Two-stage combustor for gas turbine engine
A combustor for a gas turbine engine comprises an inner annular liner and an outer annular liner. First and second combustion stages are defined between the liners. Each combustion stage has a plurality of fuel injection bores distributed in a liner wall defining the respective stage. A lobed mixer extends into the combustor, the lobed mixer arranged to receive combustion gases from each combustion stage for mixing flows of said combustion gases.
1. A combustor for a gas turbine engine comprising:
liner walls circumscribing combustion stages, the liner walls forming at least an inner annular liner, an outer annular liner and dome portions;
first and second combustion stages defined between the liners and dome portions and in a parallel arrangement relative to one another, each said combustion stage having a plurality of fuel injection bores formed into and circumferentially distributed in said dome portions defining the respective stage; and
a lobed mixer extending into the combustor and located between the first combustion stage and the second combustion stage, the lobed mixer arranged to receive combustion gases from each combustion stage for mixing flows of said combustion gases.
2. The combustor according to claim 1 , wherein the first and second stages extend generally radially inwardly, and wherein the lobed mixer extends generally radially inwardly intermediate the two stages.
3. The combustor according to claim 1 , wherein the lobed mixer is disposed entirely within the combustor and between the inner annular liner and the outer annular liner.
4. The combustor according to claim 1 , wherein the liner walls include an intermediate wall separating the first combustion stage from the second combustion stage, and wherein the lobed mixer extends from the intermediate wall into the combustor.
5. The combustor according to claim 1 , wherein valleys of the lobed mixer wall are in circumferential register with the injection bores of one of said combustion stages, while the peaks of the lobed mixer are in circumferential register with the injection bores of the other said combustion stage.
6. The combustor according to claim 1 , wherein the inner annular liner wall has an axially forward end generally radially oriented, the inner annular liner wall curving into an axial orientation in an aft direction.
7. The combustor according to claim 1 , wherein the outer annular liner wall has an axially forward end generally radially oriented, the outer annular liner wall curving into an axial orientation in an aft direction.
8. The combustor according to claim 1 , wherein the dome portions include a first dome wall and a second dome wall, the first combustion stage being defined by the inner annular liner, the first dome wall and the lobed mixer, the second combustion stage being defined by the outer annular liner, the second dome wall and the lobed mixer.
9. The combustor according to claim 8 , wherein edges of valleys of the lobed mixer wall are generally normal to a plane of their respective one of the first dome wall and second dome wall.
10. A gas turbine engine comprising:
a casing defining a plenum;
a combustor within the plenum and comprising:
liner walls circumscribing combustion stages, the liner walls forming at least an inner annular liner, an outer annular liner and dome portions;
first and second combustion stages defined between the liners and in a parallel arrangement relative to one another, each said combustion stage having a plurality of fuel injection bores formed into and circumferentially distributed in said dome portions defining the respective stage; and
a lobed mixer extending into the combustor and located between the first combustion stage and the second combustion stage, the lobed mixer arranged to receive combustion gases from each combustion stage for mixing flows of said combustion gases;
a diffuser having outlets communicating with the plenum; and
injectors and/or valves at the injection bores.
11. The gas turbine engine according to claim 10 , wherein the first and second stages extend generally radially inwardly, and wherein the lobed mixer extends generally radially inwardly intermediate the two stages.
12. The gas turbine engine according to claim 10 , wherein the lobed mixer is disposed entirely within the combustor and between the inner annular liner and the outer annular liner.
13. The gas turbine engine according to claim 10 , wherein the liner walls include an intermediate wall separating the first combustion stage from the second combustion stage, and wherein the lobed mixer extends from the intermediate wall into the combustor.
14. The gas turbine engine according to claim 10 , wherein valleys of the lobed mixer wall are in circumferential register with the injection bores of one of said combustion stages, while the peaks of the lobed mixer are in circumferential register with the injection bores of the other said combustion stage.
15. The gas turbine engine according to claim 10 , wherein the inner annular liner wall having an axially forward end generally radially oriented, the inner annular liner wall curving into an axial orientation in an aft direction.
16. The gas turbine engine according to claim 10 , wherein the dome portions include a first dome wall and a second dome wall, the first combustion stage being defined by the inner annular liner, the first dome wall and the lobed mixer, the second combustion stage being defined by the outer annular liner, the second dome wall and the lobed mixer.
17. The gas turbine engine according to claim 16 , wherein edges of valleys of the lobed mixer are generally normal to a plane of their respective one of the first dome wall and second dome wall.
18. The gas turbine engine according to claim 10 , wherein the diffuser outlets are circumferentially distributed about the combustor, with the outlets of the diffuser being offset from the injection bores of the first stage.