US20060130486A1

US20060130486A1 - Method and apparatus for assembling gas turbine engine combustors

Info

Publication number: US20060130486A1
Application number: US11/017,186
Authority: US
Inventors: Allen Danis; Timothy Held
Original assignee: Individual
Current assignee: General Electric Co
Priority date: 2004-12-17
Filing date: 2004-12-17
Publication date: 2006-06-22

Abstract

A method enables the operation of a gas turbine engine. The method comprises channeling airflow into a cooling passageway defined between the combustor casing and an inner liner of the combustor, wherein the inner liner is fabricated from a plurality of panels coupled together, channeling airflow into a cooling passageway defined between the combustor casing and an outer liner of the combustor; wherein the outer liner is fabricated from a plurality of panels coupled together, and channeling dilution airflow into a combustion chamber defined between the inner and outer liners, through a plurality of openings formed within at least one panel within at least one of the inner liner panels and the outer liner panels, wherein the plurality of openings are non-circular.

Description

BACKGROUND OF THE INVENTION

This invention relates generally to combustors and, more particularly to a method and apparatus for decreasing combustor acoustics.
At least some known gas turbine engines include a compressor for compressing air which is suitably mixed with a fuel and channeled to a combustor wherein the mixture is ignited for generating hot combustion gases. At least some known combustors include a dome assembly, a cowling, and inner and outer liners to channel the combustion gases to a turbine, which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator. The liners are coupled to the dome assembly with the cowling, and extend downstream from the cowling to define the combustion chamber. An outer support is coupled radially outward from the outer liner such that an outer cooling passage is defined radially outward from the outer liner, and an inner support is coupled radially inward from the inner liner such that an inner cooling passage is defined therebetween.
At least some known liners include a plurality of panels that are serially connected together between the upstream and aft ends of each liner such that the panels define the combustion chamber. At least some known panels are formed with primary airflow openings or secondary airflow openings. Known primary airflow openings are formed with a first diameter that is sized to enable sufficient air to enter the combustion chamber to facilitate complete oxidation of the fuel within the chamber. Known secondary airflow openings are typically formed with a smaller diameter than that of the primary airflow openings, and are positioned downstream from the primary airflow openings. The secondary airflow openings are sized to facilitate channeling airflow into the combustion chamber to facilitate diluting the combustion gases generated therein. However, the number of secondary openings that may be formed within a given panel is usually limited by structural considerations, and as such, the amount of dilution airflow that may be provided to the combustion chamber may be limited.

