Gas turbine combustion system and method of flame stabilization in such a system

Abstract

A gas turbine combustion system is provided. In an embodiment, the system includes a first radial inflow swirler having first radial outer intake openings, first radial inner outlet openings and first flow passages, each first flow passage including a first angle (a) with respect to the radial direction, a second radial inflow swirler having second radial outer intake openings, each second flow passage including a second angle (p) with respect to the radial direction, where the radial outer circumference of the second radial inflow swirler has a diameter that is smaller than the diameter of the radial inner circumference of the first radial inflow swirler and the second radial inflow swirler is located coaxially with and radially inside the first radial inflow swirler. The first angle (a) has a different sign than the second angle (p) with respect to the radial direction.

1 . A gas turbine combustion system, comprising:
 a central axis (A) and a radial direction with respect to said central axis (A);   a first radial inflow swirler comprising first radial outer intake openings located at a radial outer circumference of the first radial inflow swirler, the first radial inner outlet openings located at a radial inner circumference of the first radial inflow swirler, and first flow passages extending from the first radial outer intake openings to the first radial inner outlet openings, each first flow passage including a first angle (a) with respect to the radial direction;   a second radial inflow swirler comprising second radial outer intake openings located at a radial outer circumference of the second radial inflow swirler, the second radial inner outlet openings located at a radial inner circumference of the second radial inflow swirler, and second flow passages extending from the second radial outer intake openings to the second radial inner outlet openings, each second flow passage including a second angle (β) with respect to the radial direction;   wherein the radial outer circumference of the second radial inflow swirler has comprises a diameter that is smaller than the diameter of the radial inner circumference of the first radial inflow swirler and wherein the second radial inflow swirler is located coaxially with and radially inside an opening formed by the inner circumference of the first radial inflow swirler so that a fluid that exits the outlet openings of the first radial inflow swirler is directed towards the second radial inflow swirler, and   wherein the first angle (a) has a different sign than the second angle (β) with respect to the radial direction.   
     
     
         2 . The gas turbine combustion system as claimed in  claim 1 , wherein fuel injection openings are located in the second radial inflow swirler and are open towards the second flow passages. 
     
     
         3 . The gas turbine combustion system as claimed in  claim 2 , wherein the fuel injection openings are located inside the second flow passages. 
     
     
         4 . The gas turbine combustion system as claimed in  claim 3 , wherein the fuel injection openings are located in the radial outer half of the second flow passages. 
     
     
         5 . The gas turbine combustion system as claimed in  claim 1 , wherein the number of second flow passages is identical to the number of first flow passages. 
     
     
         6 . The gas turbine combustion system as claimed in  claim 1 , wherein a radial gap is present between the radial inner circumference of the first radial inflow swirler and the radial outer circumference of the second radial inflow swirler. 
     
     
         7 . The gas turbine combustion system as claimed in  claim 6 , wherein the flow cross section of the second flow passages is smaller than the flow cross section of the first flow passages. 
     
     
         8 . The gas turbine combustion system as claimed in  claim 1 , wherein the first angle (a) and the second angle (β) have the same absolute value. 
     
     
         9 . A method of flame stabilisation in a gas turbine combustion system in which a fluid flows along a flow path with a radial component, by use of a gas turbine combustion system as claimed in  claim 1 , wherein,
 the fluid is a fluid that comprises an oxidant and a fuel is mixed with the fluid that comprises an oxidant so as to transform the fluid into a mixture comprising fuel and the oxidant;   a first swirl with a first rotational direction is generated in the flowing fluid in a radial upstream section of the flow path by passing the fluid through the first radial inflow swirler of the gas turbine combustion system to generate a swirling fluid;   in a radial downstream section of the flow path a second swirl with a second rotational direction is generated in at least a portion of the fluid by passing said portion of the swirling fluid through the second radial inflow swirler of the gas turbine combustion system, wherein the second rotational direction represents a counter rotation with respect to the first rotational direction.   
     
     
         10 . The method as claimed in  claim 9 , wherein fuel is introduced into the fluid where the second swirl is generated. 
     
     
         11 . The method as claimed in  claim 10 , wherein the fuel is introduced into the fluid at a location where generation of the second swirl begins. 
     
     
         12 . The method as claimed in  claim 9 , wherein no second swirl is introduced into a portion of the fluid.