U.S. patent application number 13/652275 was filed with the patent office on 2013-02-14 for combustor liner cooling system.
This patent application is currently assigned to GENERAL ELECTRIC COMPANY. The applicant listed for this patent is GENERAL ELECTRIC COMPANY. Invention is credited to Ronald James Chila.
Application Number | 20130036742 13/652275 |
Document ID | / |
Family ID | 42232685 |
Filed Date | 2013-02-14 |
United States Patent
Application |
20130036742 |
Kind Code |
A1 |
Chila; Ronald James |
February 14, 2013 |
COMBUSTOR LINER COOLING SYSTEM
Abstract
A turbine engine with a combustor that includes a hollow wall
about a combustor liner. The combustor liner includes an inner
surface facing inwardly toward a combustion chamber. The turbine
engine includes a first air flow path in an upstream direction
through the hollow wall toward a head end of the combustor. The
first air flow path includes a plurality of bypass openings
extending through the combustor liner to the inner surface to
supply a first cooling film to a downstream end portion of the
combustor liner. The turbine engine further includes a second flow
path in a second direction opposite the upstream direction through
the hollow wall. The second flow path includes a plurality of film
holes extending through the combustor liner to the inner surface to
supply a second cooling film to the downstream end portion of the
combustor liner downstream of the first cooling film.
Inventors: |
Chila; Ronald James;
(Greenfield Center, NY) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
GENERAL ELECTRIC COMPANY; |
Schenectady |
NY |
US |
|
|
Assignee: |
GENERAL ELECTRIC COMPANY
Schenectady
NY
|
Family ID: |
42232685 |
Appl. No.: |
13/652275 |
Filed: |
October 15, 2012 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
12401530 |
Mar 10, 2009 |
8307657 |
|
|
13652275 |
|
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Current U.S.
Class: |
60/754 |
Current CPC
Class: |
F23R 2900/03042
20130101; F23R 2900/03044 20130101; F23R 3/06 20130101; Y02T 50/60
20130101; Y02T 50/675 20130101 |
Class at
Publication: |
60/754 |
International
Class: |
F23R 3/26 20060101
F23R003/26; F02C 7/00 20060101 F02C007/00 |
Claims
1. A system comprising: a turbine engine comprising: a combustor
comprising a hollow wall having a sleeve disposed about a combustor
liner, wherein the combustor liner comprises an inner surface
facing inwardly toward a combustion chamber; a first air flow path
in an upstream direction through the hollow wall toward a head end
of the combustor, wherein the first air flow path comprises a
plurality of bypass openings extending through the combustor liner
to the inner surface, wherein the plurality of bypass openings is
configured to supply a first cooling film to a downstream end
portion of the combustor liner; and a second air flow path in a
downstream direction opposite the upstream direction through the
hollow wall, wherein the second flow path comprises a plurality of
film holes extending through the combustor liner to the inner
surface, and the plurality of film holes is configured to supply a
second cooling film to the downstream end portion of the combustor
liner downstream of the first cooling film.
2. The system of claim 1, wherein the plurality of bypass openings
extending through the combustor liner is disposed upstream of the
plurality of film holes extending through the combustor liner.
3. The system of claim 2, wherein the plurality of bypass openings
extend through the downstream end portion of the combustor liner,
and an axial length of the downstream end portion is less than or
equal to approximately 35 percent of a total axial length of the
combustor liner.
4. The system of claim 1, wherein the second air flow path is
defined by a passage formed by an axial cooling channel on the
downstream end portion of the combustor liner and an inner surface
of a wrapper coaxially disposed generally about the downstream end
portion.
5. The system of claim 4, wherein the wrapper comprises one or more
radial openings configured to supply a portion of an air flow along
the first air flow path into the axial cooling channel.
6. The system of claim 5, wherein the plurality of film holes
extends radially through the axial cooling channel downstream of
the one or more radial openings, and the plurality of film holes is
configured to direct the portion of the air flow to supply the
second cooling film to the downstream end portion of the combustor
liner.
7. The system of claim 1, wherein the first air flow path is at
least partially defined by a first passage between a transition
piece and a transition sleeve that surrounds the transition
piece.
8. The system of claim 7, wherein the first passage is fluidly
coupled to a second passage between the combustor liner and the
transition sleeve, wherein the second passage extends in the
upstream direction from the first passage, the first passage
comprises a first plurality of inlets to receive a first portion of
air that flows through the first passage and the second passage in
the upstream direction, and the second passage comprises a second
plurality of inlets to receive a second portion of air that flows
through the second passage in the upstream direction.
9. The system of claim 8, wherein the turbine engine comprises one
or more fuel nozzles, wherein the fuel nozzles are configured to
receive the first portion of air and the second portion of air
flowing in the upstream direction through the second passage and to
mix the first portion of air and the second portion of air with a
fuel, and wherein the fuel nozzles are configured to output a
resulting air-fuel mixture into the combustion chamber surrounded
by the combustor liner for combustion.
10. A system comprising: a turbine combustor liner comprising: a
plurality of axial cooling channels arranged circumferentially
about a downstream end portion relative to a downstream direction
of combustion along a longitudinal axis of the turbine combustor
liner; an inner surface facing inwardly toward a combustion
chamber; and a plurality of bypass openings arranged
circumferentially about the downstream end portion upstream of the
plurality of axial cooling channels, wherein the plurality of
bypass openings is configured to supply a first cooling film to the
inner surface of the combustor liner, and each of the plurality of
axial cooling channels comprises a plurality of film holes
configured to supply a second cooling film to the inner surface of
the combustor liner at the downstream end portion.
11. The system of claim 10, wherein an interior of the turbine
combustor liner has a combustion path with a downstream direction
of flow of combustion gases, an exterior of the turbine combustor
liner has a first air path with an upstream direction of flow
opposite to the downstream direction, and the exterior of the
turbine combustor liner has the plurality of cooling channels with
a second air path in the downstream direction.
12. The system of claim 11, comprising a first flow sleeve disposed
about the turbine combustor liner to define a first hollow wall,
and a second flow sleeve disposed about a transition piece to
define a second hollow wall, wherein the first and second hollow
walls are coupled to one another at the downstream end portion, the
first and second hollow walls define the first air path with the
upstream direction, and the second air path in the downstream
direction is disposed radially between the plurality of cooling
channels and the transition piece.
13. The system of claim 10, wherein the plurality of cooling
channels is defined by alternating axial grooves and axial
protrusions about a circumference of the turbine combustor liner,
and the plurality of film holes extend radially through the axial
grooves into an interior of the turbine combustor liner.
