Apparatus and Method for Controlling the Secondary Injection of Fuel

• US 20090084082A1
• Filed: 08/20/2008
• Published: 04/02/2009
• Est. Priority Date: 09/14/2007
• Status: Active Grant

• Current Assignee: Siemens Energy Incorporated (Siemens AG)
• Original Assignee: Siemens Power Generation, Inc.
• Inventors: Cai, Weidong, Martin, Scott M., Harris, Arthur J. Jr.
• Granted Patent: US 8,387,398 B2
• Time in Patent Office: Days
• US Class Current: 60/39.281
• CPC Class Codes: F01D 9/023 Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
• CPC Class Codes: F02C 3/14 characterised by the arrangement of the combustion chamber in the plant combustion chambers per se F23R; F02C3/205 takes precedence
• CPC Class Codes: F02C 7/222 Fuel flow conduits, e.g. manifolds
• CPC Class Codes: F23D 2900/14004 with radially extending gas distribution spokes
• CPC Class Codes: F23K 5/10 Mixing with other fluids
• CPC Class Codes: F23N 1/002 using electronic means F23N1/04 - F23N1/10 take precedence
• CPC Class Codes: F23N 2225/14 Ambient temperature around burners
• CPC Class Codes: F23N 2237/02 two or more burners
• CPC Class Codes: F23N 2241/20 Gas turbines
• CPC Class Codes: F23R 3/346 for staged combustion

Abstract

In one embodiment, a combustor (28) for a gas turbine engine is provided comprising a primary combustion chamber (30) for combusting a first fuel to form a combustion flow stream (50) and a transition piece (32) located downstream from the primary combustion chamber (30). The transition piece (32) comprises a plurality of injectors (66) located around a circumference of the transition piece (32) for injecting a second fuel into the combustion flow stream (50). The injectors (66) are effective to create a radial temperature profile (74) at an exit (58) of the transition piece (32) having a reduced coefficient of variation relative to a radial temperature profile (64) at an inlet (54) of the transition piece (32). Methods for controlling the temperature profile of a secondary injection are also provided.

Claims

1. A method for controlling combustion in a gas turbine engine, the method comprising:

providing a primary combustion chamber and a transition piece located downstream from the primary combustion chamber;

injecting a first fuel into a compressed air stream flowing through the primary combustion chamber, wherein combustion of the first fuel forms a first radial temperature profile across the compressed air stream at an inlet of the transition piece;

injecting a second fuel preferentially into a relatively cooler portion of the compressed air stream within the transition piece, wherein combustion of the second fuel preferentially heats the relatively cooler portion of the air stream and is effective to provide a second radial temperature profile at an exit of the transition piece having a reduced coefficient of variation relative the first radial temperature profile at the inlet of the transition piece.

2. The method of claim 1, wherein the transition piece comprises at least a central zone, a first zone located peripherally outward from the central zone, and a second zone located peripherally outward from the first zone, and wherein the injecting a second fuel comprises:

injecting a first amount of the second fuel into the first zone effective to increase an average temperature of a portion of the compressed air stream in the first zone by a first amount; and

injecting a second amount of the second fuel into the second zone effective to increase an average temperature of a portion of the compressed air stream in the second zone by a second amount different than the first amount.

3. The method of claim 2, wherein the injecting of second fuel into the first and second zones is done by radially injecting the second fuel from a plurality of first injectors and a plurality of second injectors located around a perimeter of the transition piece, wherein the plurality of first injectors are effective to inject the first amount of second fuel into the first zone, and wherein the plurality of second injectors are effective to inject the second amount of second fuel into the second zone.

4. The method of claim 3, wherein the plurality of first injectors are effective to inject a radial stream of the first amount of second fuel into the first zone, and wherein the plurality of second injectors are effective to inject a homogenized dispersion of the second amount of second fuel into the second zone.

5. The method of claim 4, wherein the plurality of first injectors comprise nozzles having a center hole through which the first amount of second fuel can be injected radially and the plurality of second injectors comprise nozzles having a center hole through which a first portion of the second amount of second fuel can be injected radially and a plurality of injection holes angled relative to the center hole through which additional portions of the second amount of second fuel can be injected angularly.

