Method and apparatus for cooling combustor liner and transition piece of a gas turbine

• US 7,493,767 B2
• Filed: 04/19/2005
• Issued: 02/24/2009
• Est. Priority Date: 06/01/2004
• Status: Active Grant

• Current Assignee: GE Infrastructure Technology, LLC (General Electric Company)
• Original Assignee: General Electric Company
• Inventors: Widener, Stanley Kevin, Bailey, Jeremy Clyde, Intile, John C, Bunker, Ronald Scott, Johnson, Thomas Edward
• Primary Examiner(s): Rodriguez; William H
• Application Number:
  US10/907,866
• Publication Number:
  US 20050268615A1
• Time in Patent Office: 1,407 Days
• Field of Search: 60752-760, 608/04, 60/393.7
• US Class Current: 60/752
• CPC Class Codes: F23R 2900/03042 Film cooled combustion chamber walls or domes
• CPC Class Codes: F23R 2900/03044 Impingement cooled combustion chamber walls or subassemblies
• CPC Class Codes: F23R 3/002 Wall structures F23R3/02 and F23R3/007 take precedence
• CPC Class Codes: F23R 3/06 Arrangement of apertures along the flame tube

Abstract

A method and apparatus for cooling a combustor liner and transitions piece of a gas turbine include a combustor liner with a plurality of turbulators arranged in an array axially along a length defining a length of the combustor liner and located on an outer surface thereof; a first flow sleeve surrounding the combustor liner with a first flow annulus therebetween, the first flow sleeve having a plurality of rows of cooling holes formed about a circumference of the first flow sleeve for directing cooling air from the compressor discharge into the first flow annulus; a transition piece connected to the combustor liner and adapted to carry hot combustion gases to a stage of the turbine; a second flow sleeve surrounding the transition piece a second plurality of rows of cooling apertures for directing cooling air into a second flow annulus between the second flow sleeve and the transition piece; wherein the first plurality of cooling holes and second plurality of cooling apertures are each configured with an effective area to distribute less than 50% of compressor discharge air to the first flow sleeve and mix with cooling air from the second flow annulus.

Claims

1. A combustor for a turbine comprising:

a combustor liner including a plurality of turbulators arranged in an array axially along a length defining a length of said combustor liner and located on an outer surface thereof;

a first flow sleeve surrounding said combustor liner with a first flow annulus therebetween,said first flow sleeve having a plurality of rows of cooling holes formed about a circumference of said first flow sleeve for directing cooling air from compressor discharge air into said first flow annulus, and said cooling holes are configured as non penetrating fluid jets providing at least one of bulk flow mixing and turbulence increasing heat transfer from the liner;

a transition piece connected to said combustor liner, said transition piece adapted to carry hot combustion gases to a stage of the turbine;

a second flow sleeve surrounding said transition piece, said second flow sleeve having a second plurality of rows of cooling apertures and a plurality of flow catcher devices disposed on a surface thereof for directing cooling air from compressor discharge air into a second flow annulus between the second flow sleeve and the transition piece, said first flow annulus connecting to said second flow annulus;

wherein said plurality of cooling holes said plurality of cooling apertures are said plurality of flow catcher devices are each configured with an effective area to distribute less then 50%of compressor discharge air to said first flow sleeve and mix with cooling are from said second flow annulus.

2. The combustor of claim 1, wherein said liner is one of a cast alloy liner and a wrought alloy liner.

3. The combustor of claim 1, wherein said plurality of cooling holes said plurality of cooling apertures and said plurality of flow catcher devices are each configured with an effective area to distribute between about 25% to about 40% of compressor discharge air to said first flow sleeve and mix with cooling air from said second flow annulus.

4. The combustor of claim 1, wherein said plurality of rows of cooling holes are substantially uniformly dimensioned.

5. The combustor of claim 1, wherein the non penetrating fluid jets are configured to avoid actual fluid impingement on the liner.

6. The combustor of claim 1, wherein said plurality of rows of cooling holes are configured providing mass velocity ratios (Gc/Gjet) near unity.

7. The combustor of claim 1, wherein said cooling holes are disposed about a circumference of said first flow sleeve in an in-line manner.

8. The combustor of claim 1, wherein said plurality of rows of cooling holes are dimensioned providing mass velocity ratios (Gc/Gjet) near unity.