14. The system of claim 10, wherein the plurality of bypass
openings extend radially through the combustor liner at an angle of
approximately 90 degrees relative to the inner surface.
15. The system of claim 10, wherein the plurality of bypass
openings extend radially through the combustor liner at an angle
between approximately 30 to 60 degrees.
16. The system of claim 10, wherein at least one bypass opening of
the plurality of bypass openings has a geometry that converges or
diverges through the combustor liner into the interior of the
turbine combustor liner.
17. The system of claim 10, wherein each bypass opening of the
plurality of bypass openings is disposed in an upstream direction
of combustion along the longitudinal axis from an axial cooling
channel of the plurality of axial cooling channels.
18. The system of claim 10, wherein the bypass openings are
arranged in two or more axially spaced sets, wherein each set is
disposed circumferentially about the downstream end portion.
19. The system of claim 10, wherein an axial length of the
downstream end portion is less than or equal to approximately 35
percent of a total axial length of the turbine combustor liner, an
axial channel length of each of the plurality of cooling channels
is less than or equal to the axial length of the downstream end
portion, and the cooling channels have a depth of approximately
0.05 to 0.30 inches and a width of approximately 0.25 to 1.0
inches.
20. A system comprising: a turbine engine comprising: a combustor
comprising: a flow sleeve; and a combustor liner surrounded by the
flow sleeve and defining a flow path therebetween configured to
receive an air flow in a first direction toward a head end chamber,
wherein the combustor liner comprises: an inner surface facing
inwardly toward a combustion chamber; a plurality of axial cooling
channels arranged circumferentially about a downstream end portion
of the combustor liner; a plurality of bypass holes arranged in two
or more axially spaced sets at the downstream end portion, wherein
each bypass opening is disposed upstream in the first direction
relative to an axial cooling channel of the plurality of axial
cooling channels, and each bypass opening is configured to supply a
first cooling film to the inner surface of the combustor liner at
the downstream end portion; one or more fuel nozzles disposed in
the head end chamber of the combustor; and wherein each of the
plurality of axial cooling channels comprises a plurality of film
holes extending through the combustor liner to the inner surface,
each of the plurality of axial cooling channels is configured to
receive a portion of the air flow from the flow path, to direct a
first portion of the received air along an axial length of the
axial cooling channel in a second direction away from the head end
chamber, and to direct a second portion of the received air through
the plurality of film holes to supply a second cooling film to the
inner surface of the combustor liner at the downstream end portion.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application is a continuation of U.S. patent
application Ser. No. 12/401,530, entitled "COMBUSTOR LINER COOLING
SYSTEM", filed Mar. 10, 2009, which is herein incorporated by
reference in its entirety.
BACKGROUND OF THE INVENTION
[0002] The subject matter disclosed herein relates to gas turbine
engines and, more specifically, to a system for cooling a combustor
liner used in a combustor of a gas turbine engine.
[0003] Gas turbine engines typically include a combustor having a
combustor liner defining a combustion chamber. Within the
combustion chamber, a mixture of compressed air and fuel is
combusted to produce hot combustion gases. The combustion gases may
flow through the combustion chamber to one or more turbine stages
to generate power for driving a load and/or a compressor.
Typically, the combustion process heats the combustor liner due to
the hot combustion gases. Unfortunately, existing cooling systems
may not adequately cool the combustor liner in all conditions.
BRIEF DESCRIPTION OF THE INVENTION
[0004] Certain embodiments commensurate in scope with the
originally claimed invention are summarized below. These
embodiments are not intended to limit the scope of the claimed
invention, but rather these embodiments are intended only to
provide a brief summary of possible forms of the invention. Indeed,
the invention may encompass a variety of forms that may be similar
to or different from the embodiments set forth below.
[0005] In one embodiment, a system includes a turbine engine. The
turbine engine includes a combustor that includes a hollow wall
about a combustor liner. The combustor liner includes an inner
surface facing inwardly toward a combustion chamber. The turbine
engine also includes a first air flow path in an upstream direction
through the hollow wall toward a head end of the combustor. The
first air flow path includes a plurality of bypass openings
extending through the combustor liner to the inner surface to
supply a first cooling film to a downstream end portion of the
combustor liner. The turbine engine further includes a second flow
path in a second direction that is opposite the upstream direction
through the hollow wall. The second flow path may include a
plurality of film holes extending through the combustor liner to
the inner surface to supply a second cooling film to the downstream
end portion of the combustor liner downstream of the first cooling
film.
[0006] In another embodiment, a system includes a turbine combustor
liner. The turbine combustor liner includes a plurality of axial
cooling channels arranged circumferentially about a downstream end
portion of the turbine combustor liner, the downstream end portion
being relative to a downstream direction of combustion along a
longitudinal axis of the turbine combustor liner. The turbine
combustor includes an inner surface facing inwardly toward a
combustion chamber, and a plurality of bypass openings arranged
circumferentially about the downstream end portion upstream of the
plurality of axial cooling channels. The plurality of bypass
openings is configured to supply a first cooling film to the inner
surface of the combustion liner, and each of the plurality of axial
cooling channels includes one or more film holes configured to
supply a second cooling film to the inner surface of the combustor
liner at the downstream end portion.
[0007] In yet another embodiment, a system includes a turbine
engine having a combustor with a flow sleeve and a combustor liner
surrounded by the flow sleeve. The combustor liner and flow sleeve
define a flow path therebetween configured to receive an air flow
in a first direction towards a head end chamber. The combustor
liner includes an inner surface facing inwardly toward a combustion
chamber, a plurality of axial cooling channels arranged
circumferentially about a downstream end portion of the combustor
liner, a plurality of bypass holes arranged in two or more axially
spaced sets at the downstream end portion, and one or more fuel
nozzles disposed in the head end chamber of the combustor. Each
bypass opening is disposed upstream in the first direction relative
to an axial cooling channel of the plurality of axial cooling
channels. Each bypass opening is configured to supply a first
cooling film to the inner surface of the combustor liner at the
downstream end portion. Each of the plurality of axial cooling
channels includes a plurality of film holes extending through the
combustor liner to the inner surface. Each of the plurality of
axial cooling channels is configured to receive a portion of the
air flow from the flow path, to direct a first portion of the
received air along an axial length of the axial cooling channel in
a second direction away from the head end chamber, and to direct a
second portion of the received air through the plurality of film
holes to supply a second cooling film to the inner surface of the
combustor liner at the downstream end portion.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] These and other features, aspects, and advantages of the
present invention will become better understood when the following
detailed description is read with reference to the accompanying
drawings in which like characters represent like parts throughout
the drawings, wherein:
[0009] FIG. 1 is a block diagram of a turbine system having a
combustor liner having cooling channels for enhanced cooling, in
accordance with an embodiment of the present technique;
[0010] FIG. 2 is a cutaway side view of the turbine system, as
shown in FIG. 1, in accordance with an embodiment of the present
technique;
[0011] FIG. 3 is a cutaway side view of the combustor, as shown in
FIG. 1, having a combustor liner with cooling channels at a
downstream end portion, in accordance with an embodiment of the
present technique;
[0012] FIG. 4 is an exploded perspective view of certain components
of the combustor, as shown in FIG. 3, in accordance with an
embodiment of the present technique.