BRIEF DESCRIPTION OF THE INVENTION

In one aspect, a method for operating a gas turbine engine is provided. The method comprises channeling airflow into a cooling passageway defined between the combustor casing and an inner liner of the combustor, wherein the inner liner is fabricated from a plurality of panels coupled together, channeling airflow into a cooling passageway defined between the combustor casing and an outer liner of the combustor; wherein the outer liner is fabricated from a plurality of panels coupled together, and channeling dilution airflow into a combustion chamber defined between the inner and outer liners, through a plurality of openings formed within at least one panel within at least one of the inner liner panels and the outer liner panels, wherein the plurality of openings are non-circular.
In another aspect, a combustor for a gas turbine engine is provided. The combustor includes an inner liner, an outer liner, and a combustion chamber defined therebetween. The inner and outer liners each include a plurality of panels coupled together. At least one of the plurality of inner liner panels and the plurality of outer liner panels includes a plurality of openings extending therethrough for channeling dilution airflow into the combustion chamber. The plurality of openings are non-circular.
In a further aspect, a gas turbine engine is provided. The gas turbine engine includes a combustor including an inner liner, an outer liner, and a combustion chamber defined therebetween. The inner and outer liners each include a plurality of panels coupled together. At least one of the plurality of inner liner panels and the plurality of outer liner panels includes a plurality of openings extending therethrough for channeling dilution airflow into the combustion chamber. The plurality of openings are non-circular.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic illustration of an exemplary gas turbine engine;
FIG. 2 is a cross-sectional view of a combustor that may be used with the gas turbine engine;
FIG. 3 is an enlarged perspective view of a portion of a liner used with the combustor shown in FIG. 2 and taken along area 3; and
FIG. 4 is a plan view of a portion of the liner used with the combustor shown in FIG. 2 and taken along area 4.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 is a schematic illustration of an exemplary gas turbine engine 10 including a low pressure compressor 12, a high pressure compressor 14, and a combustor 16. Engine 10 also includes a high pressure turbine 18, and a low pressure turbine 20 arranged in a serial, axial flow relationship. Compressor 12 and turbine 20 are coupled by a first shaft 24, and compressor 14 and turbine 18 are coupled by a second shaft 26. In one embodiment, gas turbine engine 10 is an LMS100 engine commercially available from General Electric Company, Cincinnati, Ohio.
In operation, air flows through low pressure compressor 12 from an upstream side 28 of engine 10. Compressed air is supplied from low pressure compressor 12 to high pressure compressor 14. Highly compressed air is then delivered to combustor assembly 16 where it is mixed with fuel and ignited. Combustion gases are channeled from combustor assembly 16 to drive turbines 18 and 20.
FIG. 2 is a cross-sectional view of a combustor 30 that may be used with gas turbine engine 10. FIG. 3 is an enlarged perspective view of a portion of a liner 40 used with combustor 30 and taken along area 3. FIG. 4 is a plan view of a portion of liner 40 used with combustor 30 shown in FIG. 2 and taken along area 4. Combustor 30 includes a dome assembly 32. A fuel injector 34 extends into dome assembly 32 and injects atomized fuel through dome assembly 32 into a combustion zone or chamber 36 of combustor 30 to form an air-fuel mixture that is ignited downstream of fuel injector 34.
Combustor dome assembly 32 defines an upstream end of combustion zone 36 and includes a plurality of mixer assemblies 37 that are spaced circumferentially around combustor dome assembly 32 for delivering a mixture of fuel and air to combustion zone 36. In the exemplary embodiment, combustor dome assembly 32 is a single annular combustor (SAC) that includes one annular combustor dome. However, it should be understood that in alternative embodiments combustor dome assembly 32 may include any number of combustor domes. For example, in one embodiment, combustor dome assembly 32 is a dual annular combustor (DAC), and, in another embodiment, combustor dome assembly 32 is a triple annular combustor.
Combustion zone 36 is defined by combustor liners 40 that shield components external to combustor 30 from heat generated within combustion zone 36. Combustion zone 36 extends from dome assembly 32 downstream to a turbine nozzle assembly 41. Liners 40 include an inner liner 42 and an outer liner 44. Each liner 42 and 44 is annular and includes a plurality of separate panels 50. In the exemplary embodiment, each panel 50 includes a series of steps 52, each of which form a distinct portion of combustor liner 40.
Outer liner 44 and inner liner 42 each include a respective aft-most panel 64 and 66. Panels 64 and 66 are each located at the aft end 68 of combustion zone 36 and are adjacent turbine nozzle assembly 41. Specifically, each panel 64 and 66 couples an aft end 70 and 72 of each respective liner 44 and 42 to turbine nozzle assembly 41.
Each liner 42 and 44 also includes an annular support mount, or aft mount, 80 and 82, respectively. Specifically, each support mount 80 and 82 couples an aft end 70 and 72 of each respective liner 44 and 42 to turbine nozzle assembly 41 and to a combustor casing 84 that extends substantially circumferentially around combustor 30. More specifically, each support mount 80 and 82 extends radially outward from each respective liner 42 and 44 such that a radially outer cooling passageway 86 and a radially outer cooling passageway 88 are defined between combustor casing 84 and combustor liner 40. Accordingly, cooling passageway 86 is defined between liner 42 and combustor casing 84 and cooling passageway 88 is defined between liner 44 and combustor casing 84.
