U.S. patent application number 14/988068 was filed with the patent office on 2017-07-06 for cooled combustor for a gas turbine engine.
The applicant listed for this patent is General Electric Company. Invention is credited to William Thomas Bennett.
Application Number | 20170191664 14/988068 |
Document ID | / |
Family ID | 57614195 |
Filed Date | 2017-07-06 |
United States Patent
Application |
20170191664 |
Kind Code |
A1 |
Bennett; William Thomas |
July 6, 2017 |
COOLED COMBUSTOR FOR A GAS TURBINE ENGINE
Abstract
A combustor for a gas turbine engine comprises a liner having a
forward opening and an aft opening. A deflector closes the forward
opening and the aft opening is in fluid communication with a
turbine section downstream of the combustor. A dome disposed in the
deflector provides a fuel/air mixture to be ignited to drive the
turbine section. A gap can be disposed between the deflector and
the liner, having film holes disposed in the deflector for
providing a flow of cool air to the gap for cooling the combustion
liner downstream of the gap.
Inventors: |
Bennett; William Thomas;
(Cincinnati, OH) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company |
Schenectady |
NY |
US |
|
|
Family ID: |
57614195 |
Appl. No.: |
14/988068 |
Filed: |
January 5, 2016 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
Y02T 50/60 20130101;
F23R 3/002 20130101; Y02T 50/675 20130101; F23R 3/06 20130101; F23R
3/10 20130101; F23R 3/26 20130101; F23R 2900/03042 20130101; F23R
2900/03044 20130101 |
International
Class: |
F23R 3/00 20060101
F23R003/00; F23R 3/26 20060101 F23R003/26 |
Claims
1. A method of cooling a combustor for a gas turbine engine, the
method comprising: supplying an impinging airflow onto a deflector
of the combustor; and after impingement onto the deflector,
supplying the impinging airflow through film holes in the combustor
to define a film of cooling air along at least a portion of a
liner.
2. The method of claim 1 wherein the supplying an impinging airflow
comprises supplying cooling air through an impingement baffle
spaced from the deflector.
3. The method of claim 2 wherein the supplying an impinging airflow
comprises supplying compressor air from a compressor section of the
gas turbine engine.
4. The method of claim 2 wherein supplying the impinging airflow
through the film holes comprises supplying impinging airflow
through film holes passing through at least one of the deflector
and liner.
5. The method of claim 4 wherein the film holes pass through both
the deflector and liner.
6. The method of claim 1 wherein the film of cooling air flows
along an interior of the liner.
7. A combustor for a gas turbine engine comprising: a liner
defining a forward opening and an aft opening; a deflector closing
the forward opening; a dome located in the deflector; and film
holes extending through the combustor from the deflector to an
interior of the liner.
8. The combustor of claim 7 wherein the film holes extend through
at least one of the deflector and the liner.
9. The combustor of claim 8 wherein the film holes extend through
both the deflector and liner.
10. The combustor of claim 8 further comprising an impingement
baffle upstream of the deflector.
11. The combustor of claim 10 wherein the impingement baffle and
deflector at least partially define an impingement chamber and the
film holes are fluidly coupled to the impingement chamber.
12. The combustor of claim 11 wherein the deflector terminates
prior to the liner to define a gap and the film holes are fluidly
coupled to the gap.
13. The combustor of claim 7 further comprising an impingement
baffle upstream of the deflector.
14. The combustor of claim 13 wherein the impingement baffle and
deflector at least partially define an impingement chamber and the
film holes are fluidly coupled to the impingement chamber.
15. The combustor of claim 7 wherein the deflector terminates prior
to the liner to define a gap and the film holes are fluidly coupled
to the gap.
16. A combustor for a gas turbine engine comprising a deflector, a
liner spaced from the deflector and defining a gap there between,
and at least one film hole extending through the deflector and
having an outlet opening into the gap.
17. The combustor of claim 16 further comprising an impingement
baffle upstream of the deflector.
18. The combustor of claim 17 wherein the impingement baffle and
deflector at least partially define an impingement chamber and the
at least one film hole has an inlet fluidly coupled to the
impingement chamber.
19. The combustor of claim 16 wherein the at least one film hole
extends through at least one of the deflector and the liner.
20. The combustor of claim 16 wherein the at least one film hole
extend primarily radially through the deflector.