[0013] FIG. 5 is a partial perspective view of a portion of the
cooling channels on the downstream end portion of the combustor
liner, taken within line 5-5 as shown in FIG. 4, in accordance with
an embodiment of the present technique;
[0014] FIG. 6 is a partial cross-sectional side view of the
downstream end portion of the combustor liner, taken within line
6-6 as shown in FIG. 3, in accordance with an embodiment of the
present technique;
[0015] FIGS. 7A and 7B are partial cross-sectional end views of a
cooling channel within the downstream end portion of the combustor
liner, taken along line 7-7 as shown in FIG. 6, in accordance with
embodiments of the present technique;
[0016] FIGS. 8A-8D are partial cross-sectional side views
illustrating configurations of holes within a cooling channel of
the downstream end portion of the combustor liner, taken within
line 8-8 as shown in FIG. 6, in accordance with embodiments of the
present technique; and
[0017] FIG. 9 is a partial cross-sectional side view of the
downstream end portion of the combustor liner, taken within line
6-6 as shown in FIG. 3, in accordance with a further embodiment of
the present technique.
DETAILED DESCRIPTION OF THE INVENTION
[0018] One or more specific embodiments of the present invention
will be described below. In an effort to provide a concise
description of these embodiments, all features of an actual
implementation may not be described in the specification. It should
be appreciated that in the development of any such actual
implementation, as in any engineering or design project, numerous
implementation-specific decisions must be made to achieve the
developers' specific goals, such as compliance with system-related
and business-related constraints, which may vary from one
implementation to another. Moreover, it should be appreciated that
such a development effort might be complex and time consuming, but
would nevertheless be a routine undertaking of design, fabrication,
and manufacture for those of ordinary skill having the benefit of
this disclosure.
[0019] When introducing elements of various embodiments of the
present invention, the articles "a," "an," "the," and "said" are
intended to mean that there are one or more of the elements. The
terms "comprising," "including," and "having" are intended to be
inclusive and mean that there may be additional elements other than
the listed elements. Any examples of operating parameters and/or
environmental conditions are not exclusive of other
parameters/conditions of the disclosed embodiments. Additionally,
it should be understood that references to "one embodiment" or "an
embodiment" of the present invention are not intended to be
interpreted as excluding the existence of additional embodiments
that also incorporate the recited features.
[0020] Before continuing, several terms used extensively throughout
the present disclosure will be first defined in order to provide a
better understanding of the claimed subject matter. As used herein,
the terms "upstream" and "downstream," when discussed in
conjunction with a combustor liner, shall be understood to mean the
proximal end of the combustor liner and the distal end of the
combustor liner, respectively, with respect to the fuel nozzles.
That is, unless otherwise indicated, the terms "upstream" and
"downstream" are generally used with respect to the flow of
combustion gases inside the combustor liner. For example, a
"downstream" direction refers to the direction in which a fuel-air
mixture combusts and flows from the fuel nozzles towards a turbine
and an "upstream" direction refers to a direction opposite the
downstream direction, as defined above. Additionally, the term
"downstream end portion," "coupling portion," or the like, shall be
understood to refer to an aft-most (downstream most) portion of the
combustor liner. As will be discussed further below, the axial
length of the downstream end portion of the combustor liner, in
certain embodiments, may be as much as approximately 20 percent the
total axial length of the combustor liner. The downstream end
portion (or coupling portion), in some embodiments, may also be
understood to be the portion of the liner that is generally
configured to couple to a downstream transition piece of the
combustor, generally in a telescoping, concentric, or coaxial
overlapping annular relationship. Further, where the term "liner"
appears alone, it should be understood that this term is generally
synonymous with "combustor liner."
[0021] Keeping in mind the above-defined terms, the present
disclosure is generally directed towards a combustor liner capable
of providing more effective cooling during the operation of a
turbine engine. In one embodiment, the liner has a downstream end
portion that includes a plurality of channels (also referred to
herein as "cooling channels") arranged circumferentially about the
outer surface of the downstream end portion. The channels may
define a flow path that is parallel to the longitudinal axis of the
liner. Further, each channel may each include one or more openings
that fluidly couple the channel to the combustion chamber. While
certain embodiments will refer to these openings as "film holes,"
it should be understood that the openings may include holes, slots,
or a combination of holes and slots, and may be formed using any
suitable technique, such as laser drilling, for example.
[0022] In operation, an annular wrapper having a plurality of
openings extending radially therethrough may be coupled to the
liner generally about the downstream end portion. The inner surface
of the wrapper and the cooling channels on the downstream end
portion may define one or more passages through which an air flow
may be supplied via one or more openings on the wrapper. The air
flow may be a portion of the compressed air supplied to the
combustor for combustion of fuel. As the air flows through a
cooling channel, heat may be transferred away from the combustor
liner, particularly the downstream end portion of the liner, via
forced air convention. Additionally, a portion of the air flow
within the cooling channel may flow through the film holes and into
the combustion chamber to provide a film of cooling air which
insulates the liner, particularly the downstream end portion of the
liner, from the relatively hotter combustion gases within the
combustor, thereby cooling the liner via film cooling.
[0023] In some embodiments, the liner may further include a
plurality of "bypass openings" in addition to the film holes
discussed above. The bypass openings may be located upstream from
the cooling channels and may extend radially through the combustor
liner. The bypass openings may provide a direct flow path into the
combustion chamber (e.g., interior of the combustor liner) that
bypasses the cooling channels discussed above. Air may flow into
the combustion chamber along this path, thus providing an
additional cooling film along the interior of the surface of the
combustor liner upstream from the cooling film provided via the
film holes, thereby further insulating the liner from combustion
gases within the liner. In this manner, embodiments of the present
invention may provide for enhanced heat transfer using forced
convection and film cooling principles simultaneously.
