Combustor and method

• US 20060042263A1
• Filed: 08/27/2004
• Published: 03/02/2006
• Est. Priority Date: 08/27/2004
• Status: Active Grant

• Current Assignee: Pratt & Whitney Canada Corporation (Rtx Corporation)
• Original Assignee: Pratt & Whitney Canada Corporation (Rtx Corporation)
• Inventors: Sampath, Parthasarathy, Morenko, Oleg, Patel, Bhawan Bhai
• Granted Patent: US 7,260,936 B2
• Time in Patent Office: Days
• US Class Current: 60/776
• CPC Class Codes: F23R 3/06 Arrangement of apertures along the flame tube
• CPC Class Codes: F23R 3/12 inducing a vortex
• CPC Class Codes: F23R 3/54 Reverse-flow combustion chambers
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A gas turbine engine combustor includes an apparatus for directing air around the axis along which fuel is injected into the combustor to thereby improve the combustor.

Claims

1. A gas turbine engine combustor comprising a liner enclosing a combustion chamber, the liner including a dome portion at an upstream end thereof, the dome portion having defined therein a plurality of openings each adapted to receive a fuel nozzle and a plurality of holes defined around each opening, each opening having an axis generally aligned with an fuel injection axis of a fuel nozzle received by the opening, the holes adapted to direct air into the combustion chamber in a spiral around the axis of an associated one of said openings.

2. The combustor of claim 1 wherein the holes are defined substantially circumferentially around the openings.

3. The combustor of claim 1 wherein the holes are defined concentrically around the axis of its associated opening.

4. The combustor of claim 1 wherein the holes are defined in a plurality of rows around at least one opening.

5. The combustor of claim 4 wherein the rows are concentric with one another.

6. The combustor of claim 1 wherein the combustor includes a region wherein at least some holes associated with one opening are interlaced with at least some holes associated with another opening.

7. The combustor of claim 1 wherein the combustor includes a region wherein at least some holes associated with one opening are interlaced with a second set of holes, said second set adapted to direct a non-spiralling flow of air into the combustor.

8. The combustor of claim 1 wherein the holes are angled to admit air into the combustor generally tangentially relative to the opening.

9. The combustor of claim 1 wherein the holes are adapted to direct air into a vortex of sufficient strength to, in use, constrain a lateral extent of fuel entering the combustor via said fuel nozzles.

10. The combustor of claim 1 wherein the openings and holes are provided in a portion of the dome portion which is substantially perpendicular to the axis.

11. The combustor of claim 1 wherein the openings and holes are provided in a generally planar portion of the dome portion.

12. A gas turbine engine combustor comprising a liner enclosing a combustion chamber, the liner having defined therein a plurality of openings each adapted to receive a fuel nozzle for directing fuel into the combustion chamber generally along an axis of the opening, the liner also having means associated with each opening for directing air into the combustion chamber in a spiral pattern around an axis of the associated opening.

13. The combustor of claim 12 wherein the means for directing air comprises means for directing said air into the combustion chamber generally tangentially relative the associated opening.

14. The combustor of claim 12 wherein the means for directing air is disposed substantially around each of said openings.

15. The combustor of claim 12 wherein the means for directing air is disposed concentrically with each of said openings.

16. The combustor of claim 12 wherein the means for directing air is disposed substantially perpendicularly to the axis.

17. The combustor of claim 12 wherein the means for directing air is provided in a generally planar portion of the liner.

18. A method of combusting fuel in a gas turbine combustor, the method comprising the steps of:

injecting a mixture of fuel and air into the combustor along an axis;

igniting the mixture to create at least one combustion zone in which the mixture is combusted; and

directing air into the combustor around said axis in a spiral pattern.

19. The method of claim 18 wherein the air is directed in a spiral pattern around the combustion zone.

20. The method of claim 18 further comprising the step of using the spiralling air to control the lateral width of the combustion zone relative to the axis.