Description
BACKGROUND OF THE INVENTION
[0001] Gas turbine engines are rotary engines that extract energy
from a flow of combusted gases passing through the engine onto a
multitude of rotating turbine blades. Gas turbine engines have been
used for land and nautical locomotion and power generation, but are
most commonly used for aeronautical applications such as for
aircraft, including helicopters. In aircraft, gas turbine engines
are used for propulsion of the aircraft. In terrestrial
applications, turbine engines are often used for power
generation.
[0002] Gas turbine engines for aircraft are designed to operate at
high temperatures to maximize engine efficiency, so cooling of
certain engine components, such as the high pressure turbine and
the low pressure turbine, can be beneficial. Typically, cooling is
accomplished by ducting cooler air from the high and/or low
pressure compressors to the engine components that require cooling.
Temperatures in the high pressure turbine are around 1000.degree.
C. to 2000.degree. C. and the cooling air from the compressor is
around 500.degree. C. to 700.degree. C. While the compressor air is
a high temperature, it is cooler relative to the turbine air, and
can be used to cool the turbine.
[0003] Contemporary combustors have liners to define the combustion
chamber for burning fuel upstream from the turbine. The liners can
be cooled with a flow of cooling air from a combination of nugget
hole cooling and or arrays of film holes. However, arrays of film
holes alone can fail to properly cool the liner and nugget cooling
involves heightened cost and is prone to failure.
BRIEF DESCRIPTION OF THE INVENTION
[0004] A method of cooling a combustor for a gas turbine engine
comprising supplying an impinging airflow onto a deflector of the
combustor and after impingement onto the deflector, supplying the
impinging airflow through film holes in the combustor to define a
film of cooling air along at least a portion of the liner.
[0005] A combustor for a gas turbine engine comprising a liner
defining a forward opening and an aft opening, a deflector closing
the forward opening, a dome located in the deflector, and film
holes extending through the combustor deflector to an interior of
the liner.
[0006] A combustor for a gas turbine engine comprising a deflector,
a liner spaced from the deflector and defining a gap therebetween,
and at least one film holes extending through the deflector and
having an outlet opening into the gap.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] In the drawings:
[0008] FIG. 1 is a schematic cross-sectional diagram of a gas
turbine engine for an aircraft.
[0009] FIG. 2 is a schematic cross-sectional view of the combustor
section transitioning into a turbine section of the engine of FIG.
1.
[0010] FIG. 3 is a perspective view of the combustor of FIG. 2.
[0011] FIG. 4 is an exploded view of the combustor of FIG. 3.
[0012] FIG. 5 is a cross-sectional view of the deflector section of
the combustor of FIG. 4.
[0013] FIG. 6 is the cross-sectional view of the deflector section
of FIG. 5 illustrating fluid flow paths.
DESCRIPTION OF EMBODIMENTS OF THE INVENTION
[0014] The described embodiments of the present invention are
directed to a turbine combustor, and in particular to cooling a
combustion liner. For purposes of illustration, the present
invention will be described with respect to a combustor for an
aircraft gas turbine engine. It will be understood, however, that
the invention is not so limited and can have general applicability
in non-aircraft applications, such as other mobile applications and
non-mobile industrial, commercial, and residential
applications.
[0015] As used herein, the terms "axial" or "axially" refer to a
dimension along a longitudinal axis of an engine. The term
"forward" or "upstream" used in conjunction with "axial" or
"axially" refers to moving in a direction toward the engine inlet,
or a component being relatively closer to the engine inlet as
compared to another component. The term "aft" or "downstream" used
in conjunction with "axial" or "axially" refers to a direction
toward the rear or outlet of the engine relative to the engine
centerline.
[0016] As used herein, the terms "radial" or "radially" refer to a
dimension extending between a center longitudinal axis of the
engine and an outer engine circumference. The use of the terms
"proximal" or "proximally," either by themselves or in conjunction
with the terms "radial" or "radially," refers to moving in a
direction toward the center longitudinal axis, or a component being
relatively closer to the center longitudinal axis as compared to
another component. The use of the terms "distal" or "distally,"
either by themselves or in conjunction with the terms "radial" or
"radially," refers to moving in a direction toward the outer engine
circumference, or a component being relatively closer to the outer
engine circumference as compared to another component.
