Two-stage combustor for gas turbine engine

Abstract

A combustor for a gas turbine engine comprises an inner annular liner wall and an outer annular liner wall cooperating to form a combustion chamber of the combustor. A first dome wall has a circumferential array of first fuel injection bores. A second dome has a circumferential array of second fuel injection bores. An intermediate wall extends between the first dome wall and the second dome wall. A first combustion stage is defined by the inner liner wall forward end, the first dome wall and the intermediate wall. A second combustion stage is defined at least by the outer liner wall forward end, the second dome wall and the intermediate wall, the first combustion stage communicating with the first fuel injection bores, the second combustion stage communicating with the second fuel injection bores.

What is claimed is: 
     
         1 . A combustor for a gas turbine engine, the engine having a central axis, the combustor comprising:
 an inner annular liner wall having an axially forward end generally radially oriented, the inner annular liner wall curving into an axial orientation as it extends in an aft direction to an axially aft end of the inner liner;   an outer annular liner wall having an axially forward end generally radially oriented, the outer annular liner wall curving into an axial orientation as it extends in an aft direction to an axially aft end of the outer liner, with the inner annular liner wall and the outer annular liner wall spaced apart and cooperating to form a combustion chamber of the combustor;   a first dome wall extending generally axially adjacent to the inner annular liner wall and circumscribing the forward end on the inner liner, the first dome wall having a circumferential array of first fuel injection bores radially extending through the first dome wall;   a second dome wall extending generally axially adjacent to the outer annular liner wall and circumscribing the forward end of the outer liner, the second dome wall having a circumferential array of second fuel injection bores radially extending through the second dome wall; and   an intermediate wall extending between the first dome wall and the second dome wall and being an intermediate single annular wall, a first combustion stage defined at least by the inner liner wall forward end, the first dome wall and the intermediate wall, a second combustion stage defined at least by the outer liner wall forward end, the second dome wall and the intermediate wall, the first combustion stage communicating with the first fuel injection bores, the second combustion stage communicating with the second fuel injection bores.   
     
     
         2 . The combustor according to  claim 1 , wherein the intermediate wall is radially oriented and wherein the second dome wall is spaced radially outward from the first dome wall. 
     
     
         3 . The combustor according to  claim 1 , further comprising mixing features projecting into the combustion chamber from the dome and further separating the stages of combustion. 
     
     
         4 . The combustor according to  claim 3 , wherein the mixing features extend from the intermediate dome wall. 
     
     
         5 . The combustor according to  claim 1 , wherein the second dome wall has a radius of curvature larger than a radius of the first dome wall. 
     
     
         6 . The combustor according to  claim 1 , wherein the first combustion stage is a main combustion stage and wherein the second combustion stage is a pilot combustion stage. 
     
     
         7 . The combustor according to  claim 6 , wherein the main combustion stage has a volume larger than the pilot combustion stage. 
     
     
         8 . The combustor according to  claim 1 , wherein the first dome wall is axially wider than the second dome wall. 
     
     
         9 . The combustor according to  claim 1 , wherein the circumferential array of first fuel injection bores is circumferentially offset from the circumferential array of second fuel injection bores. 
     
     
         10 . The combustor according to  claim 1 , wherein the first combustion stage is entirely axially forward of the second combustion stage. 
     
     
         11 . The combustor according to  claim 1 , wherein the first combustion stage is generally radially inwardly oriented. 
     
     
         12 . The combustor according to  claim 11 , wherein the second combustion stage is generally radially inwardly oriented. 
     
     
         13 . The combustor according to  claim 1 , wherein the first dome wall is in a non-parallel orientation relative to a centerline of the combustor. 
     
     
         14 . The combustor according to  claim 1 , wherein the second dome wall is a generally parallel orientation relative to a centerline of the combustor. 
     
     
         15 . The combustor according to  claim 1 , wherein the intermediate dome wall is in a normal orientation relative to a centerline of the combustor. 
     
     
         16 . The combustor according to  claim 1 , further comprising dilution ports in the liner walls downstream of the stages. 
     
     
         17 . A gas turbine engine comprising:
 a casing defining a plenum;   a combustor within the plenum comprising:
 an inner annular liner wall having an axially forward end generally radially oriented, the inner annular liner wall curving into an axial orientation as it extends in an aft direction to an axially aft end of the inner liner; 
 an outer annular liner wall having an axially forward end generally radially oriented, the outer annular liner wall curving into an axial orientation as it extends in an aft direction to an axially aft end of the outer liner, with the inner annular liner wall and the outer annular liner wall spaced apart and cooperating to form a combustion chamber of the combustor; 
 a first dome wall extending generally axially adjacent to the inner annular liner wall and circumscribing the forward end on the inner liner, the first dome wall having a circumferential array of first fuel injection bores radially extending through the first dome wall; 
 a second dome wall extending generally axially adjacent to the outer annular liner wall and circumscribing the forward end of the outer liner, the second dome wall having a circumferential array of second fuel injection bores radially extending through the second dome wall; and 
 an intermediate wall extending generally radially between the first dome wall and the second dome wall and being an intermediate single annular wall, a first combustion stage defined at least by the inner liner wall forward end, the first dome wall and the intermediate wall, a second combustion stage defined at least by the outer liner wall forward end, the second dome wall and the intermediate wall, the first combustion stage communicating with the first fuel injection bores, the second combustion stage communicating with the second fuel injection bores; 
 a compressor diffuser having at least one outlet within the plenum; and 
 fuel injectors and/or valves disposed in communication the fuel injection bores. 
   
     
     
         18 . The gas turbine engine according to  claim 17 , wherein the diffuser comprises a plurality of passages circumferentially distributed about the combustor, with the outlets of the diffuser passages being radially offset from the first fuel injection bores of the first dome wall. 
     
     
         19 . The gas turbine engine according to  claim 17 , wherein the second fuel injection bore holes of the second dome wall are radially aligned with the diffuser outlets. 
     
     
         20 . The gas turbine engine according to  claim 1 , wherein the first combustion stage is a main combustion stage and wherein the second combustion stage is a pilot combustion stage.