Each combustor panel 50 includes a combustor liner surface 90 and an exterior surface 92 that is radially outward from liner surface 90. When panels 50 are coupled together, combustor liner surface 90 extends generally from dome assembly 32 to turbine nozzle assembly 41. In the exemplary embodiment, each panel 50 is generally rectangular and includes a pair of circumferentially-spaced side edges 100 that are connected together by a leading edge side 102 and an opposed trailing edge side 104.
Each liner 42 and 44 also includes at least one panel 110 that is downstream from fuel injector 34, and includes a plurality of circumferentially-spaced primary airflow openings 111 that extend through panel 110 between combustor liner surface 90 and an exterior surface 92. Openings 111 are substantially circular and have a diameter D₁. In the exemplary embodiment, openings 111 extend substantially circumferentially around combustion chamber 36. Accordingly, openings 111 connect each cooling passageway 86 and 88 in flow communication with combustion chamber 36. In the exemplary embodiment, panel 110 is at least two panels 50 upstream from turbine nozzle assembly 41.
Each liner 42 and 44 also includes at least one panel 112 that is downstream from panel 110 and includes a plurality of circumferentially-spaced secondary or dilution airflow openings 116. In the exemplary embodiment, openings 116 are spaced substantially circumferentially around combustion chamber 36. Openings 116 extend through panel 112 between combustor liner surface 90 and an exterior surface 92 and are non-circular. More specifically, in the exemplary embodiment, openings 116 are substantially race-tracked shaped or generally elliptical and are defined by a pair of opposed, generally parallel sidewalls 120 that are connected by a pair of opposed arcuate sidewalls 122.
In the exemplary embodiment, sidewalls 122 are formed with a pre-determined radius of curvature R₁that is smaller than an associated radius 1/2 D₁of each primary cooling opening 111. More specifically, in the exemplary embodiment, each sidewall 122 is substantially semi-circular. Accordingly, because sidewalls 120 are substantially parallel, within each opening, sidewalls 120 are separated by the diameter D₃(twice the radius of curvature R₁) of each arcuate sidewall 122.
In the exemplary embodiment, openings 116 are oriented such that sidewalls 120 are aligned generally axially. Accordingly, a distance of separation, known as web spacing, D₂between circumferentially adjacent openings 116 is measured between adjacent opening sidewalls 120. In the exemplary embodiment, distance D₂is at least twice that of the diameter D₃of each opening 116. The distance of separation D₂facilitates maintaining structural integrity of each panel 112.
During operation, an annular diffuser 124 channels air discharged from compressor 14 into the combustor dome assemblies 32 and, more specifically, into mixer assemblies 37, wherein the compressed air is mixed with fuel provided by fuel injector 34. The fuel/air mixture is then ignited within combustion zone 36 to form combustion gases, which are discharged from the combustion zone 36 through turbine nozzle assembly 41.
A portion of the compressed air discharged from compressor 14 is channeled into each cooling passageway 86 and 88 for cooling combustor assembly 30. More specifically, the compressed air channeled through passageways 86 and 88 is also channeled into combustion zone 36 through primary cooling openings 111 defined within panels 110. The compressed air channeled through openings 111 facilitates convectively cooling liners 42 and 44 in regions adjacent openings 111. Moreover, air channeled through openings 111 also mixes with the fuel-air mixture within combustion chamber 36 to facilitate complete oxidation of all of the fuel supplied to chamber 36.
As the fuel-air mixture is channeled downstream, the mixture is mixed with air channeled through dilution openings 116. Openings 116 facilitate diluting the burned combustion products within chamber 36 to facilitate reducing the temperature of the combustion gases channeled downstream to the turbines. Moreover, the elongation of openings 116 facilitates increasing the penetration of the airflow jets discharged into chamber 36 from openings 116 in comparison to other known dilution openings, such as circular openings. The increased penetration of the dilution airflow enables openings 116 to facilitate shaping the radial temperature profile to a predetermined profile shape in an area where the mainstream velocities are relatively high.
The elongated shape of openings 116 facilitates enough penetration of the dilution air such that the air is not readily turned or forced over by the mainstream flow. Moreover, because openings 116 are oriented with their narrowest dimension in the circumferential direction, airflow discharged through openings 116 is more streamlined than airflow discharged through circular openings, which enables the airflow to penetrate the mainstream flow to a greater extent than is possible through a round opening.
The above-described gas turbine engine combustor includes a liner that includes at least one panel including a plurality of non-circular, dilution openings extending therethrough. The dilution openings are oriented such that their narrowest dimension extends circumferentially across the panel. The shape and orientation of the dilution openings enables airflow discharged from the openings to penetrate the mainstream flow to a greater extent than is possible through known round openings. As a result, the openings facilitate enhanced control of the radial temperature profile generated within the combustion chamber and increasing the useful life of the combustor in a cost-effective and reliable manner.
Exemplary embodiments of a combustor for a gas turbine engine are described above in detail. The systems and assembly components of the combustor are not limited to the specific embodiments described herein, but rather, components of each system may be utilized independently and separately from other components described herein. Each system and assembly component can also be used in combination with other combustor systems and assemblies or with other gas turbine engine components.
While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.