[0017] All directional references (e.g., radial, axial, proximal,
distal, upper, lower, upward, downward, left, right, lateral,
front, back, top, bottom, above, below, vertical, horizontal,
clockwise, counterclockwise, upstream, downstream, aft, etc.) are
only used for identification purposes to aid the reader's
understanding, and do not create limitations, particularly as to
the position, orientation, or use. Connection references (e.g.,
attached, coupled, connected, and joined) are to be construed
broadly and can include intermediate members between a collection
of elements and relative movement between elements unless otherwise
indicated. As such, connection references do not necessarily infer
that two elements are directly connected and in fixed relation to
each other. The exemplary drawings are for purposes of illustration
only and the dimensions, positions, order and relative sizes
reflected in the drawings attached hereto can vary.
[0018] FIG. 1 is a schematic cross-sectional diagram of a gas
turbine engine 10 for an aircraft. The engine 10 has a generally
longitudinally extending axis or centerline 12 extending forward 14
to aft 16. The engine 10 includes, in downstream serial flow
relationship, a fan section 18 including a fan 20, a compressor
section 22 including a booster or low pressure (LP) compressor 24
and a high pressure (HP) compressor 26, a combustion section 28
including a combustor 30, a turbine section 32 including a HP
turbine 34, and a LP turbine 36, and an exhaust section 38.
[0019] The fan section 18 includes a fan casing 40 surrounding the
fan 20. The fan 20 includes a plurality of fan blades 42 disposed
radially about the centerline 12. The HP compressor 26, the
combustor 30, and the HP turbine 34 form a core 44 of the engine
10, which generates combustion gases. The core 44 is surrounded by
core casing 46, which can be coupled with the fan casing 40.
[0020] A HP shaft or spool 48 disposed coaxially about the
centerline 12 of the engine 10 drivingly connects the HP turbine 34
to the HP compressor 26. A LP shaft or spool 50, which is disposed
coaxially about the centerline 12 of the engine 10 within the
larger diameter annular HP spool 48, drivingly connects the LP
turbine 36 to the LP compressor 24 and fan 20. The portions of the
engine 10 mounted to and rotating with either or both of the spools
48, 50 are also referred to individually or collectively as a rotor
51.
[0021] The LP compressor 24 and the HP compressor 26 respectively
include a plurality of compressor stages 52, 54, in which a set of
compressor blades 58 rotate relative to a corresponding set of
static compressor vanes 60, 62 (also called a nozzle) to compress
or pressurize the stream of fluid passing through the stage. In a
single compressor stage 52, 54, multiple compressor blades 56, 58
can be provided in a ring and can extend radially outwardly
relative to the centerline 12, from a blade platform to a blade
tip, while the corresponding static compressor vanes 60, 62 are
positioned downstream of and adjacent to the rotating blades 56,
58. It is noted that the number of blades, vanes, and compressor
stages shown in FIG. 1 were selected for illustrative purposes
only, and that other numbers are possible. The blades 56, 58 for a
stage of the compressor can be mounted to a disk 53, which is
mounted to the corresponding one of the HP and LP spools 48, 50,
with each stage having its own disk. The vanes 60, 62 are mounted
to the core casing 46 in a circumferential arrangement about the
rotor 51.
[0022] The HP turbine 34 and the LP turbine 36 respectively include
a plurality of turbine stages 64, 66, in which a set of turbine
blades 68, 70 are rotated relative to a corresponding set of static
turbine vanes 72, 74 (also called a nozzle) to extract energy from
the stream of fluid passing through the stage. In a single turbine
stage 64, 66, multiple turbine blades 68, 70 can be provided in a
ring and can extend radially outwardly relative to the centerline
12, from a blade platform to a blade tip, while the corresponding
static turbine vanes 72, 74 are positioned upstream of and adjacent
to the rotating blades 68, 70. It is noted that the number of
blades, vanes, and turbine stages shown in FIG. 1 were selected for
illustrative purposes only, and that other numbers are
possible.
[0023] In operation, the rotating fan 20 supplies air to the LP
compressor 24, which then supplies pressurized air to the HP
compressor 26, which further pressurizes the air. The pressurized
air from the HP compressor 26 is mixed with fuel in the combustor
30 and ignited, thereby generating combustion gases. Some work is
extracted from these gases by the HP turbine 34, which drives the
HP compressor 26. The combustion gases are discharged into the LP
turbine 36, which extracts additional work to drive the LP
compressor 24, and the exhaust gas is ultimately discharged from
the engine 10 via the exhaust section 38. The driving of the LP
turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP
compressor 24.