6. The method of claim 5, wherein the nozzles of the plurality of first injectors comprise nozzles of a first length and the nozzles of the plurality of second injectors comprise nozzles of a second length shorter than the first length.

7. The method of claim 1, further comprising, prior to the injecting the second fuel, mixing the fuel with a diluent or combination of diluents selected from the group consisting of steam, nitrogen, carbon dioxide and an inert gas.

8. A method for controlling combustion in a gas turbine engine, the method comprising:

providing a primary combustion chamber and a transition piece located downstream from the primary combustion chamber;

injecting a first fuel into an air stream flowing through the primary combustion chamber to create a combustion stream having a first momentum in an axial direction through the primary combustion chamber and the transition piece and having a radial temperature differential between an average temperature in a hotter central region and an average temperature in a cooler peripheral region of the combustion stream flowing at an inlet to the transition piece; and

injecting a second fuel into the combustion stream in a radial direction perpendicular to the axial direction within the transition piece at a momentum ratio of 25-50, wherein combustion of the second fuel is effective to reduce the radial temperature differential of the combustion stream as it flows through the transition piece.

9. The method of claim 8, wherein the cooler peripheral region comprises a first zone located circumferentially peripherally outward from the central region and a second zone located peripherally outward from the first zone, and wherein the injecting a second fuel comprises.

injecting a portion of the second fuel in the first annular zone at a momentum ratio of between 30 and 50 effective to increase an average temperature in the first zone as the combustion stream flows through the transition piece;

and

injecting a portion of the second fuel in the second annular zone of between 25 and 45 momentum ratio effective to increase an average temperature in the second zone as the combustion stream flows through the transition piece.

10. The method of claim 8, wherein the second fuel consists essentially of fuel, and wherein the injecting is done at a momentum ratio of at least about 50.

11. A combustor for a gas turbine engine comprising:

a primary combustion chamber for combusting a first fuel to form a combustion flow stream; and

a transition piece located downstream from the primary combustion chamber, the transition piece comprising a plurality of injectors located around a circumference of the transition piece for injecting a second fuel into the combustion flow stream;

wherein the injectors are effective to create a radial temperature profile at an exit of the transition piece having a reduced coefficient of variation relative to a radial temperature profile at an inlet of the transition piece.

12. The combustor of claim 11, wherein the plurality of injectors are located at an axial position along the transition piece and are effective to ensure completion of burning of the second fuel within the combustion flow stream within a target distance of a downstream end of the transition piece.

13. The combustor of claim 12, wherein the target distance is from 0 to 20 percent of an axial length of the transition piece.

14. The combustor of claim 13, wherein the target distance is less than 10 percent of the axial length of the transition piece.

15. A transition piece for a gas turbine engine comprising a plurality of fuel injectors located between an inlet end of the transition piece and a point from 0 to 75 percent along an axial length of the transition piece from the inlet end, wherein the injectors are effective to reduce a radial temperature differential between an average temperature in a hotter central region of the combustion stream and an average temperature in a cooler peripheral region of the combustion stream as it flows through the transition piece.

16. The transition piece of claim 15, wherein the injectors are located at a point from 25 to 75 percent along the axial length of the transition piece.

17. A combustor for a gas turbine combustion system comprising:

a primary combustion chamber adapted for combusting a first fuel to produce a combustion stream;

a transition piece for receiving the combustion stream from the primary combustion chamber and for combusting a second fuel, the transition piece comprising a central region, a first zone located peripherally outward from the central region, a second zone located peripherally outward from the first zone, and a plurality of injectors located around a perimeter of the transition piece for injecting the second fuel;

wherein the plurality of injectors comprise a plurality of first injectors and a plurality of second injectors; and

wherein the first injectors are effective to inject a radial stream of the second fuel into the first zone and the second injectors are effective to inject a homogenized dispersion of the second fuel into the second zone.