Claims

1. A method for operating a gas turbine engine, said method comprising:

channeling airflow into a cooling passageway defined between the combustor casing and an inner liner of the combustor, wherein the inner liner is fabricated from a plurality of panels coupled together;

channeling airflow into a cooling passageway defined between the combustor casing and an outer liner of the combustor; wherein the outer liner is fabricated from a plurality of panels coupled together; and

channeling dilution airflow into a combustion chamber defined between the inner and outer liners, through a plurality of openings formed within at least one panel within at least one of the inner liner panels and the outer liner panels, wherein the plurality of openings are non-circular.

2. A method in accordance with claim 1 wherein channeling dilution airflow into a combustion chamber further comprises channeling dilution airflow into the combustion chamber to facilitate controlling an exit temperature profile of the combustor.

3. A method in accordance with claim 1 wherein channeling dilution airflow into a combustion chamber further comprises channeling dilution airflow into the combustion chamber through the plurality of openings, wherein the openings are shaped to enable cooling air to penetrate into the combustion chamber to facilitate achieving a desired radial temperature profile within the combustion chamber.

4. A method in accordance with claim 1 wherein channeling dilution airflow into a combustion chamber further comprises channeling dilution airflow into the combustion chamber through the plurality of openings, wherein the openings are generally elliptically shaped.

5. A method in accordance with claim 1 wherein channeling dilution airflow into a combustion chamber further comprises channeling dilution airflow into the combustion chamber through the plurality of openings, wherein the openings are defined by a pair of substantially parallel walls that are connected together by a pair of opposed arcuate sidewalls formed with a predetermined radius of curvature.

6. A combustor for a gas turbine engine, said combustor comprising:

an inner liner comprising a plurality of panels coupled together;

an outer liner comprising a plurality of panels coupled together; and

a combustion chamber defined between said inner and outer liners, at least one of said plurality of inner liner panels and said plurality of outer liner panels comprises a plurality of openings extending therethrough for channeling dilution airflow into said combustion chamber, said plurality of openings are non-circular.

7. A combustor in accordance with claim 6 wherein said plurality of openings facilitate controlling an exit temperature profile of said combustor.

8. A combustor in accordance with claim 6 wherein said plurality of openings are each substantially elliptically-shaped.

9. A combustor in accordance with claim 6 wherein said plurality of openings are shaped to enable cooling air to penetrate into said combustion chamber to facilitate achieving a desired radial temperature profile within said combustion chamber.

10. A combustor in accordance with claim 6 wherein said plurality of openings are defined by a pair of opposed substantially parallel sidewalls connected together by a pair of opposed arcuate walls formed with a pre-determined radius.

11. A combustor in accordance with claim 10 wherein adjacent of said plurality of openings are separated by a distance that is approximately equal to twice the diameter of said arcuate walls.

12. A combustor in accordance with claim 6 wherein said at least one panel comprises a pair of opposed circumferential edges coupled together by a leading edge and a side edge, said plurality of openings comprises at least three openings spaced approximately equi-distantly between said pair of opposed circumferential edges.

13. A gas turbine engine comprising a combustor comprising an inner liner, an outer liner, and a combustion chamber defined between said inner and outer liners, each of said inner and outer liners comprises a plurality of panels coupled together, at least one of said panels within at least one of said inner liner and said outer liner comprises a plurality of openings extending therethrough for channeling dilution air into said combustion chamber, said plurality of openings are non-circular.

14. A gas turbine engine in accordance with claim 13 wherein said combustor plurality of openings extending through said at least one panel facilitate controlling an exit temperature profile of said combustor.

15. A gas turbine engine in accordance with claim 14 wherein said combustor plurality of openings extending through said at least one panel are each generally elliptically-shaped.

16. A gas turbine engine in accordance with claim 14 wherein said combustor plurality of openings extending through said at least one panel are shaped to enable cooling air to penetrate into said combustion chamber to facilitate achieving a desired radial temperature profile within said combustion chamber.

17. A gas turbine engine in accordance with claim 14 wherein said combustor plurality of openings extending through said at least one panel are defined by a pair of opposed substantially parallel sidewalls that are connected together by a pair of opposed arcuate walls formed with a pre-determined radius.

18. A gas turbine engine in accordance with claim 17 wherein adjacent of said plurality of openings extending through said at least one panel are separated within said panel by a distance that is approximately equal to twice the diameter of said arcuate walls.

19. A gas turbine engine in accordance with claim 14 wherein each of said plurality of panels comprises a pair of opposed circumferential edges coupled together by a leading edge and a side edge, said plurality of openings extending through said at least one panel comprises at least three openings spaced approximately equi-distantly between said pair of opposed circumferential edges.