9. A turbine engine comprising:

a combustion section;

a compressor air discharge section upstream of the combustion section;

a transition region between the combustion and air discharge section;

a turbulated combustor liner defining a portion of the combustion section and transition region, said turbulated combustor liner including a plurality of turbulators arranged in an array axially along a length defining a length of said combustor liner and located on an outer surface thereof;

a first flow sleeve surrounding said combustor liner with a first flow annulus therebetween, said first flow sleeve having a plurality of rows of cooling holes formed about a circumference of said first flow sleeve for directing cooling air from compressor discharge air into said first flow annulus, and said cooling holes are configured as non penetrating fluid jets providing at least one of bulk flow mixing and turbulence increasing heat transfer from the liner;

a transition piece connected to at least one of said combustor liner and said first flow sleeve, said transition piece adapted to carry hot combustion gases to a stage of the turbine corresponding to the air discharge section;

a second flow sleeve surrounding said transition piece, said second flow sleeve having a plurality of rows of cooling apertures and a plurality of flow catcher devices disposed on a surface thereof for directing said cooling air into a second flow annulus between the second flow sleeve and the transition piece, said first flow annulus connecting to said second flow annulus;

wherein said plurality of cooling holes, and said plurality of cooling apertures and said plurality of flow catcher devices are each configured with an effective area to distribute less than 50% of compressor discharge air to said first flow sleeve and mix with cooling air from said second flow annulus.

10. The engine of claim 9, wherein said first plurality of cooling holes, and second plurality of cooling apertures and said plurality of flow catcher devices are each configured with an effective area to distribute between about 25% to about 40% of compressor discharge air to said first flow sleeve and mix with cooling air from said second flow annulus.

11. The engine of claim 9, wherein said plurality of rows of cooling holes are substantially uniformly dimensioned.

12. The engine of claim 9, wherein the non penetrating fluid jets are configured to avoid actual fluid impingement on the liner.

13. The engine of claim 9, wherein said plurality of rows of cooling holes are configured providing mass velocity ratios (Gc/Gjet) near unity.

14. The engine of claim 9, wherein said cooling holes are disposed about the circumference of said first flow sleeve in an in-line manner.

15. The engine of claim 9, wherein said plurality of rows of cooling holes are dimensioned providing mass velocity ratios (Gc/Gjet) near unity.

16. A method of cooling a combustor liner of a gas turbine combustor, said combustor liner having a substantially circular cross-section, and a first flow sleeve surrounding said liner in substantially concentric relationship therewith creating a first flow annulus therebetween for feeding air to the gas turbine combustor, and wherein a transition piece is connected to said combustor liner, with the transition piece surrounded by a second flow sleeve, thereby creating a second flow annulus in communication with said first flow annulus;

the method comprising;

providing a plurality of axially spaced rows of cooling holes in said flow sleeves, each row extending circumferentially around said flow sleeves, a first of said rows in said second sleeve is located proximate an end where said first flow sleeve and said second flow sleeve interface;

providing a plurality of axially spaced rows of flow catcher devices in said second flow sleeve, each row extending circumferentially around at least a portion of said second flow sleeve;

supplying cooling air from compressor discharge to said cooling holes;

configuring said cooling holes as non penetrating fluid jets providing at least one of bulk flow mixing and turbulence increasing heat transfer from the liner, and configuring said flow catcher devices to aerodynamically cooperate with said cooling holes in said second flow sleeve, the cooling holes having and effective area and the flow catcher devices having an effective aerodynamic profile to distribute less than one half of compressor discharge air to said first flow sleeve and mix with the cooling air flowing from said second flow annulus.

17. The method combustor of claim 16, further comprising:

configuring said cooling holes with an effective area and said flow catcher devices with an effective aerodynamic profile to distribute between about 25% to about 40% of compressor discharge air to said first flow sleeve and mix with cooling air from said second flow annulus.

18. The method of claim 16, further comprising disposing said cooling holes about a circumference of said first flow sleeve in an in-line manner,

wherein said cooling holes are substantially uniformly dimensioned.

19. The method of claim 16, wherein the non penetrating fluid jets are configured to avoid actual fluid impingement on the liner.

20. The method of claim 16, wherein said cooling holes are at least one of configured and dimensioned providing mass velocity ratios (Gc/Gjet) near unity.