Two-stage combustor for gas turbine engine

Patent metadata

Field	Value
Publication number	US-9416972-B2
Application number	US-201113313305-A
Country	US
Kind code	B2
Filing date	Dec 7, 2011
Priority date	Dec 7, 2011
Publication date	Aug 16, 2016
Grant date	Aug 16, 2016

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.

First claim

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 an inner liner;
an outer annular liner wall having 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 an 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;
an intermediate wall extending between the first dome wall and the second dome wall and being an intermediate single wall wherein the second dome wall is spaced radially outward from the first dome wall and the first dome wall is connected to the intermediate wall at a location different than the second dome wall; and
a mixing feature defined by a single wall projecting into the combustion chamber from the intermediate wall, the single wall separating the first combustion stage from the second combustion stage;
wherein a first combustion stage is defined at least by the inner annular liner wall forward end, the first dome wall, and the intermediate single wall or mixing feature, a second combustion stage defined at least by the outer annular liner wall forward end, the second dome wall, and the intermediate single wall or mixing feature, 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.

3. 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.

4. 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.

5. The combustor according to claim 4 , wherein the main combustion stage has a volume larger than the pilot combustion stage.

6. The combustor according to claim 1 , wherein the first dome wall is axially wider than the second dome wall.

7. The combustor according to claim 1 , wherein the circumferential array of the first fuel injection bores is circumferentially offset from the circumferential array of the second fuel injection bores.

8. The combustor according to claim 1 , wherein the first combustion stage is generally radially inwardly oriented.

9. The combustor according to claim 8 , wherein the second combustion stage is radially inwardly oriented.

10. The combustor according to claim 1 , wherein the first dome wall is in a non-parallel orientation relative to the central axis.

11. The combustor according to claim 1 , wherein the second dome wall is a generally parallel orientation relative to the central axis.

12. The combustor according to claim 1 , wherein the intermediate wall is in a normal orientation relative to the central axis.

13. The combustor according to claim 1 , further comprising dilution ports in the inner annular liner wall and the outer annular liner wall downstream of the first combustion stage and of the second combustion stage.

14. 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.

15. The gas turbine engine according to claim 1 , wherein the single wall is a single integral annular piece.

16. 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 an 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 an outer liner, with the inner annular liner wall and the outer annular liner wad 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;
an intermediate wall extending generally radially between the first dome wall and the second dome wall and being an intermediate single wall wherein the second dome wall is spaced radially outward from the first dome wall and the first dome wall is connected to the intermediate wall at a location different than the second dome wall;
a mixing feature defined by a single wall projecting into the combustion chamber from the intermediate wall, the single wall separating the first combustion stage from the second combustion stage, wherein a first combustion stage is defined at least by the inner annular liner wall forward end, the first dome wall, and the intermediate single wall or mixing feature, a second combustion stage defined at least by the outer annular liner wall forward end, the second dome wall, and the intermediate single wall or mixing feature, 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.

17. The gas turbine engine according to claim 16 , 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.

18. The gas turbine engine according to claim 16 , wherein the second fuel injection bore holes of the second dome wall are circumferentially aligned with the diffuser outlets.

Assignees

• Hawie Eduardo
• Davenport Nigel
• Pratt & Whitney Canada

Inventors

• Hawie Eduardo
• Davenport Nigel

Classifications

• F23R3/002

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

• F05D2250/141

  circular · CPC title

• F23R3/26

  Controlling the air flow · CPC title

• F23R3/346 Primary

  for staged combustion · CPC title

• F05D2240/35

  Combustors or associated equipment · CPC title