cooling hole and said second cooling hole converge at about the same place on said shroud segment.

2. The shroud assembly of claim 1 wherein at least one cooling hole of said plurality of cooling holes is oriented so as to impinge a flow of cooling air directly against said axially-facing aft edge of said shroud segment.

3. The shroud assembly of claim 2 wherein the at least one cooling hole of the plurality of cooling holes is disposed at an angle R in a radial plane towards said centerline axis.

4. The shroud assembly of claim 3 wherein said angle R is about 32 to about 36 degrees.

5. The shroud assembly of claim 2 wherein the at least one cooling hole of said plurality of cooling holes is disposed at an angle T in a tangential plane to direct a flow of cooling air at an outer aft corner of said shroud segment.

6. The shroud assembly of claim 5 wherein said angle T is about 48 to about 58 degrees.

7. The shroud assembly of claim 2 wherein the at least one cooling hole of said plurality of cooling holes is positioned so as to direct a flow of cooling air against a selected portion of an adjacent shroud segment.

8. A method of cooling a shroud for a gas turbine engine comprising:

providing at least one arcuate shroud segment surrounding a row of rotating turbine blades, said shroud segment having a forward end, an aft flange defining an aft edge, and an inwardly-facing flowpath surface, wherein said aft flange lacks cooling holes;

providing at least one stationary turbine nozzle disposed axially next to said shroud segment, said nozzle including:

an airfoil;

an arcuate outer band disposed at a radially outer end of said airfoil and disposed axially next to said shroud, said outer band defining a nozzle plenum, wherein gas flow over said airfoil generates a bow wave which causes localized heating of a portion of said aft flange;

directing a flow of cooling air from said nozzle plenum to said aft edge of said shroud segment so as to substantially negate said localized heating; and

forming a plurality of cooling holes including at least a first cooling hole and at least a second cooling hole in said shroud and having an inlet in fluid communication with said nozzle and an outlet in fluid communication with said aft edge, wherein said first cooling hole and said second cooling hole converge at about the same place on said shroud segment.

9. The method of claim 8 wherein at least one cooling hole of said plurality of cooling holes is oriented so as to direct a flow of cooling air directly against a selected portion of said aft end of said shroud segment.

10. The method of claim 9 wherein the at least one cooling hole of said plurality of cooling holes is disposed at an angle R in a radial plane towards said longitudinal axis.

11. The method of claim 10 wherein said angle R is about 32 to about 36 degrees.

12. The method of claim 11 wherein the at least one cooling hole of said plurality of cooling holes is disposed at an angle T in a tangential plane so as to direct a flow of cooling air at an outer aft corner of said shroud segment.

13. The method of claim 12 wherein said angle T is about 48 to about 58 degrees.

14. A cooled shroud assembly for a gas turbine engine having a longitudinal centerline, said shroud assembly comprising:

at least one arcuate shroud segment surrounding a row of rotating turbine blades, said shroud segment having a forward flange, an aft flange defining an axially-facing aft edge, and an inwardly-facing flowpath surface, wherein said shroud segment lacks cooling holes for cooling said aft flange; and

at least one stationary turbine nozzle disposed axially next to said shroud segment, comprising:

an airfoil; and

an arcuate outer band disposed at a radially outer end of said airfoil and positioned axially adjacent to said shroud;

wherein a plurality of cooling holes including at least a first cooling hole and at least a second cooling hole are formed in said outer band and in fluid communication with a source of cooling air, said first cooling hole and said second cooling hole being positioned so as to impinge a flow of cooling air on the aft flange of the shroud segment.

15. The shroud assembly of claim 14 wherein at least one cooling hole of said plurality of cooling holes is disposed at an angle R in a radial plane towards said centerline axis.

16. The shroud assembly of claim 15 wherein said angle R is about 32 to about 36 degrees.

17. The shroud assembly of claim 14 wherein the at least one cooling hole of said plurality of cooling holes is disposed at an angle T in a tangential plane to direct a flow of cooling air at an outer aft corner of said shroud segment.

18. The shroud assembly of claim 17 wherein said angle T is about 48 to about 58 degrees.