Advantageously, this may improve overall turbine performance and
increase the life of the combustor and/or combustor liner.
[0024] Turning now to the drawings and referring first to FIG. 1, a
block diagram of an embodiment of a turbine system 10 is
illustrated. As discussed in detail below, the disclosed turbine
system 10 may employ a combustor liner having cooling channels
formed on a downstream end portion of the liner. The cooling
channels may include film holes that provide for improved cooling
of the downstream end portion, as will be discussed further below.
The turbine system 10 may use liquid or gas fuel, such as natural
gas and/or a hydrogen rich synthetic gas, to run the turbine system
10. As depicted, a plurality of fuel nozzles 12 intakes a fuel
supply 14, mixes the fuel with air, and distributes the air-fuel
mixture into a combustor 16. The air-fuel mixture combusts in a
chamber within combustor 16, thereby creating hot pressurized
exhaust gases. The combustor 16 directs the exhaust gases through a
turbine 18 toward an exhaust outlet 20. As the exhaust gases pass
through the turbine 18, the gases force one or more turbine blades
to rotate a shaft 22 along an axis of the system 10. As
illustrated, the shaft 22 may be connected to various components of
turbine system 10, including a compressor 24. The compressor 24
also includes blades that may be coupled to the shaft 22. As the
shaft 22 rotates, the blades within the compressor 24 also rotate,
thereby compressing air from an air intake 26 through the
compressor 24 and into the fuel nozzles 12 and/or combustor 16. The
shaft 22 may also be connected to a load 28, which may be a vehicle
or a stationary load, such as an electrical generator in a power
plant or a propeller on an aircraft, for example. As will be
understood, the load 28 may include any suitable device that
capable of being powered by the rotational output of turbine system
10.
[0025] FIG. 2 illustrates a cutaway side view of an embodiment of
the turbine system 10 schematically depicted in FIG. 1. The turbine
system 10 includes one or more fuel nozzles 12 located inside one
or more combustors 16. The combustor 16 may include one or more
combustor liners disposed within one or more respective flow
sleeves. As mentioned above, the combustor liner (or liners) may
include a plurality of cooling channels formed on a downstream end
portion of the liner. Each cooling channel may also include
openings, such as film holes, which fluidly couple the cooling
channel to the combustion chamber defined by the liner.
[0026] In operation, air enters the turbine system 10 through the
air intake 26 and may be pressurized in the compressor 24. The
compressed air may then be mixed with gas for combustion within
combustor 16. For example, the fuel nozzles 12 may inject a
fuel-air mixture into the combustor 16 in a suitable ratio for
optimal combustion, emissions, fuel consumption, and power output.
The combustion generates hot pressurized exhaust gases, which then
drive one or more blades 17 within the turbine 18 to rotate the
shaft 22 and, thus, the compressor 24 and the load 28. The rotation
of the turbine blades 17 causes a rotation of shaft the 22, thereby
causing blades 19 within the compressor 22 to draw in and
pressurize the air received by the intake 26.
[0027] As will be discussed in further detail below, each of the
cooling channels on the downstream end portion of the combustor
liner may receive a portion of the air supplied to the combustor 16
through the air intake 26. In one embodiment, the total air
supplied to the cooling channels may make up approximately 2% of
the total air supplied to the combustor 16 via compressor 24 and
intake 26. As the compressor-supplied air (which is generally
substantially cooler relative to the combustion gases within the
combustor 16) flows through the cooling channels, heat is
transferred away from the downstream end portion of the liner
(e.g., via forced convection cooling). Further, a portion of the
airflow within each cooling channels may flow through the film
holes and form a cooling film along a portion of the inner surface
of the liner. The cooling film insulates the liner from the
relatively hot combustion gases flowing within the combustor 16.
Thus, in operation, cooling of the liner, particularly the
downstream end portion of the liner, is enhanced by utilizing both
forced convection and film cooling techniques.
[0028] Continuing now to FIG. 3, a more detailed cutaway side view
of an embodiment of the combustor 16, as shown FIG. 2, is
illustrated. As will be appreciated, the combustor 16 is generally
fluidly coupled to the compressor 24 and the turbine 18. The
compressor 24 may include a diffuser 29 and a discharge plenum 31
that are coupled to each other in fluid communication as to
facilitate the channeling of air downstream to the combustor 16. In
the illustrated embodiment, the combustor 16 includes a cover plate
30 at the upstream head end of the combustor 16. The cover plate 30
may at least partially support the fuel nozzles 12 and provide a
path through which air and fuel are directed to the fuel nozzles
12.
[0029] The illustrated combustor 16 comprises a hollow annular wall
configured to facilitate cooling air flow. For example, the
combustor 16 includes a combustor liner 34 disposed within a flow
sleeve 32. The arrangement of the liner 34 and the flow sleeve 32,
as shown in FIG. 3, is generally concentric and may define an
annular passage 36. In certain embodiments, the flow sleeve 32 and
the liner 34 may define a first or upstream hollow annular wall of
the combustor 16. The interior of the liner 34 may define a
substantially cylindrical or annular combustion chamber 38. The
flow sleeve 32 may include a plurality of inlets 40, which provide
a flow path for at least a portion of the air from the compressor
24 into the annular passage 36. In other words, the flow sleeve 32
may be perforated with a pattern of openings to define a perforated
annular wall.
[0030] Downstream from the liner 34 and the flow sleeve 32 (e.g. in
the direction 39), a second flow sleeve 42, which may be referred
to as an "impingement sleeve," may be coupled to the flow sleeve
32. Thus, the direction 39 may represent a downstream direction
with respect to the flow of combustion gases away from the fuel
nozzles 12 inside the liner 34. As used herein, the terms
"upstream" and "downstream," when discussed in conjunction with a
combustor liner, shall be understood to mean the proximal end of
the combustor liner and the distal end of the combustor liner 34,
respectively, with respect to the fuel nozzles 12. That is, unless
otherwise indicated, the terms "upstream" and "downstream" are
generally used with respect to the flow of combustion gases inside
the combustor liner. For example, a "downstream" direction refers
to the direction 39 in which a fuel-air mixture combusts and flows
from the fuel nozzles 12 towards the turbine 18, and an "upstream"
direction refers to a direction opposite the downstream direction,
as defined above.
