Methods and apparatus for assembling gas turbine engines

• US 20060133922A1
• Filed: 12/20/2004
• Published: 06/22/2006
• Est. Priority Date: 12/20/2004
• Status: Active Grant

• Current Assignee: General Electric Company
• Original Assignee: General Electric Company
• Inventors: Ford, Gregg, Ervin, Elizabeth, Vonderau, Christen, Heyward, John
• Granted Patent: US 7,296,966 B2
• Time in Patent Office: Days
• US Class Current: 415/115
• CPC Class Codes: F01D 25/12 Cooling
• CPC Class Codes: F01D 9/042 fixing blades to stators fixing stator-rings in the casing or to each other F01D25/246
• CPC Class Codes: F05D 2230/60 Assembly methods
• CPC Class Codes: F05D 2240/40 Use of a multiplicity of similar components
• CPC Class Codes: F05D 2260/201 by impingement of a fluid
• CPC Class Codes: F05D 2260/202 by film cooling

Abstract

A method of assembling a gas turbine engine includes coupling at least one turbine nozzle segment within the gas turbine engine, each turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band, wherein the airfoil vane includes a leading edge and a trailing edge, and wherein the outer band includes a front face, a rear face, and an inner surface extending therebetween. The method also includes coupling at least one turbine shroud segment downstream from the at least one turbine nozzle segment, wherein each turbine shroud segment includes a front face, a rear face, and an inner surface extending therebetween, and coupling a cooling fluid source to each turbine nozzle segment such that cooling fluid may be channeled to each turbine nozzle inner surface proximate to one of the leading edge and the trailing edge of each airfoil vane, such that cooling fluid channeled to each turbine nozzle outer band rear face is directed towards the front face of at least one turbine shroud segment.

Claims

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

coupling at least one turbine nozzle segment within the gas turbine engine, each turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band, wherein the airfoil vane includes a leading edge and a trailing edge, and wherein the outer band includes a front face, a rear face, and an inner surface extending therebetween;

coupling at least one turbine shroud segment downstream from the at least one turbine nozzle segment, wherein each turbine shroud segment includes a front face, a rear face, and an inner surface extending therebetween; and

coupling a cooling fluid source to each turbine nozzle segment such that cooling fluid may be channeled to each turbine nozzle inner surface proximate to one of the leading edge and the trailing edge of each airfoil vane, such that cooling fluid channeled to each turbine nozzle outer band rear face is directed towards the front face of at least one turbine shroud segment.

2. A method in accordance with claim 1 wherein said coupling a cooling fluid source to each turbine nozzle segment further comprises coupling the cooling fluid source to each turbine nozzle segment such that cooling fluid may be channeled to a throat area defined between adjacent turbine nozzle airfoils.

3. A method in accordance with claim 1 wherein said coupling a cooling fluid source to each turbine nozzle segment further comprises coupling the cooling fluid source to each turbine nozzle segment such that cooling fluid may be channeled to each turbine nozzle outer band to facilitate impingement cooling of the front face of at least one turbine shroud.

4. A method in accordance with claim 1 wherein coupling at least one turbine nozzle segment within the gas turbine engine further comprises positioning cooling holes defined within each turbine nozzle outer band rear face in substantially alignment with a combustion gas flow path channeled through the turbine nozzle segment to facilitate reducing an operating temperature of the turbine shroud during engine operation.

5. A method in accordance with claim 1 wherein coupling at least one turbine shroud segment downstream from the at least one turbine nozzle segment further comprises:

positioning each turbine shroud segment such that a gap is defined between the rear face of each turbine nozzle outer band and the front face of each turbine shroud segment; and

applying a coating to the rear face of each turbine nozzle outer band to facilitate reducing a width of the gap.

6. A nozzle assembly comprising:

an inner band;

an outer band comprising a front face, a rear face, and an inner surface extending therebetween, said outer band rear face comprising a plurality of cooling holes configured to direct cooling fluid onto at least one turbine shroud segment, said inner surface comprising a plurality of cooling holes configured to facilitate cooling said inner surface; and

at least one airfoil vane extending between said inner band and said outer band, each said at least one airfoil vane comprising a first sidewall and a second sidewall connected at a leading edge and a trailing edge.

7. A nozzle assembly in accordance with claim 6 wherein said inner surface cooling holes are configured to facilitate film cooling of said inner surface.

8. A nozzle assembly in accordance with claim 6 wherein said inner surface cooling holes are positioned proximate a throat area defined between adjacent airfoil vanes.

9. A nozzle assembly in accordance with claim 6 wherein at least one of said inner surface cooling holes extend through said outer band proximate said airfoil leading edge, and wherein at least one of said inner surface cooling holes extends through said outer band proximate said trailing edge.

10. A nozzle assembly in accordance with claim 6 wherein said outer band rear face cooling holes facilitate impingement cooling of a front face of the at least one turbine shroud segment.

11. A nozzle assembly in accordance with claim 6 wherein said outer band rear face cooling holes are configured to direct cooling fluid onto every other turbine shroud segment to facilitate reducing an amount of cooling fluid channeled to said nozzle assembly.

12. A nozzle assembly in accordance with claim 6 wherein said outer band rear face comprises a coating of material, said coating facilitates reducing a width of a gap defined between said outer band rear face and a front face of the at least one turbine shroud segment.

13. 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, each said at least one airfoil vane comprising a first sidewall and a second sidewall connected at a leading edge and a trailing edge, said outer band comprising a front face, a rear face, and an inner surface extending therebetween, said outer band rear face comprising a plurality of cooling holes configured to direct cooling fluid onto at least one turbine shroud segment, said inner surface comprising a plurality of cooling holes configured to facilitate cooling said inner surface.

14. A gas turbine engine in accordance with claim 13 wherein said inner surface cooling holes are configured to facilitate film cooling of said inner surface.

15. A gas turbine engine in accordance with claim 13 wherein said inner surface cooling holes are positioned proximate a throat area defined between adjacent airfoil vanes.

16. A gas turbine engine in accordance with claim 13 wherein at least one of said inner surface cooling holes extend through said outer band proximate said airfoil leading edge, and wherein at least one of said inner surface cooling holes extends through said outer band proximate said trailing edge.

17. A gas turbine engine in accordance with claim 13 wherein said outer band rear face cooling holes facilitate impingement cooling of a front face of the at least one turbine shroud segment.

18. A gas turbine engine in accordance with claim 13 wherein said outer band rear face cooling holes are configured to direct cooling fluid onto every other turbine shroud segment to facilitate reducing an amount of cooling fluid channeled to said nozzle assembly.

19. A gas turbine engine in accordance with claim 13 wherein said outer band rear face comprises a coating of material, said coating facilitates reducing a width of a gap defined between said outer band rear face and a front face of the at least one turbine shroud segment.