U.S. patent application number 11/017186 was filed with the patent office on 2006-06-22 for method and apparatus for assembling gas turbine engine combustors.
Invention is credited to Allen Michael Danis, Timothy James Held.
Application Number | 20060130486 11/017186 |
Document ID | / |
Family ID | 36593980 |
Filed Date | 2006-06-22 |
United States Patent
Application |
20060130486 |
Kind Code |
A1 |
Danis; Allen Michael ; et
al. |
June 22, 2006 |
Method and apparatus for assembling gas turbine engine
combustors
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.
Inventors: |
Danis; Allen Michael;
(Mason, OH) ; Held; Timothy James; (Blanchester,
OH) |
Correspondence
Address: |
JOHN S. BEULICK (12729);C/O ARMSTRONG TEASDALE LLP
ONE METROPOLITAN SQUARE
SUITE 2600
ST. LOUIS
MO
63102-2740
US
|
Family ID: |
36593980 |
Appl. No.: |
11/017186 |
Filed: |
December 17, 2004 |
Current U.S.
Class: |
60/752 |
Current CPC
Class: |
Y02T 50/675 20130101;
Y02T 50/60 20130101; F23R 3/06 20130101; F23R 3/002 20130101; F23R
2900/00014 20130101 |
Class at
Publication: |
060/752 |
International
Class: |
F23R 3/42 20060101
F23R003/42 |
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.
Description
BACKGROUND OF THE INVENTION
[0001] This invention relates generally to combustors and, more
particularly to a method and apparatus for decreasing combustor
acoustics.
[0002] 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.
[0003] 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
[0004] 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.
[0005] 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.
[0006] 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
[0007] FIG. 1 is a schematic illustration of an exemplary gas
turbine engine;
[0008] FIG. 2 is a cross-sectional view of a combustor that may be
used with the gas turbine engine;
[0009] 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
[0010] 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
[0011] 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.
[0012] 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.
[0013] 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.
[0014] 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.
[0015] 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.
[0016] 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.
[0017] 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.
[0018] 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.
[0019] 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 D.sub.1. 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.
[0020] 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.
[0021] In the exemplary embodiment, sidewalls 122 are formed with a
pre-determined radius of curvature R.sub.1 that is smaller than an
associated radius 1/2 D.sub.1 of 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 D.sub.3 (twice the radius of curvature
R.sub.1) of each arcuate sidewall 122.
[0022] 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, D.sub.2 between
circumferentially adjacent openings 116 is measured between
adjacent opening sidewalls 120. In the exemplary embodiment,
distance D.sub.2 is at least twice that of the diameter D.sub.3 of
each opening 116. The distance of separation D.sub.2 facilitates
maintaining structural integrity of each panel 112.
[0023] 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.
[0024] 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.
[0025] 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.
[0026] 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.
[0027] 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.
[0028] 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.
[0029] 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.
* * * * *