[0031] In the present embodiment, the flow sleeve 32 may include a
mounting flange 44 configured to receive a portion of the
impingement sleeve 42. A transition piece 46 (which may be referred
to as a "transition duct") may be disposed within the impingement
sleeve 42. A concentric arrangement of the impingement sleeve 42
and the transition piece 46 may define an annular passage 47. As
shown, the annular passage 47 is fluidly coupled to the annular
passage 36. In certain embodiments, the sleeve 42 and the
transition piece 46 may define a second or downstream hollow
annular wall of the combustor 16. Thus, together, the elements 32,
34, 42, and 46 define a hollow annular wall (e.g., upstream and
downstream portions) configured to facilitate air flow to the fuel
nozzles 12, while also cooling the combustor 16 due to the heat
generated from combustion.
[0032] The impingement sleeve 42 may include a plurality of inlets
48 (e.g., perforated annular wall), which may provide a flow path
for at least a portion of the air from the compressor 24 into the
annular passage 47. An interior cavity 50 of the transition piece
46 generally provides a path by which combustion gases from the
combustion chamber 38 may be directed thru a turbine nozzle 60 and
into the turbine 18. In the depicted embodiment, the transition
piece 46 may be coupled to the downstream end of the liner 34 (with
respect to direction 39), generally about a downstream end portion
52 (coupling portion), as discussed above. An annular wrapper 54
and a seal may be disposed between the downstream end portion 52
and the transition piece 46. The seal may secure the outer surface
of the wrapper 54 to inner surface of the transition piece 46.
Further, as mentioned above, the inner surface of the wrapper 54
and the cooling channels on the downstream end portion may define
passages that receive a portion of the air flow from the annular
passage 47.
[0033] As discussed above, the turbine system 10, in operation, may
intake air through the air intake 26. The compressor 24, which is
driven by the shaft 22, rotates and compresses the air. The
compressed air is discharged into the diffuser 29, as indicated by
the arrows shown in FIG. 3. The majority of the compressed air is
further discharged from the compressor 24, by way of the diffuser
29, through a plenum 31 into the combustor 16. Though not shown in
detail here, a smaller portion of the compressed air may be
channeled downstream for cooling of other components of the turbine
engine 10. A portion of the compressed air within the plenum 31 may
enter the annular passage 47 by way of the inlets 48. The air in
the annular passage 47 is then channeled upstream (e.g., in the
direction of fuel nozzles 12) towards the annular passage 36, such
that the air flows over the downstream end portion 52 of the liner
34. That is, a flow path in the upstream direction (relative to
direction 39) is defined by the annular passages 36 (formed by
sleeve 32 and liner 34) and 47 (formed by sleeve 42 and transition
piece 46). A portion of the air flowing in the upstream direction
is diverted into the cooling channels on the downstream end portion
of the liner 34 to facilitate cooling. In one embodiment, a
plurality of inlets on the wrapper 54 may provide a flow path into
the cooling channels. As mentioned above, air flowing through the
channel may cool the liner 34 via forced convection cooling.
Additionally, a portion of the airflow within the channel may be
diverted through one or more film holes within the channel and into
the combustion chamber 38, as indicated by the air flow 53. The air
flow 53 may form a cooling film that insulates the downstream end
portion 52 of the liner 34 from the hot combustion gases within the
chamber 38.
[0034] The portion of the air flow that is not discharged into the
cooling channel continues to flow upstream into the annular passage
36 toward the cover plate 30 and fuel nozzles 12. Accordingly, the
annular passage 36 may receive air from the annular passage 47 and
the inlets 40. As shown in FIG. 3, a portion of the air flow within
the annular passage 36 may be directed into one or more bypass
openings 41 on the liner 34. The bypass openings 41 extend radially
through the liner 34 and provide a direct flow path into the
combustion chamber 38 that bypasses the cooling channels on the
downstream end portion 52. The air 43 that flows into the
combustion chamber 38 through the bypass openings 41 may provide an
additional cooling film along the inner surface of the liner 34
upstream from the cooling film provide via film holes within the
cooling channels, thus providing additional insulation for the
liner 34. The remaining air flowing into the annular passage 36 is
then channeled upstream towards the fuel nozzles 12, wherein the
air is mixed with fuel 14 and ignited within the combustion chamber
38. The resulting combustion gases are channeled from the chamber
38 into the transition piece cavity 50 and through the turbine
nozzle 60 to the turbine 18.
[0035] FIG. 4 is an exploded perspective view showing some of the
above-discussed components of the combustor 16. Particularly, FIG.
4 is intended to provide a better understanding of the relationship
between the liner 34, the wrapper 54, and the transition piece 46.
As shown, the liner 34 may have a length of L1 when measured along
a longitudinal axis, referred to here by reference number 58. In
the illustrated embodiment, a radius R1 of the upstream end of the
liner 34 may be greater than a radius R2 of the downstream end of
the liner 34. In other embodiments, however, the radii R1 and R2
may be equal or the radius R2 may be greater than the radius R1.
The liner 34 includes the downstream end portion 52. As discussed
above, the downstream end portion 52 is a portion of the liner
having an axial length L2 which, when measured from the downstream
(aft-most) end of the liner 34, is less than the total length L1 of
the liner 34. In one embodiment, the length L2 of the downstream
end portion 52 may be approximately 10-20 percent of the total
length L1 of the liner. However, it should be appreciated that in
other embodiments, depending on implementation specific goals, the
length L2 could be greater than 20 percent or less than 10 percent
of L1. For example, in other embodiments, the longitudinal length
L2 of the downstream end portion 52 may be at least less than
approximately 5, 10, 15, 20, 25, 30, or 35 percent of the total
length L1.
[0036] The wrapper 54 is configured to mate with the liner 34
generally about the downstream end portion 52 in a telescoping,
coaxial, or concentric overlapping relationship. The transition
piece 46 is coupled to the liner 34 generally about the downstream
end portion 52 and the wrapper 54. A sealing ring 66 may be
disposed between the wrapper 54 and the transition piece 46 to
facilitate the coupling. As shown, the wrapper 54 may include a
plurality of inlets 68 generally near the upstream end of the
wrapper 54. In the illustrated embodiment, the inlets 68 are
depicted as a plurality of openings disposed circumferentially
(relative to the axis 58) about the upstream end of the wrapper 54
and also extending radially therethrough. The openings defined by
the inlets 68 may include holes, slots, or a combination of holes
and slots, for example. An inner surface 55 of the wrapper 54 and
the cooling channels 56 on the downstream end portion 52 may form
passages to receive an air flow provided via the inlets 68. By way
of example, in one embodiment, each inlet 68 may supply an air flow
(e.g., divert a portion of the air flowing upstream towards the
fuel nozzles 12 through annular passages 36 and 47) to a respective
cooling channel 56 on the downstream end portion 52. As the air
(which is substantially cooler relative to the temperature of the
combustion gases within the combustion chamber 38) flows into and
through the channels 56, heat is transferred away from the liner
34, thus cooling the liner 34. Additionally, as discussed above,
one or more of the channels 56 may include film holes fluidly
coupling the channel 56 to the combustion chamber 38. A portion of
the air flow within the channel 56 may be diverted low through the
film holes to provide a cooling film that insulates the inner
surface of the liner 34 from the combustion gases in the chamber
38. The liner 34 also includes the bypass openings 41 which, as
discussed above, may provide an additional cooling film along the
inner surface of the liner 34, thus providing additional insulation
for the liner 34.