[0024] Some of the air supplied by the fan 20 can bypass the engine
core 44 and be used for cooling of portions, especially hot
portions, of the engine 10, and/or used to cool or power other
aspects of the aircraft. In the context of a turbine engine, the
hot portions of the engine are normally downstream of the combustor
30, especially the turbine section 32, with the HP turbine 34 being
the hottest portion as it is directly downstream of the combustion
section 28. Other sources of cooling fluid can be, but is not
limited to, fluid discharged from the LP compressor 24 or the HP
compressor 26.
[0025] FIG. 2 is a side section view of the combustor 30 of the
engine 10 of FIG. 1. The combustor 30 includes a combustion liner
76 and a deflector assembly 78 to define a combustion chamber 80.
The deflector assembly 78 comprises a plurality of annularly
arranged dome 92 for introducing a fuel/air mixture into the
combustion chamber 80.
[0026] A bypass channel 86 can be disposed both radially inside of
and outside of the combustion chamber 80. The bypass channels 86
provide a flow of fluid 88 from the compressor section 26 to the
turbine section 34 bypassing the combustor 30 through a set of
openings 90. Additionally, the bypass channels 86 can provide a
flow of cooling fluid to the combustion chamber 80 through a
plurality of film holes 91 for providing surface film cooling along
the interior surface of the combustion liner 76 relative to the
combustion chamber 80.
[0027] The dome 92 can be coupled to a fuel line 94 mounted to the
combustor 30 at a mount 96 and fed with a flow of fuel from the
fuel line 94 for emitting and igniting a fuel/air mixture into the
combustion chamber 80. The fuel line 94 is inserted through a cowl
82, which can mount the dome 92 and forward section of the
combustor 30 to the core casing 46.
[0028] Turning to FIG. 3, a perspective view of the combustor 30
illustrates the cowl 82 having a body 110. An arm 112 extends from
the body 110, terminating at a mount 114 having an aperture 116 for
mounting the combustor 30 to the core casing with a fastener, such
as a bolt. The cowl 82 couples to an impingement baffle 120. The
impingement baffle 120 couples to a deflector 122, having a
deflector surface 124 defining the surface of the deflector
assembly 78 confronting the combustion chamber 80. The deflector
surface 124 can be made of a material suitable for use in high
temperature regions typical in a combustor 30. The deflector 122
comprises two extensions 128 adapted to receive the body 110 of the
cowl 82.
[0029] The combustion liner 76 comprises a radially outer liner 132
and a radially inner liner 130, each liner 130, 132 having a
plurality of film holes 134 to provide surface cooling along the
interior of the liner 76 confronting the combustion chamber 80. The
combustion liner 76 can define a forward opening 140 and an aft
opening 142, having the deflector 122 closing the forward opening
140. The radially outer and inner liners 130, 132 comprise forward
extensions 136, extending over the impingement baffle extensions
128. The combination of the body 110, the baffle extensions 128,
and the liner extensions 136 overlay one another, each comprising
one or more fastener apertures 138 for coupling the cowl 82,
impingement baffle 120, and the liners 130, 132.
[0030] Looking at FIG. 4, an exploded view best illustrates a
plurality of apertures as impingement holes 150 disposed in the
impingement baffle 120 and film holes 160 in the deflector 122. The
cowl 82 further comprises an opening 144. The opening 144 receives
the insertion of the fuel line 94 as well as provides space for a
flow of fluid F to be provided into the interior of the cowl 82.
The impingement baffle 120 comprises a plurality of impingement
holes 150 arranged along an impingement wall 152. The impingement
holes 150 are arranged in two rows being disposed radially above
and below the dome 92. The arrangement of the impingement holes 150
is exemplary and should not be construed as limiting. The
impingement holes 150 can be alternatively arranged around the dome
92, spaced, staggered, offset, or any other arrangement.
[0031] The deflector 122 comprises a deflector wall 154
transitioning to sidewalls 156 on the sides of the deflector 122.
One or more fillets 158 can provide a transition between the
deflector wall 154 and the sidewalls 156. A plurality of film holes
160 are disposed in the fillets 158 extending to a radially outer
surface 162 and a radially inner surface 164 of the deflector 122.
