U.S. patent application number 11/743126 was filed with the patent office on 2008-11-06 for cooling holes for gas turbine combustor having a non-uniform diameter therethrough.
This patent application is currently assigned to General Electric Company. Invention is credited to David Louis Burrus, George Chia-Chun Hsiao, Marie Ann McMasters, Hukam Chand Mongia, Hongtao Zhang.
Application Number | 20080271457 11/743126 |
Document ID | / |
Family ID | 39645559 |
Filed Date | 2008-11-06 |
United States Patent
Application |
20080271457 |
Kind Code |
A1 |
McMasters; Marie Ann ; et
al. |
November 6, 2008 |
Cooling Holes For Gas Turbine Combustor Having A Non-Uniform
Diameter Therethrough
Abstract
A gas turbine combustor liner, including a shell having a first
end adjacent to an upstream end of the combustor and a second end
adjacent to a downstream end of the combustor, where the shell also
has a hot side, a cold side, and a centerline axis therethrough. A
plurality of small, closely-spaced film cooling holes are formed in
the shell through which air flows for providing a cooling film
along the hot side of the shell. Each cooling hole has a
non-uniform diameter as it extends through the shell. In
particular, each cooling hole includes a first opening located at
the cold side of the shell having a first diameter and a second
opening located at the hot side of the shell having a second
diameter, wherein the second diameter of the second opening is
larger than the first diameter of the first opening. It is
preferred that the shape of each cooling hole be substantially
frusto-conical.
Inventors: |
McMasters; Marie Ann;
(Mason, OH) ; Burrus; David Louis; (Cincinnati,
OH) ; Hsiao; George Chia-Chun; (West Chester, OH)
; Zhang; Hongtao; (Mason, OH) ; Mongia; Hukam
Chand; (West Chester, OH) |
Correspondence
Address: |
JAMES P. DAVIDSON, ESQ.
8375 ASHMONT WAY
MASON
OH
45040
US
|
Assignee: |
General Electric Company
Schenectady
NY
|
Family ID: |
39645559 |
Appl. No.: |
11/743126 |
Filed: |
May 1, 2007 |
Current U.S.
Class: |
60/754 ;
416/95 |
Current CPC
Class: |
Y02T 50/675 20130101;
Y02T 50/60 20130101; F23R 3/002 20130101; F23R 2900/03042 20130101;
F23R 3/06 20130101 |
Class at
Publication: |
60/754 ;
416/95 |
International
Class: |
F23R 3/04 20060101
F23R003/04 |
Claims
1. A gas turbine combustor liner, comprising: (a) a shell having a
first end adjacent to an upstream end of said combustor and a
second end adjacent to a downstream end of said combustor, said
shell also having a hot side, a cold side, and a centerline axis
therethrough; and, (b) a plurality of small, closely-spaced film
cooling holes formed in said shell through which air flows for
providing a cooling film along said hot side of said shell, said
cooling holes having a non-uniform diameter through said shell.
2. The combustor liner of claim 1, each of said cooling holes
further comprising: (a) a first opening located at said cold side
of said shell having a first diameter; and, (b) a second opening
located at said hot side of said shell having a second diameter;
wherein said second diameter of said second opening is larger than
said first diameter of said first opening.
3. The combustor liner of claim 2, wherein a ratio of said second
diameter for said second opening to said first diameter for said
first opening is approximately 3.0-5.0.
4. The combustor liner of claim 2, wherein spacing between adjacent
first openings of said cooling holes is greater than spacing
between adjacent second openings of said cooling holes.
5. The combustor liner of claim 2, wherein spacing between adjacent
first openings of said cooling holes is approximately 3.0-6.0 times
said first diameter.
6. The combustor liner of claim 2, wherein spacing between said
adjacent second openings of said cooling holes is approximately
0.2-0.7 times said second diameter.
7. The combustor liner of claim 1, wherein a diameter of each
cooling hole grows progressively larger from said cold side of said
shell to said hot side of said shell.