[0037] FIG. 5 is a partial perspective view showing the cooling
channels 56 on the downstream end portion 52 of the liner 34 within
the circular region defined by the arcuate line 5-5, as shown in
FIG. 4. As shown in the depicted embodiment, a plurality of axial
cooling channels 56 are arranged circumferentially about the
downstream end portion 52 of the liner 34. The channels 56 may
define flow paths generally parallel to one another and the
longitudinal axis 58 of the liner 34. In one embodiment, the
channels 56 may be formed by removing a portion of the outer
surface of the downstream end portion 52, such that each cooling
channel 56 is a recessed groove between adjacent raised dividing
members 62. Thus, the cooling channels 56 may be defined by
alternating axial grooves and axial protrusions (e.g., 62) about a
circumference of the combustor liner 34. As will be appreciated,
the channels 56 may be formed using any suitable technique,
including milling, casting, molding, or laser etching/cutting, for
example. The cooling channels 56, in one embodiment, may have an
axial length (with respect to axis 58) that is substantially
equivalent to the axial length L2 of the downstream end portion 52,
as discussed above. In other embodiments, the cooling channels 56
may have an axial length that is less than L2. By way of example
only, the axial length of each cooling channel 56 may be at least
less than approximately 3, 4, 5, 6, 7, or 8 inches. In other
embodiments, however, the axial length of the cooling channels 56
may be less than 3 inches or greater than 8 inches. The cooling
channels may also have various depths and widths. In one
embodiment, the cooling channels may have a width of at least less
than approximately 0.25 inches, 0.5 inches, 0.75 inches, or 1 inch.
In other embodiments, the width may be less than 0.25 inches or
greater than 1 inch. Further, in one embodiment, the depth of the
cooling channels 56 may be at least less than approximately 0.05
inches, 0.10 inches, 0.15 inches, 0.20 inches, 0.25 inches, or 0.30
inches. In further embodiments, the depth of the cooling channels
56 may be less than 0.05 inches or greater than 0.30 inches.
[0038] The film holes 64 extend radially through the axial grooves
into an interior of the combustor liner 34. In certain embodiments,
the film holes 64 may be arranged in a group, as shown in FIG. 5,
at a particular axial position along each cooling channel 56. For
example, the film holes 64 may include between approximately 1 and
20 or 1 and 10 openings in a group, which may be disposed at an
axial position of approximately 20, 40, 60, or 80 percent of the
length L2 of the cooling channel 56 relative to the downstream end.
In some embodiments, the film holes 64 may be disposed at multiple
axial positions, equally or non-equally spaced relative to one
another, along the length L2 of the channels 56. In the present
figure, film holes 64 are shown in only one channel 56 for purposes
of simplicity. It should be appreciated that in an actual
implementation, similar arrangements of the illustrated film holes
64 may be provided in more than one cooling channel 56 on the
downstream end portion 52 (e.g., each cooling channel 56 may
include film holes 64 in one or more locations).
[0039] As discussed above, the film holes 64 fluidly couple the
channels 56 to the combustion chamber 38 and may provide an
insulating film of cooling air along the inner surface of the liner
34. In one embodiment, the film holes 64 may have a diameter of at
least less than approximately 0.01, 0.02, 0.03, 0.04, 0.05, 0.06,
0.07, 0.08, 0.09, or 0.10 inches. In other embodiments, the film
holes 64 may be less than 0.01 inches or greater than 0.10 inches.
FIG. 5 also shows the bypass openings 41 located upstream from the
cooling channels 56. As discussed above, the bypass openings 41 may
provide a flow of air directly into the combustion chamber 38
(e.g., bypassing the cooling channels 56), thus providing an
additional cooling film along the inner surface of the liner 34,
thereby further enhancing cooling of the liner 34. In one
embodiment, the bypass openings 41 may have dimensions similar to
the film holes 64, as discussed above. That is, the bypass openings
41, in one embodiment, may have a diameter of at least less than
approximately 0.01, 0.02, 0.03, 0.04, 0.05, 0.06, 0.07, 0.08, 0.09,
or 0.10 inches or, in other embodiments, less than 0.01 inches or
greater than 0.10 inches. As will be appreciated, the presently
illustrated embodiment is only intended by provide an example of a
particular implementation that utilizes both the film holes 64 and
bypass openings 41 to cool the liner 34 via film cooling and forced
convection cooling. In another embodiment, the liner 34 may include
only the film holes 64 and not the bypass openings 41.
[0040] Referring now to FIG. 6, a partial cross-sectional side view
of the combustor 16 within the circular region defined by the
arcuate line 6-6 in FIG. 3 is illustrated. Particularly, FIG. 6
shows in more detail the air flow into the cooling channels 56 on
downstream end portion 52 of the liner 34. Compressed air
discharged by the compressor 24 may be received in the annular
passage 47 (defined by the impingement sleeve 42 and the transition
piece 46) through the inlets 48. In the present embodiment, the
inlets 48 are circular-shaped holes, although in other
implementations, the inlets 48 may be slots, or a combination of
holes and slots of other geometries. As the air 72 within the
annular passage 47 is channeled upstream relative to the direction
of the combustion gas flow (e.g., direction 39), the majority of
the air 72 is discharged into the annular passage 36 (defined by
the flow sleeve 32 and the liner 34). As discussed above, the flow
sleeve 32 may include the mounting flange 44 at a downstream end 74
configured to receive a member 76 extending radially outward from
the upstream end 78 of the impingement sleeve 42, thereby fluidly
coupling the flow sleeve 32 and impingement sleeve 42. In addition
to receiving the air flow 72 from the annular passage 47, the
annular passage 36 also receives a portion 80 of the compressed air
from the plenum 31 by way of the inlets 40. That is, the airflow
within the annular passage 36 may include air 72 discharged from
the annular passage 47 and air 80 flowing through the inlets 40.
