System and method for fuel and steam injection within a combustor

Patent metadata

Field	Value
Publication number	US-9441835-B2
Application number	US-201213647359-A
Country	US
Kind code	B2
Filing date	Oct 8, 2012
Priority date	Oct 8, 2012
Publication date	Sep 13, 2016
Grant date	Sep 13, 2016

Abstract

A system includes a gas turbine combustor configured to combust a fuel and an oxidant, such as O 2 and O 2 mixtures. The system also includes an aerodynamic peg disposed in the gas turbine combustor. The aerodynamic peg includes a first passage configured to convey a first fluid into the gas turbine combustor and a second passage configured to convey a second fluid into the gas turbine combustor. The first fluid and second fluid are different from one another.

First claim

The invention claimed is:

1. A system, comprising:
a gas turbine combustor comprising a combustion region downstream from a head end region, wherein the head end region comprises one or more fuel nozzles, wherein the combustion region is configured to combust one or more fuels and an oxidant and flow combustion gases in a downstream direction away from the head end region, wherein the gas turbine combustor comprises a first wall disposed about at least one of the head end region or the combustion region, and a second wall disposed about the first wall to define a flow path configured to flow fluid in an upstream direction opposite the downstream direction into the head end region; and
an aerodynamic peg coupled to at least one of the first or second wall within the flow path in the gas turbine combustor, wherein the aerodynamic peg is oriented in a radial direction relative to an axis of the gas turbine combustor, wherein the aerodynamic peg comprises:
one or more non-oxidant, non-fuel outlets coupled to a non-oxidant, non-fuel passage, wherein the one or more non-oxidant, non-fuel outlets are configured to convey only a non-oxidant, non-fuel fluid into the flow path of the gas turbine combustor;
and
one or more fuel outlets coupled to a fuel passage, wherein the one or more fuel outlets are configured to convey only a fuel into the flow path of the gas turbine combustor, wherein the one or more non-oxidant, non-fuel outlets are disposed upstream from the one or more fuel outlets.

2. The system of claim 1 , wherein the aerodynamic peg comprises an airfoil shaped cross-section.

3. The system of claim 2 , wherein the one or more non-oxidant, non-fuel outlets and the one or more fuel outlets are disposed between a leading edge and a trailing edge of the airfoil shaped cross-section.

4. The system of claim 3 , wherein the aerodynamic peg comprises:
a first side extending between the leading edge and the trailing edge, wherein the first side comprises one or more fuel outlets and one or more non-oxidant, non-fuel outlets; and
a second side between the leading edge and the trailing edge, opposite the first side, wherein the second side comprises one or more fuel outlets and one or more non-oxidant, non-fuel outlets.

5. The system of claim 1 , comprising a manifold disposed surrounding the gas turbine combustor and coupled to the non-oxidant, non-fuel passage and the fuel passage, wherein the manifold is configured to convey the non-oxidant, non-fuel fluid and the fuel to the non-oxidant, non-fuel passage and the fuel passage.

6. The system of claim 5 , wherein the manifold comprises a non-oxidant, non-fuel manifold configured to convey the non-oxidant, non-fuel fluid and a fuel manifold configured to convey the fuel.

7. The system of claim 1 , wherein the first wall is disposed about the combustion region, and the aerodynamic peg is disposed within the flow path about the combustion region.

8. The system of claim 1 , wherein the aerodynamic peg is disposed about the head end region.

9. The system of claim 7 , wherein the gas turbine combustor comprises:
a plurality of aerodynamic pegs, including the aerodynamic peg, equidistantly spaced circumferentially from one another, wherein each of the plurality of aerodynamic pegs is coupled to at least one of the first or second wall within the flow path.

10. The system of claim 7 , wherein the aerodynamic peg extends only partially across the flow path between the first and second walls.

11. The system of claim 1 , wherein the non-oxidant, non-fuel fluid comprises steam.

12. The system of claim 1 , wherein the non-oxidant, non-fuel fluid comprises nitrogen.

13. The system of claim 1 , comprising a gas turbine engine having the gas turbine combustor.

14. The system of claim 1 , comprising a third passage disposed in the aerodynamic peg and configured to convey a third fluid into the gas turbine combustor, wherein the non-oxidant, non-fuel fluid, the fuel, and the third fluid are different from one another.