8. The combustor liner of claim 7, wherein each said cooling hole
has a substantially frusto-conical shape.
9. The combustor liner of claim 7, wherein each said cooling hole
has an angle of diffusion with respect to an axis therethrough of
approximately 1.degree. to approximately 15.degree..
10. The combustor liner of claim 7, wherein each said cooling hole
has an angle of diffusion with respect to an axis therethrough of
approximately 3.degree. to approximately 10.degree..
11. The combustor liner of claim 7, wherein each said cooling hole
has an angle of diffusion with respect to an axis therethrough of
approximately 5.degree. to approximately 9.degree..
12. The combustor liner of claim 1, wherein an axis through each
said cooling hole is oriented at an axial angle to said centerline
axis in a range of approximately 15.degree. to approximately
35.degree..
13. The combustor liner of claim 1, wherein an axis through each
said cooling hole is oriented at a circumferential angle to said
centerline axis in a range of approximately 30.degree. to
approximately 60.degree..
14. The combustor liner of claim 1, wherein a ratio of a jet
velocity of air at said hot side of said shell to a jet velocity of
air at said cold side of said shell is approximately 0.25-0.50.
15. The combustor liner of claim 1, each said cooling hole further
comprising: (a) a first portion having a substantially uniform
diameter through said liner; and, (b) a second portion having a
non-uniform diameter through said liner.
16. The combustor liner of claim 15, wherein said first portion of
said cooling hole is located adjacent said cold side of said
shell.
17. The combustor liner of claim 15, wherein said first portion of
said cooling hole is located adjacent said hot side of said
shell.
18. The combustor liner of claim 1, wherein said cooling holes are
provided over essentially an entire length of said shell.
19. The combustor liner of claim 1, wherein said cooling holes are
provided at certain designated locations of said shell.
20. The combustor liner of claim 1, wherein said cooling holes are
provided at an upstream portion of said shell.
21. The combustor liner of claim 1, wherein said cooling holes are
provided in said shell immediately downstream of an air-fuel mixer
for said combustor.
22. A method of forming a cooling hole in a liner of a gas turbine
engine combustor, wherein said cooling hole has a non-uniform
diameter therethrough, comprising the following steps: (a) forming
a first portion of said cooling hole from a hot side of said liner,
wherein said first portion is substantially conical in shape and
extends substantially through said liner; and, (b) forming a second
portion of said cooling hole from said first portion of said
cooling hole to a cold side of said liner, wherein said second
portion is substantially uniform in diameter.
23. The method of claim 22, wherein said first portion of said
cooling hole has a diameter which progressively decreases from said
hot side of said liner.
Description
BACKGROUND OF THE INVENTION
[0001] The present invention relates generally to a liner for a gas
turbine engine combustor and, in particular, to the configuration
of the cooling holes utilized in a multihole cooling scheme for
such liner.
[0002] Combustor liners are generally used in the combustion
section of a gas turbine engine located between the compressor and
turbine sections of the engine, although such liners may also be
used in the exhaust sections of aircraft engines that employ
augmentors. Combustors generally include an exterior casing and an
interior combustor where fuel is burned to produce a hot gas at an
intensely high temperature (e.g., 3000.degree. F. or even higher).
To prevent this intense heat from damaging the combustor case and
the surrounding engine before it exits to a turbine, a heat shield
or combustor liner is provided in the interior of the
combustor.
[0003] Various liner designs have been disclosed in the art having
different types of cooling schemes. One example of a liner design
includes a plurality of cooling holes being formed in an annular
one-piece liner to provide film cooling along the hot side of the
liner (e.g., U.S. Pat. No. 5,181,379 to Wakeman et al., U.S. Pat.