Thus, a flow path that is directed upstream (with respect to the
direction 39) is defined by the annular passages 36 and 47.
Additionally, it should be understood that like the inlets 48 on
the impingement sleeve 42, the inlets 40 may also include holes,
slots, or a combination thereof, of various shapes.
[0041] While a majority of the air 72 flowing through the annular
passage 47 is discharged into the annular passage 36, a portion of
the air flow, shown here by the reference number 84, may be
directed into the cooling channels 56 on the downstream end portion
52 by way of a flow path F provided by the plurality of inlets 68
on the wrapper 54. The flow path F may define an air flow through
the cooling channels 56. As shown, the flow path F is directed
downstream with respect to direction 39, and is opposite of the
flow path through the annular passages 36 and 47. Though only one
cooling channel 56 is shown in the cross-sectional view of FIG. 6,
it should be understood that a similar air flow scheme may be
applied to each of the cooling channels 56 on the downstream end
portion 52. In one embodiment, the total air flow directed into and
through the cooling channels 56 about the downstream end portion 52
may represent at least less than approximately 1, 2, 3, 4, 5, 6, 7,
8, 9, or 10 percent of the total compressed air supplied to the
combustor 16. In other embodiments, the total air directed into the
cooling channels 56 may be more than 10 percent of the total
compressed air supplied to the combustor 16.
[0042] As discussed above, the air 84 that flows into the depicted
cooling channel 56 is generally substantially cooler relative to
the temperature of the combustion gases within the combustion
chamber 38. Thus, as the air 84 flows through the cooling channels
56 along the flow path F, heat may be transferred away from the
combustor liner 34, particularly the downstream end portion 52 of
the liner. By way of example, the mechanism employed in cooling the
liner 34 may be forced convective heat transfer resulting from the
contact between the cooling air 84 and the outer surface of
downstream end portion 52, which may include the grooves and
dividing members 62 defining the channels 56, as discussed above
with reference to FIG. 5. The flow path F may continue along the
axial length of the cooling channel 56, wherein the cooling air 84
exits the cooling channel 56 at a downstream end (not shown),
thereby discharging into the transition piece cavity 50, whereby
the cooling air 84 is directed towards combustion gases flowing
downstream (away from the fuel nozzles 12) through the transition
piece cavity 50.
[0043] As shown in the present embodiment, a portion 53 of the
cooling air 84 may flow through the film holes 64 within the
cooling channel 56 and into the combustion chamber 38. The air 53
may provide a cooling film 86 that insulates the liner 34 from the
combustion gases within the chamber 38, as discussed above. The
depicted cooling film 86 may also include the air flow 43, which
may be provided through the bypass openings 41 on the liner 34.
Thus, air directed through the film holes 64 and the bypass
openings 41 may both contribute to the formation of the cooling
film 86.
[0044] Referring now to FIGS. 7A and 7B, cross-sectional end views
of the cooling channel 56 with respect to the cut line 7-7 of FIG.
6 are illustrated in accordance with embodiments of the invention.
Referring first to FIG. 7A, the transition piece 46, seal 66,
wrapper 54, and downstream end portion 52 of the liner 34 are shown
in the arrangement described above. As discussed, the cooling
channels 56 may be formed by removing a portion of the liner 34 to
define a groove between dividing members 62. In the illustrated
embodiment, the dividing members 62 may have a height 94 of
approximately at least less than approximately 0.05 inches, 0.10
inches, 0.15 inches, 0.20 inches, 0.25 inches, or 0.30 inches,
which may corresponding to the depth of the cooling channel 56, as
mentioned above. In further embodiments, the height 94 of the
dividing members 62 may be less than 0.05 inches or greater than
0.30 inches. Additionally, the width 90 of the cooling channel 56
may be defined as a circumferential distance between the sidewalls
92 of two adjacent dividing members 62. As discussed above, in one
embodiment, the width 90 (e.g., circumferential width) of each
cooling channel 56 may be at least less than approximately 0.25
inches, 0.5 inches, 0.75 inches, or 1 inch. In other embodiments,
the width 90 may be less than 0.25 inches or greater than 1
inch.
[0045] In the depicted embodiment, the cooling channel 56 may have
a substantially flat and/or smooth surface 95. For example, the
surface 95 may be flat in the axial and/or circumferential
directions, or the surface 95 may have a slight curvature in the
circumferential direction due to the annular shape of the liner 34.
By further example, the surface 95 may be substantially or entirely
free of protrusions, recesses, or surface texture except for the
film holes 64. As cooling air (e.g., air 84) flows through the
channel 56 in the downstream direction 39 (i.e., perpendicular to
the page) and contacts the surface 95 and sidewalls 92, heat may be
transferred away from the liner 34, particularly the downstream end
portion 52 of the liner 34, via forced convection cooling.
Additionally, as mentioned above, a portion 53 of the cooling air
84 may flow through one or more film holes 64 that extend radially
through the channel 56 and fluidly couple the channel 56 to the
combustion chamber 38. As the air 53 flows through the film holes
64 and into the chamber 38, a cooling film 86 is formed. As
discussed above, the cooling film 86 may insulate the liner 34 from
the hot combustion gases within the chamber 38.
[0046] While the present view depicted by FIG. 7A shows three film
holes 64 distributed circumferentially across the width 90 of the
cooling channel 56, it should be understood that this is meant to
provide merely one example of how the film holes 64 may be arranged
within the channel 56. Indeed, any other suitable arrangement of
film holes 64 may be employed. For instance, a plurality of film
holes 64 may be arranged in both circumferential and axial
directions within the cooling channel 56. Further, as will be
discussed further below with respect to FIG. 9, in some
embodiments, the film holes 64 may be arranged in a plurality of
groups axially spaced along the axial length of the cooling channel
56.
[0047] Referring to FIG. 7B, an alternate embodiment of the cooling
channel 56 is illustrated. In contrast to the flat and/or smooth
surface 95 shown in FIG. 7A, the surface 95 of the embodiment
depicted in FIG. 7B may include a plurality of surface features 96,
which may be discrete protrusions extending from the surface 95. By
way of example, the surface features may include fin-shaped
protrusions, cylindrical-shaped protrusions, ring-shaped
protrusions, chevron-shaped protrusions, raised portions between
cross-hatched grooves formed with in the cooling channel 56, or
some combination thereof, as well as any other type of suitable
geometric shape. It should be appreciated that the dimensions of
the surface features 96 may be selected to optimize cooling while
satisfying the geometric constraints of the cooling channels 56
(e.g., based upon the cooling channel dimensions discussed
above).
