Systems and apparatus relating to downstream fuel and air injection in gas turbines

Patent metadata

Field	Value
Publication number	US-9714768-B2
Application number	US-201313836503-A
Country	US
Kind code	B2
Filing date	Mar 15, 2013
Priority date	Mar 15, 2013
Publication date	Jul 25, 2017
Grant date	Jul 25, 2017

Abstract

A gas turbine engine that includes: a combustor coupled to a turbine and a downstream injection system that includes two injection stages, a first stage and a second stage, positioned within an interior flowpath, wherein the first stage comprises an axial position that is aft of the primary air and fuel injection system and the second stage comprising an axial position that is aft of the first stage. Each of the first stage and the second stage include a plurality of circumferentially spaced injectors, each injector of which is configured to inject air and fuel into a flow through the interior flowpath. The first stage and the second stage have a configuration that limits a fuel injected at the second stage to less than 50% of a fuel injected at the first stage.

First claim

We claim:

1. A gas turbine engine that includes:
a combustor coupled to a turbine that together define an interior flowpath, the interior flowpath extending rearward about a longitudinal axis from a primary air and fuel injection system positioned at a forward end of the combustor, through an interface at which the combustor connects to the turbine, and through at least a row of stator blades in the turbine; and
a staged injection system that includes two injection stages, a first stage and a second stage, positioned within the interior flowpath, wherein the first stage and the second stage are each axially spaced along the longitudinal axis such that the first stage comprises an axial position that is aft of the primary air and fuel injection system and the second stage comprising an axial position that is aft of the first stage;
wherein each of the first stage and the second stage include a plurality of circumferentially spaced injectors, each injector of which is configured to inject air and fuel into a flow through the interior flowpath; and
wherein the first stage and the second stage comprise a configuration that limits a fuel injected at the second stage to less than 50% of a fuel injected at the first stage;
wherein the first stage is positioned aft of a longitudinal midpoint of the interior flowpath within the combustor;
wherein immediately aft of the primary air and fuel injection system, the interior flowpath includes a primary combustion zone defined by a surrounding liner and, immediately aft of the liner, the interior flowpath includes a transition zone defined by a surrounding transition piece;
wherein the transition piece is configured to fluidly couple the primary combustion zone to an inlet of the turbine while transitioning a flow through the transition piece from an approximate cylindrical cross-sectional area of the liner to an annular cross-sectional area of the inlet of the turbine;
wherein the transition piece comprises an aft frame that forms the interface between the combustor and the inlet of the turbine;
wherein the first stage of the staged injection system is positioned within the transition zone and the second stage of the staged injection system is positioned within or aft of the aft frame;
wherein the staged injection system comprises a third stage positioned within the interior flowpath, the third stage being configured to inject both air and fuel into the interior flowpath, wherein the second stage and the third stage are each axially spaced from each other along the longitudinal axis with the third stage comprising an axial position that is aft of the second stage; and
wherein the third stage includes a plurality of circumferentially spaced injectors, each of which is configured to inject both air and fuel into a combustion flow in the interior flowpath during operation.

2. The gas turbine of claim 1 , wherein each of the plurality of injectors at each of the first stage and the second stage are positioned on a common injection plane, each common injection plane aligned approximately perpendicular to the longitudinal axis of the interior flowpath;
wherein each of the first stage and the second stage comprise between 3 and 10 injectors; and
wherein the injectors of the first stage are circumferentially staggered with respect to the injectors of the second stage.

3. The gas turbine of claim 2 , wherein the injectors of the first stage and the second stage each comprise a configuration that, in operation, injects air and fuel in a direction between +30° and −30° to a reference line that is perpendicular relative a predominant direction of the flow through the interior flowpath; and
wherein the first stage comprises between 3 and 6 injectors and the second stage comprises between 5 and 10 injectors.

4. The gas turbine of claim 2 , wherein the injectors of the first stage and the second stage are configured so that the injected air and fuel from the first stage penetrate the combustion flow through the interior flowpath more than the injected air and fuel from the second stage; and
wherein the second stage comprise more injectors positioned about the interior flowpath than the first stage.

5. The gas turbine of claim 4 , wherein the injectors of the first stage are configured toward mixing the injected air and fuel with a combustion flow in a center region of the interior flowpath; and
wherein the injectors of the second stage are configured toward mixing the injected air and fuel with a combustion flow in a periphery region of the interior flowpath.

6. The gas turbine of claim 2 , wherein the circumferential placement of the injectors of the first stage comprises one from which the injected air and fuel therefrom penetrates predetermined areas of the interior flowpath based on an expected flow from the primary air and fuel injection system so to increase reactant mixing and temperature uniformity in a combustion flow downstream of the first stage; and
wherein the circumferential placement of the injectors of the second stage comprises one that compliments the circumferential placement of injectors of the first stage given a characteristic of an expected flow downstream of the first stage.

7. The gas turbine of claim 1 , wherein the first stage and the second stage comprise a configuration that limits a fuel injected at the second stage to between 10% and 50% of a fuel injected at the first stage.

8. The gas turbine of claim 1 , wherein the primary air and fuel injection system and the first stage and the second stage of the staged injection system are configured such that the following percentages of a total fuel supply are delivered to each during operation: between 50% and 80% delivered to the primary air and fuel injection system; between 20% and 40% delivered to the first stage; and between 2% and 10% delivered to the second stage;
wherein the primary air and fuel injection system and the first stage and the second stage of the staged injection system are configured such that the following percentages of a total combustor air supply are delivered to each during operation: between 60% and 85% delivered to the primary air and fuel injection system; between 15% and 35% delivered to the first stage; and between 1% and 5% delivered to the second stage.

9. The gas turbine of claim 8 , wherein the primary air and fuel injection system and the first stage and the second stage of the staged injection system are configured such that the following percentages of a total fuel supply are delivered to each during operation: about 70% delivered to the primary air and fuel injection system; about 25% delivered to the first stage; and about 5% delivered to the second stage;
wherein the primary air and fuel injection system and the first stage and the second stage of the staged injection system are configured such that the following percentages of a total air supply are delivered to each during operation: about 80% delivered to the primary air and fuel injection system; about 18% delivered to the first stage; and about 2% delivered to the second stage.

10. The gas turbine of claim 8 , wherein the first stage and the second stage comprise a common valve for controlling a fuel supply, wherein the first stage and the second stage are configured so that a desired fuel split therebetween is achieved through relative orifice sizing; and
wherein the first stage and the second stage comprise a common valve for controlling an air supply to each, wherein the first stage and the second stage are configured so that a desired air split is achieved through relative orifice sizing.

11. The gas turbine of claim 8 , further comprising a dual manifold through which a fuel supply i

Assignees

• Gen Electric

Inventors

• Davis Jr Lewis Berkley
• Venkataraman Krishna Kumar
• Graham Kaitlin Marie

Classifications

• F23R3/286

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

• F23R3/346

  for staged combustion · CPC title

• F23R3/20

  incorporating fuel injection means · CPC title

• F23R3/28 Primary

  characterised by the fuel supply (burners F23D) · CPC title