15. The system of claim 1 , comprising a fuel supply and a non-oxidant, non-fuel fluid supply, wherein the one or more fuel outlets are coupled to the fuel supply and are not coupled to the non-oxidant, non-fuel fluid supply, wherein the one or more non-oxidant, non-fuel outlets are coupled to the non-oxidant, non-fuel fluid supply and are not coupled to the fuel supply.

16. A system, comprising:
a gas turbine engine, comprising:
a turbine: and
a gas turbine combustor, wherein the gas turbine combustor comprises:
a first wall disposed about a combustion region wherein the combustion region is configured to combust one or more fuels and an oxidant downstream from a head end region and the combustion region is configured to flow combustion gases in a downstream direction away from the head end region;
a second wall disposed about the first wall to define a flow path configured to flow fluid in an upstream direction opposite the downstream direction toward the head end region; and
an aerodynamic peg coupled to at least one of the first or second wall within the flow path, wherein the aerodynamic peg is oriented in a radial direction relative to an axis of the gas turbine combustor, wherein the aerodynamic peg comprises:
one or more non-oxidant, non-fuel outlets coupled to a non-oxidant, non-fuel passage, wherein the one or more non-oxidant, non-fuel outlets are configured to convey only a non-oxidant, non-fuel fluid into the flow path of the gas turbine combustor; and
one or more fuel outlets coupled to a fuel passage, wherein the one or more fuel outlets are configured to convey only a fuel into the flow path of the gas turbine combustor, wherein the one or more non-oxidant, non-fuel outlets are disposed upstream of the one or more fuel outlets.

17. The system of claim 16 , wherein the non-oxidant, non-fuel passage is coupled to a non-oxidant, non-fuel manifold to supply the non-oxidant, non-fuel fluid, and the fuel passage is coupled to a fuel manifold to supply the fuel.

18. The system of claim 16 , wherein the aerodynamic peg is coupled to the second wall and extends toward the first wall of the gas turbine combustor, the first wall comprises a combustion liner, and the second wall comprises a flow sleeve configured to receive a compressed flow from a compressor.

19. The system of claim 16 , wherein the aerodynamic peg extends only partially across the flow path between the first and second walls.

20. The system of claim 16 , wherein the aerodynamic peg comprises an airfoil shaped cross-section extending from a leading edge in a downstream direction to a trailing edge, wherein the one or more non-oxidant, non fuel outlets and the one or more fuel outlets are disposed between the leading edge and the trailing edge.

21. The system of claim 20 , wherein the aerodynamic peg comprises:
a first side extending between the leading edge and the trailing edge, wherein the first side comprises one or more fuel outlets and one or more non-oxidant, non-fuel outlets; and
a second side between the leading edge and the trailing edge, opposite the first side, wherein the second side comprises one or more fuel outlets and one or more non-oxidant, non-fuel outlets.

22. A method, comprising:
injecting only a non-oxidant, non-fuel fluid through one or more non-oxidant, non-fuel outlets disposed in an aerodynamic peg into a gas turbine combustor, the gas turbine combustor comprising a combustion region downstream from a head end region, where

Assignees

• Gen Electric

Inventors

• Overby Brandon Taylor
• Leach David

Classifications

• F23R3/286

  having fuel-air premixing devices (F23R3/30 takes precedence) · CPC title

• F23R3/02 Primary

  characterised by the air-flow or gas-flow configuration (reverse- flow combustion chambers F23R3/54; cyclone or vortex type combustion chambers F23R3/58) · CPC title

• F23L2900/07002

  Injecting inert gas, other than steam or evaporated water, into the combustion chambers · CPC title

• F23L2900/07009

  Injection of steam into the combustion chamber · CPC title

• F23L7/00 Primary

  Supplying non-combustible liquids or gases, other than air, to the fire, e.g. oxygen, steam · CPC title