Counter swirl doublet combustor

Patent metadata

Field	Value
Publication number	US-9765969-B2
Application number	US-201314137391-A
Country	US
Kind code	B2
Filing date	Dec 20, 2013
Priority date	Mar 15, 2013
Publication date	Sep 19, 2017
Grant date	Sep 19, 2017

Abstract

An improved combustor for a gas turbine is operable to provide high combustion efficiency in a compact combustion chamber. The combustor includes a counter swirl doublet for improved fuel/air mixing. The enhanced combustor assembly and method of operation improves operation of the turbine.

First claim

What is claimed is:

1. A combustor comprising:
an inner liner and an outer liner extending circumferentially around an axis of rotation of an engine spaced apart from one another to form a combustion chamber therebetween;
a bulkhead extending between the inner and outer liners proximate a first end of the combustor;
a plurality of fuel nozzle ports for receiving fuel nozzles therein, each of the plurality of fuel nozzle ports having a centerline axis extending axially therethrough;
a first outer primary inlet hole and a second outer primary inlet hole formed as a first doublet pair of circumferentially spaced chutes that are integrally connected at a first flange attached to the outer liner, the first doublet pair extending radially inward from a first circumferential plane through the outer liner to a first side of each respective centerline axis; and
a first inner primary inlet hole and a second inner primary inlet hole formed as a second doublet pair of circumferentially spaced chutes that are integrally connected at a second flange attached to the inner liner, the second doublet pair extending radially outward from a second circumferential plane through the inner liner to a second side of each respective centerline axis.

2. The combustor of claim 1 , further comprising:
a plurality of fuel nozzles engageable with a respective one of the plurality of fuel nozzle ports.

3. The combustor of claim 1 , wherein fuel is injected into the combustor with a swirling motion about the centerline axis of at least one fuel nozzle port.

4. The combustor of claim 1 , wherein primary combustion air is injected through at least one of the first and second doublet pairs such that an airflow creates a counter swirl relative to swirl of a fuel.

5. The combustor of claim 1 , wherein the inner and outer liners include a plurality of effusion holes formed downstream of at least one of the first and second doublet pairs.

6. The combustor of claim 5 , wherein the plurality of effusion holes receive a cooling fluid from a cooling fluid source and fluidly connect the cooling fluid source to the combustion chamber.

7. The combustor of claim 1 , wherein at least one of the inner liner with the second doublet pair or the outer liner with the first doublet pair includes a plurality of thermal resistant tiles.

8. The combustor of claim 1 , wherein the combustor includes a portion made from at least one of a metal, ceramic or intermetal based material.

9. The combustor of claim 1 , wherein at least one of the inner liner with the second doublet pair or the outer liner with the first doublet pair includes dual walls spaced apart from one another.

10. The combustor of claim 9 , wherein the dual walls include a fluid passageway for passage of a cooling fluid flow therethrough.

11. The combustor of claim 1 , wherein at least one of the chutes of the first and second doublet pairs extend into the combustion chamber.

12. The combustor of claim 1 , wherein an axial length of the combustion chamber is two times a radial height of the bulkhead.

13. The combustor of claim 1 , wherein at least one of the first outer primary inlet hole, second outer primary inlet hole, first inner primary inlet hole, or second inner primary inlet hole includes an oval or elliptical shape.

14. The combustor of claim 1 , further including one or more additional holes in at least one of the inner liner or the outer liner.

15. A gas turbine engine comprising:
a compressor section for compressing ambient air;
a combustor section for receiving compressed air and mixing a fuel for combustion;
a turbine section positioned downstream of the combustor section for receiving hot exhaust gases formed in a combustion chamber; and
wherein the combustor section includes a combustor having:
an inner liner and an outer liner extending circumferentially around an axis of rotation of the gas turbine engine and spaced apart from one another to form the combustion chamber therebetween;
a bulkhead extending between the inner and outer liners proximate a first end of the combustor;
a plurality of fuel nozzle ports positioned in the bulkhead and receiving corresponding fuel nozzles therein, each of the plurality of fuel nozzle ports having a centerline axis extending axially therethrough;
a first outer primary inlet hole and a second outer primary inlet hole formed as a first doublet pair of circumferentially spaced chutes that are integrally connected at a first flange attached to the outer liner, the first doublet pair extending radially inward from a first circumferential plane through the outer liner positioned to a first side of at least one of the corresponding fuel nozzles; and
a first inner primary inlet hole and a second inner primary inlet hole formed as a second doublet pair of circumferentially spaced chutes that are integrally connected at a second flange attached to the inner liner, the second doublet pair extending radially outward from a second circumferential plane through the inner liner positioned on an opposing side of the at least one of the corresponding fuel nozzles.

16. The gas turbine engine of claim 15 , wherein the combustor produces a temperature profile in an exit plane to minimize overheating.

17. The gas turbine engine of claim 15 , wherein combustion air flows through at least one of the chutes of the first and second doublet pairs.

18. The gas turbine engine of claim 15 , wherein a fuel has a swirl around the centerline axis of at least one of the plurality of fuel nozzle ports.

19. The gas turbine engine of claim 15 , wherein an axial length of the combustor is two times a radial height of the bulkhead.

20. A method comprising:
injecting a fuel spray into a combustion chamber with a swirl in a first circumferential direction;
injecting primary combustion air through a first outer primary inlet hole and a second outer primary inlet hole formed as a first doublet pair of circumferentially spaced chutes that are integrally connected at a first flange attached to an outer liner, the first doublet pair extending radially inward from a first circumferential plane through the outer liner adjacent a first side of the fuel spray; and
injecting the primary combustion air through a first inner primary inlet hole and a second inner primary inlet hole formed as a second doublet pair of circumferentially spaced chutes that are integrally connected at a second flange attached to an inner liner, the second doublet pair extending radially outward from a second circumferential plane through the inner liner adjacent a second side of the fuel spray;
wherein the primary combustion air is injected in an opposite direction to the first circumferential direction of the fuel spray.

Assignees

• Rolls Royce Corp

Inventors

• Graves Charles B
• Wicksall Donald M
• Richardson Jr Thomas F

Classifications

• F02C3/00

  Gas-turbine plants characterised by the use of combustion products as the working fluid (generated by intermittent combustion F02C5/00) · CPC title

• F23R3/002

  Wall structures (F23R3/02 and F23R3/007 take precedence) · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F05D2240/35

  Combustors or associated equipment · CPC title

• F23R3/06 Primary

  Arrangement of apertures along the flame tube · CPC title