[0048] The surface features 96 may further enhance the forced
convective cooling of the liner 34 by increasing the surface area
of the downstream end portion 52 via which the cooling air 84 may
contact as it flows through the channel 56. Thus, in the present
embodiment, as the air 84 flows through the channel 56 and contacts
the surface features 96, the amount of heat transferred away from
the liner 34 may be greater relative to the embodiment shown in
FIG. 7A, in which the cooling channel 56 has a substantially flat
and/or smooth surface 95. Further, while the presently illustrated
embodiments show surface features 96 formed only on the surface 95,
in other embodiments, the surface features 96 may also be formed on
the sidewalls 92 of the channel 56.
[0049] Continuing now to FIGS. 8A-8D, cross-sectional side views of
the cooling channel 56 within the circular region defined by the
arcuate line 8-8 in FIG. 6 is illustrated. Particularly, FIGS.
8A-8D illustrate several shapes in which the film holes 64 may be
formed, in accordance with embodiments of the present invention.
For instance, referring to the embodiment shown in FIG. 8A, a film
hole 64 extending through the downstream end portion 52 within the
cooling channel 56 may include edges 104 that are parallel to each
other and perpendicular to the longitudinal axis 58 of the liner
34. In other words, the edges 104 of the film hole 64 may define a
straight cylindrical passage with an angle of approximately 90
degrees relative to the inner and outer surfaces of the liner 34.
Thus, the outer opening 100 (adjacent to the cooling channel 56)
and the inner opening 102 (adjacent to the combustion chamber 38)
are substantially equal in size. As described above, the film hole
64 may provide a path through which a portion 53 of cooling air 84
flowing through the channel 56 may flow directly into the
combustion chamber 38 to provide an insulting cooling film 86.
[0050] FIG. 8B shows an alternate embodiment of the film hole 64 in
which the edges 104 are parallel to each other, but are angled with
respect to the longitudinal axis 58 of the liner 34. In other
words, the edges 104 of the film hole 64 may define a straight
cylindrical passage with an angle between approximately 0 and 90
degrees, 30 and 60 degrees, or about 45 degrees relative to the
inner and outer surfaces of the liner 34. Thus, the outer opening
100 and the inner opening 102 are also substantially equal in size,
but the path through which the portion of air 53 flows into the
combustion chamber 38 may be angled based upon the angle of the
edges 104.
[0051] FIG. 8C shows a further embodiment in which the film hole 64
is tapered, such that the outer opening 100 is smaller relative to
the inner opening 102. In other words, the edges 104 of the film
hole 64 may define a diverging passage, e.g., generally conical
shaped, from the inner surface (e.g., along cooling channel 56) to
the outer surface (e.g., in combustion chamber 38). Furthermore, a
centerline of the film hole 64 may have an angle between
approximately 0 and 90 degrees, 30 and 60 degrees, or about 45 or
90 degrees relative to the inner and outer surfaces of the liner
34.
[0052] FIG. 8D shows yet another embodiment in which the film hole
64 is tapered, such that the outer opening 100 is larger relative
to the inner opening 102. In other words, the edges 104 of the film
hole 64 may define a converging passage, e.g., generally conical
shaped, from the inner surface (e.g., along cooling channel 56) to
the outer surface (e.g., in combustion chamber 38). Furthermore, a
centerline of the film hole 64 may have an angle between
approximately 0 and 90 degrees, 30 and 60 degrees, or about 45 or
90 degrees relative to the inner and outer surfaces of the liner
34.
[0053] As mentioned above, the diameters of the film holes 64 may
be at least less than approximately 0.01, 0.02, 0.03, 0.04, 0.05,
0.06, 0.07, 0.08, 0.09, or 0.10 inches. In other embodiments, the
film holes 64 may be less than 0.01 inches or greater than 0.10
inches. Further, while the film holes 64 depicted in FIGS. 8A-8D
are shown as being generally circular in shape, it should be
appreciated that in other embodiments, the film holes 62 may be
square-shaped, rectangular shaped, oval-shaped, or any other type
of suitable geometric shape, and may be formed using any suitable
technique, such as laser drilling. Still further, it should be
understood that the various embodiments of the film holes 64
depicted herein may be similarly applied in forming the bypass
openings 41 located upstream from the downstream end portion
52.
[0054] FIG. 9 shows a partial cross-sectional side view of the
combustor 16 within the circular region defined by the arcuate line
6-6 in FIG. 3, in accordance with a further embodiment of the
invention. Particularly, FIG. 9 depicts an embodiment in which a
plurality of sets of film holes 64 are provided and axially spaced
along the axial length of the cooling channel 56. For instance, in
the illustrated embodiment, the channel 56 may include a first set
of film holes 64a, a second set of film holes 64b located
downstream from the first set 64a, and a third set of film holes
64c located downstream from the second set 64b. Thus, as the
cooling air 84 flows into the cooling channel 56 by way of the
inlets 68 on the wrapper 54, portions 53 of the cooling air 84 may
flow through each set of film holes 64a, 64b, and 64c in series. As
will be appreciated, this arrangement may not only increases the
amount of air 53 supplied to the combustion chamber 38, but also
distribute the air 53 more evenly across the inner surface of the
downstream end portion 52, thus providing a more uniform cooling
film 86 for insulating the liner 34.
[0055] Additionally, FIG. 9 also illustrates the use of multiple
sets of bypass openings 41. For instance, referring back to the
embodiment shown in FIGS. 4 and 5, a single set of bypass openings
41 disposed circumferentially about the liner 34 is illustrated. In
FIG. 9, three such sets of axially spaced bypass openings, referred
to here by reference numbers 41a, 41b, and 41c, may be utilized in
cooling the liner 34. That is, each of the bypass openings shown in
the cross-sectional view of FIG. 9 may correspond to a respective
set of bypass openings arranged circumferentially about the liner
34. A portion of air 43 from the annular passage 36 may flow into
each of the bypass openings 41a, 41b, and 41c into the combustion
chamber 38. As discussed above, this air flow 43 may provide an
additional cooling film, or may contribute to the cooling film 86
that is supplied via the air flow 53 through the film holes 64a,
64b, and 64c. As will be appreciated, the use of multiple sets of
bypass openings 41a, 41b, and 41c may further increase the area and
uniformity of the cooling film 86, thus further improving the
insulation of the liner 34 from the combustion gases within the
chamber 38.
[0056] This written description uses examples to disclose the
invention, including the best mode, and also 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 may 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.
* * * * *