No. 5,233,828 to Napoli, and U.S. Pat. No. 5,465,572 to Nicoll et
al.). It will also be appreciated that various patterns, sizes and
densities of cooling holes have been employed in such multihole
cooling of liners. This is disclosed in U.S. Pat. No. 6,205,789 to
Patterson et al., U.S. Pat. No. 6,655,149 to Farmer et al., and
U.S. Pat. No. 7,086,232 to Moertle et al. In each case, it will be
seen that the individual cooling holes are formed straight through
the liner with a constant or uniform diameter.
[0004] While each of the aforementioned patents has progressed the
state of the art, it has been found that hot streaks still occur
between adjacent rows of holes in the current multihole cooling
patterns. These hot streaks eventually result in cracks to the
liner, thereby necessitating removal of the liner for repair.
[0005] Thus, it would be desirable for a combustor liner to be
developed for use with a gas turbine engine combustor which
includes a multihole cooling scheme that minimizes hot streaks,
reduces the amount of metal surface of the liner exposed along the
hot side thereof, and increases the durability of the liner. It
would also be desirable for the configuration of the individual
cooling holes to reduce the temperature along the hot side of the
liner, as well as enhance bore cooling of the liner itself.
Further, it is desirable for the cooling holes to reduce the jet
velocity of cooling air along the hot side of the liner, and
thereby promote more effective film cooling.
BRIEF SUMMARY OF THE INVENTION
[0006] In accordance with a first exemplary embodiment of the
invention, a gas turbine combustor liner is disclosed as including
a shell having a first end adjacent to an upstream end of the
combustor and a second end adjacent to a downstream end of the
combustor, where the shell also has a hot side, a cold side, and a
centerline axis therethrough. A plurality of small, closely-spaced
film cooling holes are formed in the shell through which air flows
for providing a cooling film along the hot side of the shell. Each
cooling hole has a non-uniform diameter as it extends through the
shell. In particular, each cooling hole includes a first opening
located at the cold side of the shell having a first diameter and a
second opening located at the hot side of the shell having a second
diameter, wherein the second diameter of the second opening is
larger than the first diameter of the first opening. It is
preferred that the shape of each cooling hole be substantially
frusto-conical.
[0007] In a second exemplary embodiment of the invention, a gas
turbine combustor liner is disclosed as including a shell having a
first end adjacent to an upstream end of the combustor and a second
end adjacent to a downstream end of the combustor, where the shell
also has a hot side, a cold side, and a centerline axis
therethrough. A plurality of small, closely-spaced film cooling
holes are formed in the shell through which air flows for providing
a cooling film along the hot side of the shell. In particular, each
cooling hole includes a first portion having a substantially
uniform diameter through said liner and a second portion having a
non-uniform diameter through said liner.
[0008] In a third exemplary embodiment of the invention, a method
of forming a cooling hole in a liner of a gas turbine engine
combustor is disclosed, wherein the cooling hole has a non-uniform
diameter therethrough. The method includes the following steps:
forming a first portion of the cooling hole from a hot side of the
liner, wherein the first portion is substantially conical in shape
and extends substantially through the liner; and, forming a second
portion of the cooling hole from the first portion of the cooling
hole to a cold side of the liner, wherein the second portion is
substantially uniform in diameter. According to this method, the
first portion of the cooling hole has a diameter which
progressively decreases from the hot side of the liner.
BRIEF DESCRIPTION OF THE DRAWINGS
[0009] FIG. 1 is a perspective view of a combustor for a gas
turbine engine, where liners having cooling holes in accordance
with the present invention are depicted;
[0010] FIG. 2 is a partial sectional view of the outer liner for
the combustor depicted in FIG. 1, wherein cooling holes in
accordance with the present invention are shown;
[0011] FIG. 3 is a partial top perspective view of a portion of the
combustor outer liner depicted in FIG. 2;
[0012] FIG. 4 is a partial bottom perspective view of a portion of
the combustor outer liner depicted in FIGS. 2 and 3;
[0013] FIG. 5 is an enlarged partial sectional view of the
combustor outer liner depicted in FIGS. 1-4 taken in the
axial-radial plane;
[0014] FIG. 6 is an enlarged partial section view of the combustor
outer liner depicted in FIG. 1-4 taken in the
circumferential-radial plane;
[0015] FIG. 7 is an enlarged partial sectional view of the
combustor outer liner depicted in FIG. 2 taken in the axial-radial
plane, where the cooling hole has an alternate configuration;
[0016] FIG. 8 is an enlarged partial top view of the combustor
outer liner depicted in FIGS. 1-4; and,
[0017] FIG. 9 is an enlarged partial bottom view of the combustor
outer liner depicted in FIGS. 1-4.
