Combustion system for a gas turbine engine and method of operating same

Patent metadata

Field	Value
Publication number	US-9683744-B2
Application number	US-201414193575-A
Country	US
Kind code	B2
Filing date	Feb 28, 2014
Priority date	Feb 28, 2014
Publication date	Jun 20, 2017
Grant date	Jun 20, 2017

Abstract

A gas turbine engine comprises a combustion system comprising a secondary annular combustor and a primary combustor in fluid communication with the secondary combustor, a secondary fuel injector associated with the secondary combustor, a primary fuel injector associated with the primary combustor, and a ECU controlling fuel delivery to the secondary and primary fuel injectors. The primary fuel injector delivers fuel to the primary combustor. The ECU allows fuel to be delivered to the secondary fuel injector in addition to the primary fuel injector only when a fuel amount higher is requested delivered by the primary fuel injector. A method of operating a gas turbine engine is also presented.

First claim

The invention claimed is:

1. A gas turbine engine comprising:
a combustion system having a primary combustor in fluid communication with a secondary combustor downstream thereof relative to a flow of fuel circulating therein, the primary combustor disposed radially outward of the secondary combustor;
a primary fuel injector assembly associated with the primary combustor;
a secondary fuel injector assembly associated with the secondary combustor;
a fuel conduit network fluidly connected to the primary and secondary fuel injector assemblies; and
an electronic control unit (ECU) configured in a first mode for delivering fuel from a source via the conduit network only to the primary fuel injector assembly and in a second mode from a source via the conduit network to both the primary and secondary fuel injector assemblies.

2. The gas turbine engine as defined in claim 1 , wherein the secondary fuel injector assembly comprises a plurality of fuel injection points to deliver a substantially uniform annular flow of fuel to the secondary combustor.

3. The gas turbine engine as defined in claim 1 , wherein the primary combustor is an annular combustor.

4. The gas turbine engine as defined in claim 1 , wherein the primary combustor converges downstream as it communicates with the secondary combustor.

5. The gas turbine engine as defined in claim 1 , wherein the primary combustor and the secondary combustor converge at an angle comprised between 20° and 30°.

6. The gas turbine engine as defined in claim 1 , wherein the primary and secondary combustors are arranged in series, the primary emptying into the secondary, and a combined combustion chamber therefore has a single outlet.

7. The gas turbine engine as defined in claim 1 , wherein the secondary combustor is arranged generally parallel to a central axis of the gas turbine engine, and the primary combustor is disposed along a primary combustor axis which intersects with the central axis of the gas turbine engine.

8. A method of operating a gas turbine engine, the engine having a primary combustor fed by a primary fuel injector assembly and a secondary combustor serially downstream of the primary combustor and fed by a secondary fuel injector assembly, the method comprising, in sequence:
a) in response to a low power command input which is below a selected power threshold level, delivering fuel only to the primary fuel injector assembly of the primary combustor; and
b) in response to a high power command input which is above said selected power threshold level, delivering fuel serially downstream of the primary combustor to the secondary fuel injector assembly of the secondary combustor while also delivering fuel to the primary fuel injector assembly of the primary combustor.

9. The method as defined in claim 8 , wherein more fuel is delivered to the secondary fuel injector assembly than the primary fuel injector assembly in step b).

10. The method as defined in claim 8 , where the fuel delivered to the secondary combustor in step b) is between 75% and 82% of total fuel flow provided to the primary and secondary combustors.

11. The method as defined in claim 8 , wherein fuel delivered to the secondary combustor is about 80% of a total fuel flow provided to the primary and secondary combustors.

12. The method as defined in claim 8 , wherein delivering fuel only to the primary fuel injector in response to the low power command input comprises delivering fuel only to the primary fuel injector in response to the low power command input requiring a fuel amount lower than a maximum fuel amount delivered by the primary fuel injector.

13. The method as defined in claim 8 , wherein a fuel flow amount delivered to the primary combustor in step b) is between 18% and 25% of a total fuel flow amount delivered to the primary and secondary combustion chambers.

14. The method as defined in claim 8 , wherein the fuel amount delivered to the primary combustor in step b) is 20% of the total fuel amount delivered to the primary and secondary combustion chambers.

15. The method as defined in claim 8 , wherein a fuel flow rate provided to the primary combustor is about 50% of a total fuel flow rate delivered to the primary and secondary combustion chambers in steps a) and b).

16. The method as defined in claim 8 , wherein step b) corresponds to at least one of a take-off and an altitude cruising flight condition.

17. The method as defined in claim 8 , wherein step a) corresponds to at least one of start of a combustion system, taxiing and idle operating conditions.

Assignees

• Pratt & Whitney Canada

Inventors

• Patel Bhawan B
• Morenko Oleg

Classifications

• F23C6/047 Primary

  with fuel supply in stages · CPC title

• F23N1/002

  using electronic means (F23N1/04 - F23N1/10 take precedence) · CPC title

• F23R3/346

  for staged combustion · CPC title

• F23R3/28 Primary

  characterised by the fuel supply (burners F23D) · CPC title