Combustor for a low-emissions gas turbine engine

• US 6,098,397 A
• Filed: 06/08/1998
• Issued: 08/08/2000
• Est. Priority Date: 06/08/1998
• Status: Expired due to Fees

• Current Assignee: Solar Turbines Incorporated (Caterpillar Incorporated)
• Original Assignee: Caterpillar Incorporated
• Inventors: Dutta, Partha, Glezer, Boris, Moon, Hee-Koo, Greenwood, Stuart A.
• Primary Examiner(s): FREAY, CHARLES GRANT
• Application Number:
  US09/093,375
• Time in Patent Office: 792 Days
• Field of Search: 60/39.02, 60/39.32, 60/752, 60/753, 60/754, 60/756, 60/757, 60/760, 60/758
• US Class Current: 60/772
• CPC Class Codes: F05B 2260/222 by creating turbulence
• CPC Class Codes: F23R 2900/03045 Convection cooled combustion chamber walls provided with turbolators or means for creating turbulences to increase cooling
• CPC Class Codes: F23R 3/002 Wall structures F23R3/02 and F23R3/007 take precedence
• CPC Class Codes: F23R 3/005 Combined with pressure or heat exchangers
• CPC Class Codes: F23R 3/02 characterised by the air-flow or gas-flow configuration reverse- flow combustion chambers F23R3/54; cyclone or vortex type combustion chambers F23R3/58

Abstract

Many government entities regulated emission from gas turbine engines including CO. CO production is generally reduced when CO reacts with excess oxygen at elevated temperatures to form CO2. Many manufactures use film cooling of a combustor liner adjacent to a combustion zone to increase durability of the combustion liner. Film cooling quenches reactions of CO with excess oxygen to form CO2. Cooling the combustor liner on a cold side (backside) away from the combustion zone reduces quenching. Furthermore, placing a plurality of concavities on the cold side enhances the cooling of the combustor liner. Concavities result in very little pressure reduction such that air used to cool the combustor liner may also be used in the combustion zone. An expandable combustor housing maintains a predetermined distance between the combustor housing and combustor liner.

Claims

1. A combustor for a gas turbine engine, said combustor comprising:

a combustor cooling shield;

a combustor liner having an inlet portion and an outlet portion, said combustor liner being positioned within said combustor cooling shield, said combustor liner being connected with said combustor cooling shield at said outlet portion, said combustor liner having a hot side and a cold side, said cold side and said combustor cooling shield defining a cooling channel therebetween, said hot side defining a combustion zone therein, said combustion zone being adapted to receive compressed air and a fuel at said inlet portion, said combustion zone being adapted to exhaust a combustion gas into a turbine being in fluid communication with said outlet portion, said cooling channel being.adapted to receive a compressed air stream; and

a plurality of concavities disposed on said cold side, said concavities being adapted to increase convective cooling of said combustor liner.

2. The combustor of claim 1 wherein said cooling channel being adapted to receive compressed air intermediate said inlet portion and said outlet portion.

3. The combustor of claim 1 wherein said cooling channel being fluidly connected with said combustion zone proximate said inlet portion.

4. The combustor of claim 1 wherein said combustor liner being a nickel-base alloy.

5. The combustor of claim 1 wherein said hot side being treated with a thermal barrier coating being adapted to thermally insulate said hot side from said combustion zone.

6. The combustor of claim 5 wherein said thermal barrier coating being a zirconia-base material.

7. The combustor of claim 6 wherein said thermal barrier coating being applied by a plasma spray.

8. The combustor of claim 7 wherein said thermal barrier coating being about 0.010 inches thick.

9. The combustor of claim 1 wherein said combustor cooling shield being formed from a plurality of circumferential segments further comprising:

a resilient radial spacer being engagingly connectable with said circumferential segments and said combustor liner, said spacer being adapted to maintain a predetermined distance between said circumferential segments and said combustor liner; and

a resilient band being connectable with said combustor cooling shield, said resilient band being adapted to maintain connection between said circumferential segments and said radial spacer, said resilient band being adapted to maintain connection between said spacer and said combustor liner.

10. The combustor of claim 9 wherein said combustor is an annular combustor.

11. The combustor of claim 1 wherein each of said concavities being equally spaced from an adjacent concavity.

12. The combustor of claim 11 wherein said equal spacing being about 0.275 inches.

13. The combustor of claim 1 wherein said concavities extending into said cold side about 0.0415 inches.

14. The combustor of claim 1 wherein said concavities having a diameter of about 0.22 inches.

15. A method for improved cooling of a combustor for a gas turbine engine comprising the steps of:

positioning a combustor liner having a cold side inside a combustor cooling shield with said cold side facing said combustor cooling shield;

establishing a predetermined distance between said combustor cooling shield and said cold side, said predetermined distance, said cooling shield, and said cold side defining a cooling channel; and

maintaining said predetermined distance in response to expansion and contraction of said combustor liner.

16. The method for improved cooling of claim 15 further comprising the step of interrupting a growing thermal boundary layer on said cold side.

17. The method for improved cooling of claim 16 wherein said boundary layer growth being interrupted by a plurality of concavities on said cold side.

18. The method for improved cooling of claim 16 wherein said concavities being formed on said cold side by a stamping process.

19. The method for improved cooling of claim 15 wherein said predetermined distance being established by positioning a resilient radial spacer between said cold side and said combustor housing.

20. The method for improved cooling of claim 15 wherein said maintaining said predetermined distance being constraining a plurality of circumferential combustor cooling shield segments with a resilient band.

21. The method for improved cooling of claim 15 wherein said establishing said predetermined distance being forming a plurality of indentations in said combustor cooling shield extending to said cold side.

22. A method for reducing emissions of a gas turbine engine comprising the steps of:

directing a volume of air having a first pressure to a combustor, said combustor having a combustor cooling shield, a combustor liner, and a cooling channel between said combustor cooling shield and said combustor liner, said combustor liner having an inlet portion, an outlet portion, and a plurality of concavities adjacent said combustor cooling shield, said concavities being adapted to retard growth of a thermal boundary layer;

diverting a first portion of said volume of air into said cooling channel intermediate said inlet portion and said outlet portion;

diverting a remainder of said volume of air into said inlet portion;

passing said first portion over said concavities, said first portion convectively cooling said combustor liner; and

directing said first portion into said inlet portion, said first portion being at a second pressure wherein said second pressure being about equal to said first pressure.

23. The method for reducing emissions of claim 22 further comprising the step of directing said first portion through a dilution duct proximate said outlet portion.

24. The method for reducing emissions of claim 22 further comprising the step of adjusting said combustor cooling shield to maintain a predetermined distance between said combustor cooling shield and said combustor liner.