DETAILED DESCRIPTION OF THE INVENTION
[0018] Referring now to the drawings in detail, wherein identical
numerals indicate the same elements throughout the figures, FIG. 1
depicts a combustor 10 of the type suitable for use in a gas
turbine engine. Combustor 10 includes an outer liner 12 and an
inner liner 14 disposed between an outer combustor casing 16 and an
inner combustor casing 18. Outer and inner liners 12 and 14 are
radially spaced from each other to define a combustion chamber 20.
Outer liner 12 and outer casing 16 form an outer passage 22
therebetween, and inner liner 14 and inner casing 18 form an inner
passage 24 therebetween. A cowl assembly 26 is mounted to the
upstream ends of outer and inner liners 12 and 14. An annular
opening 28 is formed in cowl assembly 26 for the introduction of
compressed air into combustor 10. The compressed air is supplied
from a compressor (not shown) in a direction generally indicated by
arrow 25 of FIG. 1. The compressed air passes principally through
annular opening 28 to support combustion and partially into outer
and inner passages 22 and 24 where it is used to cool liners 12 and
14.
[0019] Disposed between and interconnecting outer and inner liners
12 and 14 near their upstream ends is an annular dome plate 30. A
plurality of circumferentially spaced swirler assemblies 32 are
mounted in dome plate 30. Each swirler assembly 32 receives
compressed air from annular opening 28 and fuel from a
corresponding fuel tube 34. The fuel and air are swirled and mixed
by swirler assemblies 32, and the resulting fuel/air mixture is
discharged into combustion chamber 20. It is noted that although
FIG. 1 illustrates one preferred embodiment of a single annular
combustor, the present invention is equally applicable to any type
of combustor, including multiple annular combustors, which utilizes
multihole film cooling.
[0020] Outer and inner liners 12 and 14 each comprise a single
wall, metal shell having a generally annular and axially extending
configuration. Outer liner 12 includes a first end 13 adjacent to
an upstream end of combustor 10 and a second end 15 adjacent to a
downstream end of combustor 10. Likewise, inner liner 14 includes a
first end 17 adjacent to an upstream end of combustor 10 and a
second end 19 adjacent to a downstream end of combustor 10. Outer
liner 12 has a hot side 36 facing the hot combustion gases in
combustion chamber 20 and a cold side 38 in contact with the
relatively cool air in outer passage 22. Similarly, inner liner 14
has a hot side 40 facing the hot combustion gases in combustion
chamber 20 and a cold side 42 in contact with the relatively cool
air in inner passage 24. Both liners 12 and 14 include a plurality
of small, closely-spaced film cooling holes 44 formed therein
through which air flows for providing a cooling film along hot
sides 36 and 40 of outer and inner liners 12 and 14,
respectively.