While the film holes 160 are illustrated as disposed in the fillets
158, it is also contemplated that the deflector 122 does not
comprise fillets 158 and the film holes 160 extend through the
sidewalls 156 to the outer surface 162 and the inner surface 164 of
the deflector 122. Thus, the film holes 160 can extend
substantially radially relative to the engine centerline to the
radially inner and outer surfaces 162, 164 of the deflector 122.
Furthermore, the film holes can be any shape, comprising circular,
oval, elliptical, quadrilateral, or otherwise.
[0032] Looking now at FIG. 5, a section view of the coupled cowl
82, impingement baffle 120, deflector 122, and liners 130, 132 best
illustrates a plurality of internal chambers. A housing chamber 170
can be defined within the cowl 82 and forward of the impingement
baffle 120. A set of impingement chambers 172 are defined between
the impingement baffle 120 and the deflector 122 and are in fluid
communication with the housing chamber 170 through the impingement
holes 150. A set of gaps 174 are defined between the outer surface
162 and the outer liner 132 and between the inner surface 164 and
the inner liner 130. The gaps 174 are in fluid communication with
the impingement chambers 172 through the film holes 160, with the
film holes 160 having inlets adjacent the impingement chamber 172
and outlets at the gap 174. The gaps 174 are further in fluid
communication with the combustion chamber 80 downstream of the
impingement baffle 120 and the deflector 122. The gaps 174 can be
annular, extending circumferentially around the engine centerline
12 with the combustor 30. In one alternative example, the film
holes 160 can pass through both the deflector 122 and the liners
130, 132.
[0033] Turning now to FIG. 6, an airflow at 180 can be provided to
the housing chamber 170 through the opening 144 from the compressor
section upstream of the combustor. At 182, the airflow within the
housing chamber 170 can be provided to the impingement baffle 120.
The airflow at 184 passes through the impingement holes 150 into
the impingement chamber 172 where it can impinge on the deflector
122 to cool the deflector 122. From the impingement chamber 172, at
186 the airflow is provided through the film holes 160 to the gap
174 where the airflow can pass over the internal surface of the
combustion liner 76 as a cooling film 188. Thus, a flow of cooling
fluid is provided through the deflector assembly 78 to form a
cooling film over the internal surface of the combustion liner 76
at the deflector 122, providing a stronger initial cooling film
adjacent to the deflector assembly 78. In an alternative example
where the film holes 160 pass through both the deflector 122 and
the liner 130, 132, the cooling film 188 can be provided to the
exterior of the liner 130, 132 relative to the combustion chamber
80.
[0034] A method for cooling the combustor 30 of the gas turbine
engine 10 can comprise supplying an impinging airflow onto a
deflector 122 of the combustor 30 and after impingement onto the
deflector 122, supplying the impinging airflow through the film
holes 134 in the combustor 30 to define a film of cooling air along
at least a portion of the liner 76. The airflow can be supplied
from the housing chamber 170 through the impingement holes 150 as
an impinging airflow to the deflector 122 within the impingement
chamber 172. From the impingement chamber 172, the impinging
airflow is fed through the film holes 160 to the gap 174 where the
airflow can be directed along a portion of the liner 76 as a
cooling film. The airflow being supplied can comprise compressor
air from the compressor section 22 of the engine 10. The airflow
provided along the liner 76 can be fed along the interior of the
liner relative to the combustion chamber 80 defined by the liner
76.
[0035] It should be appreciated the improved film cooling can be
achieved by directing a flow of air along the combustion liner from
a gap in the deflector and the liner. Directing the airflow to the
deflector can further cool the deflectors and utilize spent
impingement air to further cool the liner. The improved cooling
film can reduce or eliminate the need for film nuggets on the
liner, which have a tendency to fail or reduce time-on-wing for the
combustor or combustion liner.
[0036] This written description uses examples to disclose the
invention, including the best mode, and to enable any person
skilled in the art to practice the invention, including making and
using any devices or systems and performing any incorporated
methods. The patentable scope of the invention is defined by the
claims, and can include other examples that occur to those skilled
in the art. Such other examples are intended to be within the scope
of the claims if they have structural elements that do not differ
from the literal language of the claims, or if they include
equivalent structural elements with insubstantial differences from
the literal languages of the claims.
* * * * *