IGNITER FOR A GAS TURBINE ENGINE

• US 20170130653A1
• Filed: 11/09/2015
• Published: 05/11/2017
• Est. Priority Date: 11/09/2015
• Status: Active Grant

• Current Assignee: General Electric Company
• Original Assignee: General Electric Company
• Inventors: Klasing, Kevin Samuel, Gonyou, Craig Alan, Bennett, William Thomas
• Granted Patent: US 10,738,707 B2
• Time in Patent Office: Days
• CPC Class Codes: F02C 7/264 Ignition
• CPC Class Codes: F05D 2260/99 Ignition, e.g. ignition by warming up of fuel or oxidizer in a resonant acoustic cavity
• CPC Class Codes: F23Q 3/008 Structurally associated with fluid-fuel burners
• CPC Class Codes: F23R 3/343 Pilot flames, i.e. fuel nozzles or injectors using only a very small proportion of the total fuel to insure continuous combustion ignition in gas-turbine plants F02C7/264; pilot flame igniters F23Q9/00

Abstract

A gas turbine engine includes a combustion section space between a compressor section in a turbine section. The combustion section defines a combustion chamber and includes a combustor member defining an opening to the combustion chamber. A mounting assembly extends around or is positioned adjacent to the opening defined by the combustor member. In igniter extends through a ferrule of the mounting assembly and includes a distal end positioned proximate the opening in the combustor member. The igniter defines a plurality of channels each channel extending between a first end and a second end. The first end is positioned away from the distal end of the igniter relative to the second end, and the second end is a terminal end spaced from the distal end of the igniter.

Claims

1. A gas turbine engine comprising:

a compressor section;

a turbine section located downstream of the compressor section; and

a combustion section disposed between the compressor section and the turbine section, the combustion section defining a combustion chamber and comprising

a combustor member defining an opening to the combustor chamber;

a mounting assembly extending around or positioned adjacent to the opening defined by the combustor member, the mounting assembly comprising a ferrule; and

an igniter extending through the ferrule and including a distal end positioned proximate the opening in the combustor member, the igniter defining a plurality of channels, each channel extending between a first end and a second end, the first end positioned away from the distal end of the igniter relative to the second end, and the second end being a terminal end spaced from the distal end of the igniter.

2. The gas turbine engine of claim 1, wherein the igniter further comprises an agitation portion between the second ends of each of the channels and the distal end of the igniter for disrupting an airflow through the channels.

3. The gas turbine engine of claim 2, wherein the agitation portion extends around a circumference of the igniter between the second ends of each of the plurality of channels and the distal end of the igniter.

4. The gas turbine engine of claim 2, wherein the agitation portion defines a dimension along an axial direction of the igniter of at least about 0.02 inches.

5. The gas turbine engine of claim 1, wherein the combustor member is a combustor liner.

6. The gas turbine engine of claim 1, wherein the ferrule includes a contact portion that contacts the igniter, and wherein the first ends of the plurality of channels are positioned outward from the contact portion of the ferrule relative to the combustor member.

7. The gas turbine engine of claim 1, wherein the igniter defines an annular opening with the combustor member around the distal end of the igniter.

8. The gas turbine engine of claim 1, wherein each of the plurality of channels extend generally along an axial direction of the igniter.

9. The gas turbine engine of claim 1, wherein the plurality of channels are evenly spaced along a circumferential direction of the igniter.

10. The gas turbine engine of claim 1, wherein each of the channels define a width between about 0.02 inches and about 0.2 inches.

11. The gas turbine engine of claim 1, wherein the combustion assembly additionally includes a combustor casing, and wherein the igniter is attached to the combustor casing and is slidably received within the ferrule.

12. The gas turbine engine of claim 1, wherein the igniter defines an outside surface, and wherein the plurality of channels extend along the outside surface of the igniter.

13. The gas turbine engine of claim 1, wherein the mounting assembly further includes a cap extending around the opening in the combustor member, and wherein the ferrule is moveably attached to the cap.

14. An igniter for a combustion section of a gas turbine engine, the combustion section including a combustor liner defining an opening and a mounting assembly extending around or adjacent to the opening defined by the combustor liner, the mounting assembly including a ferrule, the igniter comprising:

a tip including a distal end of the igniter, the tip configured to extend through the ferrule and at least partially into the opening in the combustor liner when the igniter is installed in the combustion assembly, the tip defining an exterior surface, and the exterior surface defining a plurality of channels, each channel extending between a first end and a second end, the first end positioned away from the distal end of the igniter relative to the second end, and the second end spaced from the distal end of the igniter.

15. The igniter of claim 14, further comprising

an agitation portion between the second ends of each of the channels and the distal end for disrupting a cooling airflow through the channel.

16. The igniter of claim 15, wherein the agitation portion extends around a circumference of the igniter between the second ends of each of the plurality of channels and the distal end of the igniter.

17. The igniter of claim 15, wherein the agitation portion defines a dimension along an axial direction of the igniter of at least about 0.02 inches.

18. The igniter of claim 14, wherein each of the plurality of channels extend generally along an axial direction of the igniter.

19. The igniter of claim 14, wherein the plurality of channels are evenly spaced along a circumferential direction of the igniter.

20. The igniter of claim 14, wherein each of the channels define a width between about 0.02 inches and about 0.2 inches.