[0021] As seen in FIGS. 2-6 and 8-9, cooling holes 44 disposed
through at least a portion of outer liner 12 are shown in more
detail. Although cooling holes 44 are depicted in outer liner 12,
it should be understood that the configuration of cooling holes of
inner liner 14 is substantially identical to that of outer liner
12. As such, the following description will also apply to inner
liner 14. FIGS. 3 and 4 include a frame of reference having axes
35, 37 and 39, wherein axis 35 is in the axial direction through
combustor 10, axis 37 is in the circumferential direction, and axis
39 is in the radial direction. As best seen in FIG. 5, cooling
holes 44 are preferably axially slanted from cold side 38 to hot
side 36 at a downstream angle 45, which is preferably in the range
of approximately 15.degree. to approximately 35.degree.. Cooling
holes 44 may also be circumferentially slanted or clocked at a
clock angle 55, as shown in FIG. 6. Clock angle 55 preferably
corresponds to the swirl of flow through combustor chamber 20,
which is typically in the range of approximately 30.degree. to
approximately 65.degree.. It will further be seen from FIGS. 3 and
4 that cooling holes 44 are preferably arranged in a series of
circumferentially extending rows 46. Such rows 46 are also
preferably staggered as they extend downstream in an axial
direction.
[0022] Contrary to the cooling holes of the prior art, cooling
holes 44 are configured so as to have a non-uniform diameter 50
through outer liner 12. More specifically, it will be seen that
each cooling hole 44 preferably includes a first opening 52 located
at cold side 38 (for outer liner 12) having a first diameter 54 and
a second opening 56 located at hot side 36 of outer liner 12 having
a second diameter 58. It will be appreciated that diameter 58 of
second opening 56 is preferably larger than diameter 54 of first
opening 54. In particular, a ratio of second diameter 58 to first
diameter 54 preferably is approximately 3.0-5.0.
[0023] It will further be seen from FIGS. 5 and 6 that diameter 50
of cooling hole 44 preferably gets progressively larger from cold
side 38 of outer liner 12 to hot side 36 of outer liner 12. Thus,
it will be understood that an angle of diffusion (or included
angle) 60 exists with respect to an axis 65 extending through each
cooling hole 44. Angle of diffusion 60 is defined as an angle
extending omni-directionally from a focal point on axis 65 and
preferably is in a range of approximately 1.degree. to
approximately 15.degree.. A more preferred range for diffusion
angle 60 is approximately 3.degree. to approximately 10.degree.,
while an optimal range for diffusion angle 60 is approximately
5.degree. to approximately 9.degree.. In any event, it will be
appreciated that each cooling hole 44 will have a substantially
frusto-conical shape.
[0024] It will be appreciated that spacing (represented by
reference numeral 62 in FIG. 8) between adjacent first openings 64
and 66 of adjacent cooling holes 68 and 70 is approximately 3.0-6.0
times first diameter 54 thereof. This corresponds generally to the
spacing utilized in current multihole cooling designs and therefore
does not necessitate a change to the flow of cooling air provided
to outer and inner liners 12 and 14, respectively. Spacing between
adjacent second openings 72 and 74 is represented by reference
numeral 76 in FIG. 9 and preferably is approximately 0.2-0.7 times
second diameter 58 thereof. In order to provide some means of
comparison to first openings 52 on cold sides 38 and 42 of outer
and inner liners 12 and 14, it will be understood that spacing 76
between adjacent second openings 72 and 74 is preferably
approximately 2.0-5.0 times diameter 54 of first openings 64 and
66. Because of the shorter spacing between adjacent second openings
56 of cooling holes 44, it will be appreciated that less metal of
outer and inner liners 12 and 14 is provided on hot sides 40 and 36
thereof is exposed to the harsh environment of combustion chamber
20. Also, by minimizing the spacing between second openings 56 of
cooling holes 44, the air flowing through cooling holes 44 is
better able to work in concert to eliminate or minimize hot streaks
on hot sides 40 and 36.
[0025] It will be appreciated that no dilution holes are shown
within outer and inner liners 12 and 14. Nevertheless, dilution air
may be introduced into combustor chamber 20 through a plurality of
circumferentially spaced dilution holes disposed in each of outer
and inner liners 12 and 14 to promote additional combustion when
desired. Such dilution holes would generally be far smaller in
number than cooling holes 44, with a cross-sectional area that is
substantially greater than the cross-sectional area of one of
cooling holes 44. It will be understood that cooling holes 44 will
serve to admit some dilution air into combustor chamber 20.
