Staged fuel and air injection in combustion systems of gas turbines

Patent metadata

Field	Value
Publication number	US-9976487-B2
Application number	US-201514977993-A
Country	US
Kind code	B2
Filing date	Dec 22, 2015
Priority date	Dec 22, 2015
Publication date	May 22, 2018
Grant date	May 22, 2018

Abstract

A gas turbine that includes a working fluid flowpath extending aftward from a forward injector in a combustor. The combustor may include an inner radial wall, an outer radial wall, and, therebetween, a flow annulus. A staged injector may intersect the flow annulus so to attain an injection point within the working fluid flowpath by which aftward and forward annulus sections are defined. Air directing structure may include an aftward intake section that corresponds to the aftward annulus section and a forward intake section that corresponds to the forward annulus section. The air directing structure may be configured to: direct air entering through the aftward intake section through the aftward annulus section in a forward direction to the staged injector; and direct air entering through the forward intake section through the forward annulus section in an aftward direction to the staged injector.

First claim

That which is claimed:

1. A gas turbine that comprises:
a combustor coupled to a turbine that together define a working fluid flowpath, the working fluid flowpath extending aftward along a longitudinal axis from a forward end defined by a forward injector in the combustor, through an interface at which the combustor transitions to the turbine, and then through the turbine to an aftward end defined therein, wherein the combustor includes an inner radial wall, which defines the working fluid flowpath, and an outer radial wall, which is formed about the inner radial wall such that a flow annulus is formed therebetween;
a compressor discharge cavity formed about the combustor for receiving a combustor air supply delivered thereto by a compressor;
a staged injection system that includes the forward injector and, axially spaced aftward therefrom, a staged injector, wherein the staged injector intersects the flow annulus so to attain an injection point within the working fluid flowpath, and wherein, relative an axial position of the injection point, an aftward annulus section is defined to an aftward side of the injection point and a forward annulus section is defined to a forward side of the injection point;
fuel directing structure for apportioning a combustor fuel supply between the forward injector and the staged injector; and
air directing structure for apportioning the combustor air supply between the forward injector and the staged injector;
wherein the air directing structure comprises axially defined intake sections formed through the outer radial wall that fluidly connect the compressor discharge cavity to corresponding axially defined sections of the flow annulus, the intake sections including an aftward intake section that corresponds to the aftward annulus section and a forward intake section that corresponds to the forward annulus section; and
wherein the air directing structure is configured to:
direct air entering through the aftward intake section through the aftward annulus section in a forward direction to the staged injector; and
direct air entering through the forward intake section through the forward annulus section in an aftward direction to the staged injector;

wherein the flow annulus includes an axial partition that fluidly seals the forward annulus section from the aftward annulus section; and
wherein the axial partition comprises a wall that extends between the outer radial wall and the inner radial wall and about a circumference of the flow annulus so to seal the flow annulus against fluid communication between the aftward annulus section and the forward annulus section.

2. The gas turbine according to claim 1 , wherein the combustor comprises reference planes including: a forward reference plane, a mid reference plane, an aftward reference plane, and a staged injector reference plane, each of which comprising reference planes aligned substantially perpendicular to the longitudinal axis of the working fluid flowpath, wherein:
the forward reference plane aligns with the forward end of the working fluid flowpath;
the aftward reference plane aligns with the interface at which the combustor transitions to the turbine;
the mid reference plane aligns with an axial midpoint between of the working fluid flowpath between the forward and aftward reference plane; and
the staged injector reference plane aligns with the injection point of the staged injector; and

wherein:
the aftward intake section comprises an axial range defined approximately between the staged injector reference plane and the aftward reference plane; and
the forward intake section comprises an axial range defined approximately between the staged injector reference plane and the forward reference plane.

3. The gas turbine according to claim 2 , wherein:
the aftward intake section comprises a plurality of spaced impingement ports, each formed through the outer radial wall for training an impinged air jet against an outer surface of the inner radial wall; and
the forward intake section comprises a plurality of spaced impingement ports, each formed through the outer radial wall for training an impinged air jet against the outer surface of the inner radial wall.

4. The gas turbine according to claim 3 , wherein:
the plurality of impingement ports of the aftward intake section are axially spaced between an aftward most impingement port positioned just forward of the aftward reference plane and a forward most impingement port positioned just aftward of the staged injector reference plane;
the plurality of impingement ports of the forward intake section are axially spaced between an aftward most impingement port positioned just forward of the staged injector reference plane and a forward most impingement port positioned just aftward of the forward reference plane; and
wherein the plurality of impingement ports of each of the aftward and the forward intake sections are spaced circumferentially about substantially all of a circumference of the outer radial wall.

5. The gas turbine according to claim 4 , wherein the staged injector is positioned between the forward reference plane and the mid reference plane.

6. The gas turbine according to claim 4 , wherein the staged injector is positioned approximately at the mid reference plane.

7. The gas turbine according to claim 4 , wherein the staged injector is positioned at the aftward reference plane.

8. The gas turbine according to claim 4 , wherein the staged injector is positioned approximately midway between the mid reference plane and the aftward reference plane.

9. The gas turbine according to claim 4 , further comprising a plurality of the staged injectors that are spaced circumferentially about the staged injector reference plane; and
wherein each of the plurality of the staged injectors has corresponding ones of the aftward and the forward intake sections that angularly align therewith.

10. The gas turbine according to claim 4 , wherein the inner radial wall comprises axially stacked chambers defined therewithin, the axially stacked chambers including a forward chamber that houses the forward injector and an aftward chamber that defines a combustion zone; and
wherein a nozzle of the staged injector intersects the flow annulus so to attain the injection point; and
wherein the nozzle includes a tube extending between the outer radial wall and the inner radial wall, the tube blocking a portion of the flow annulus.

11. The gas turbine according to claim 10 , wherein the staged injector comprises an air port formed through the tube of the nozzle, the air port fluidly connecting the flow annulus to an interior of the tube; and
wherein the fuel directing structure includes:
a fuel passageway extending axially from a fuel source positioned near the headend of the combustor; and
fuel ports formed through the tube of the nozzle that fluidly connect the fuel passageway to the interior of the tube.

12. The gas turbine according to claim 11 , wherein the air port comprises a plurality of air ports, including: a first air port disposed on an aftward face of the tube configured for collecting an airflow from the aftward annulus section, and a second air port disposed on an forward face of the tube configured for collecting an airflow from the forward annulus section;
wherein the fuel ports are circumferentially spaced about an inner circumference of the tube of the nozzle; and
wherein the fuel passageway extends axially through an interior of the outer radial wall.

13. The gas turbine according to claim 11 , wherein the combustor air supply comprises a total supply of air delivered to the compressor di

Assignees

• Gen Electric

Inventors

• Hughes Michael John
• Berry Jonathan Dwight

Classifications

• F23R3/346

  for staged combustion · CPC title

• F23R3/002

  Wall structures (F23R3/02 and F23R3/007 take precedence) · CPC title

• F02C7/222 Primary

  Fuel flow conduits, e.g. manifolds · CPC title

• F02C3/04

  having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16) · CPC title

• F23R3/06 Primary

  Arrangement of apertures along the flame tube · CPC title