Gas turbine engine system and an associated method thereof

1. A gas turbine engine system, comprising:

a compressor,

a combustor coupled to the compressor and disposed downstream of the compressor; wherein the combustor comprises:

a primary combustor section;

a secondary combustor section coupled to the primary combustor section and disposed downstream of the primary combustor section;

a transition nozzle coupled to the secondary combustor section and disposed downstream of the secondary combustor section;

an injector coupled to the secondary combustor section for injecting an air-fuel mixture to the secondary combustor section; and

an inlet of a turbine coupled to the combustor and disposed downstream of the transition nozzle; wherein an exit of the transition nozzle is oriented substantially tangentially to a circumference of the inlet of the turbine.

2. The gas turbine engine system of claim 1 , wherein the injector is airfoil shaped.

3. The gas turbine engine system of claim 1 , wherein the injector has curved portion within the second stage combustion section for enhancing distribution of the air-fuel mixture.

4. The gas turbine engine system of claim 1 , wherein the transition nozzle extends from a point corresponding to a location at which a combustion flame is generated in the secondary combustion section, towards the turbine.

5. The gas turbine system of claim 1 , wherein the injector is configured to inject the air-fuel mixture non-uniformly to the secondary combustor section.

6. The gas turbine system of claim 5 , wherein the injector is configured to generate a hot spot away from a plurality of side walls of the transition nozzle, and at a center portion of an exit of the transition nozzle.

7. The gas turbine engine system of claim 1 , further comprising an integrated vane disposed within the transition nozzle such that the transition nozzle is formed into a plurality of transition sections.

8. The gas turbine engine system of claim 7 , wherein the integrated vane is used to feed cooling air from the compressor into the transition nozzle.

9. The gas turbine system of claim 7 , wherein the injector is configured to generate a hot spot away from a plurality of side walls of the transition nozzle and a peripheral surface the integrated vane, at a center portion of an exit of the transition nozzle.

10. The gas turbine system of claim 1 , wherein the transition nozzle is used to inject a hot combustion gas substantially tangential to the turbine.

11. The gas turbine system of claim 1 , wherein the combustor has a central axis oriented at a predefined angle to the turbine.

12. The gas turbine system of claim 1 , wherein the transition nozzle provides a turning of flow of a hot combustion gas substantially tangential to the turbine.

13. A method comprising:

combusting a first air-fuel mixture in a primary combustor section and generating a first combustion gas;

feeding the first combustion gas from the primary combustor section to a secondary combustor section;

injecting a second air-fuel mixture via an injector coupled to the secondary combustor section to combust the second air-fuel mixture and generate a second combustion gas; and

injecting the first combustion gas and the second combustion gas via a transition nozzle substantially tangentially to a circumference of an inlet of a turbine.

14. The method of claim 13 , wherein the first air-fuel mixture comprises a lean air-fuel mixture.

15. The method of claim 14 , wherein the second air-fuel mixture comprises a rich air-fuel-diluent mixture.

16. The method of claim 13 , further comprising feeding cooling air from a compressor, via an integrated vane disposed in the transition nozzle.

17. The method of claim 13 , wherein injecting the second air-fuel mixture comprises injecting the second air-fuel mixture non-uniformly via the injector to the secondary combustor section.

18. The method of claim 17 , further comprising generating a hot spot away from a plurality of side walls of the transition nozzle, and at a center portion of an exit of the transition nozzle.

19. The method of claim 13 , wherein injecting the first combustion gas and the second combustion gas comprises providing a turning of flow of the first combustion gas and the second combustion gas substantially tangential to the turbine via the transition nozzle.

20. The method of claim 13 , wherein injecting the first and second combustion gases further comprises providing a turning of flow of the first and second combustion gases substantially tangential to the turbine via an integrated vane disposed in the transition nozzle.

21. The method of claim 20 , further comprising generating a hot spot away from a plurality of side walls of the transition nozzle and a peripheral surface of the integrated vane, at a center portion of an exit of the transition nozzle.

22. A method, comprising:

forming a combustor downstream of a compressor; wherein the combustor comprises a primary combustor section; a secondary combustor section coupled to the primary combustor section and disposed downstream of the primary combustor section, and a transition nozzle coupled to the secondary combustor section and disposed downstream of the secondary combustor section, such that an exit of the transition nozzle is oriented substantially tangentially to a circumference of an inlet of a turbine; and

coupling an injector to the secondary combustor section.

23. The method of claim 22 , wherein the injector is airfoil shaped.

24. The method of claim 22 , wherein the injector has a curved portion disposed within the secondary combustor section for enhancing distribution of an air-fuel mixture.

25. The method of claim 22 , further comprising forming an integrated vane within the transition nozzle such that the transition nozzle is formed into a plurality of transition sections.

26. The method of claim 22 , further comprising orienting a central axis of the combustor at a predefined angle to the turbine.