Additionally, the disclosed configuration of cooling holes 44 is
able to enhance bore cooling of outer and inner liners 12 and 14
since the overall volume thereof has increased.
[0026] As indicated by an arrow 75 (see FIG. 3), it is preferred
that cooling air enter first opening 54 of each cooling hole 44
with a predetermined jet velocity on the order of approximately
200-300 feet per second. Due to diffusion angle 60 of cooling hole
44, wherein second opening 56 has a larger diameter 58 than
diameter 54 of first opening 52, cooling air (indicated by arrow
85) at hot side 38 of outer liner 12 has a jet velocity that is
approximately 75-100 feet per second. Accordingly, the jet velocity
of cooling air 85 is less than that for a conventional straight
(i.e., uniform diameter) cooling hole. By comparison, the jet
velocity of cooling air 85 at second opening 56 is approximately
30%-50% less than the jet velocity of cooling air 75 at first
opening 52. This reduction in the jet velocity of cooling air 85
along hot side 38 of outer liner 12 assists to promote more
effective film cooling and is less apt to penetrate
therethrough.
[0027] As shown in FIG. 7, an alternate configuration for cooling
holes 44 is provided for outer liner 12. In this embodiment, each
cooling hole 144 includes a first portion 146 located adjacent cold
side 38 of outer liner 12 and a second portion 148 located adjacent
hot side 36 of outer liner 12. It will be seen that first portion
146, which includes a first opening 152, has a substantially
uniform diameter 154 and extends a predetermined length 78 from
cold side 38 to a second end 80 located within a thickness 82 of
outer liner 12. Second portion 148, for its part, extends from
second end 80 of first portion 146 to second opening 156 on hot
side 36 of outer liner 12 so as to have a desired length 84 and
preferably a non-uniform diameter 158. While not shown, it will be
understood that second portion 156 having non-uniform diameter 158
may be located adjacent to cold side 38 of outer liner 12 and first
portion 146 having substantially uniform diameter 154 may be
located adjacent to hot side 36 of outer liner 12.
[0028] By configuring the cooling holes in outer and inner liners
12 and 14 like that described for cooling holes 144, the
manufacturing of such cooling holes is made less complex. In
accordance therewith, a method of forming a cooling hole 144 in
outer and inner liners 12 and 14 of combustor 10, where cooling
hole 144 has a non-uniform diameter therethrough is hereby
disclosed. In a first step, second portion 148 of cooling hole 144
is formed from hot side 36 of outer liner 12. It will be understood
that second portion 148 has a diameter 150 that progressively
decreases from hot side 36 of outer liner and extends a desired
length 84 through thickness 82 of outer liner 12. Thus, second
portion 148 is substantially conical in shape. Secondly, first
portion 146 of cooling hole 144 is formed through second portion
148 so that first portion 146 has a substantially uniform
diameter.
[0029] While it is primarily intended for cooling holes 44 and/or
cooling holes 144 to be provided over essentially an entire axial
length and circumference of outer and inner liners 12 and 14, it is
also possible that cooling holes have such configuration could be
provided only at certain designated locations thereof. This
includes, for example, areas of outer and inner liners 12 and 14
where hot streaks are known to occur. Exemplary locations for such
cooling holes may include adjacent to dilution holes 48, adjacent
to cooling nuggets present in the liners, immediately downstream of
a swirler assembly 32, upstream ends 13 and 17 of the liners, or
downstream ends 15 and 19 of the liners.
[0030] Having shown and described the preferred embodiment of the
present invention, further adaptations of cooling holes, as well as
the process for forming such cooling holes, can be accomplished by
appropriate modifications by one of ordinary skill in the art
without departing from the scope of the invention. Moreover, it
will be understood that the cooling holes described herein may be
utilized with other components of a gas turbine engine not depicted
herein, such as an afterburner liner.
* * * * *