CA2685342A1 - Cooling holes for gas turbine combustor liner having a non-uniform diameter therethrough - Google Patents
Cooling holes for gas turbine combustor liner having a non-uniform diameter therethrough Download PDFInfo
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- CA2685342A1 CA2685342A1 CA002685342A CA2685342A CA2685342A1 CA 2685342 A1 CA2685342 A1 CA 2685342A1 CA 002685342 A CA002685342 A CA 002685342A CA 2685342 A CA2685342 A CA 2685342A CA 2685342 A1 CA2685342 A1 CA 2685342A1
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- Prior art keywords
- shell
- liner
- cooling
- diameter
- combustor
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- 238000001816 cooling Methods 0.000 title claims abstract description 120
- 238000011144 upstream manufacturing Methods 0.000 claims abstract description 10
- 238000009792 diffusion process Methods 0.000 claims description 8
- 238000000034 method Methods 0.000 claims description 7
- 239000000446 fuel Substances 0.000 claims description 6
- 230000007423 decrease Effects 0.000 claims description 3
- 239000007789 gas Substances 0.000 description 10
- 238000002485 combustion reaction Methods 0.000 description 8
- 238000010790 dilution Methods 0.000 description 6
- 239000012895 dilution Substances 0.000 description 6
- 239000002184 metal Substances 0.000 description 3
- 230000000712 assembly Effects 0.000 description 2
- 238000000429 assembly Methods 0.000 description 2
- 239000000567 combustion gas Substances 0.000 description 2
- 230000006978 adaptation Effects 0.000 description 1
- 230000004323 axial length Effects 0.000 description 1
- 230000008859 change Effects 0.000 description 1
- RLQJEEJISHYWON-UHFFFAOYSA-N flonicamid Chemical compound FC(F)(F)C1=CC=NC=C1C(=O)NCC#N RLQJEEJISHYWON-UHFFFAOYSA-N 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 239000000203 mixture Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000008569 process Effects 0.000 description 1
- 230000009467 reduction Effects 0.000 description 1
- 230000008439 repair process Effects 0.000 description 1
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/002—Wall structures
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
- F23R3/06—Arrangement of apertures along the flame tube
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/03042—Film cooled combustion chamber walls or domes
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A gas turbine combustor liner (12), including a shell having a first end (13) adjacent to an upstream end of the combustor and a second end (15) adjacent to a downstream end of the combustor, where the shell also has a hot side (36), a cold side (38), and a centerline axis therethrough. A plurality of small, closely-spaced film cooling holes (44) 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 (38) of the shell having a first diameter and a second opening located at the ho.t side (36) 2of 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.
Description
COOLING HOLES FOR GAS TURBINE COMBUSTOR LINER HAVING A NON-UNIFORM DIAMETER THERETHROUGH
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.
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 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. Patent 5,181,379 to Wakeman et al., U.S.
Patent 5,233,828 to Napoli, and U.S. Patent 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. Patent 6,205,789 to Patterson et al., U.S. Patent 6,655,149 to Farmer et al., and U.S. Patent 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.
Patent 5,233,828 to Napoli, and U.S. Patent 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. Patent 6,205,789 to Patterson et al., U.S. Patent 6,655,149 to Farmer et al., and U.S. Patent 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
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.
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
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
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.
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 to approximately 35 . 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 to approximately 65 . 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.
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 to approximately 35 . 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 to approximately 65 . 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 521ocated at cold side 38 (for outer liner 12) having a first diameter 54 and a second opening 561ocated 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 3.0-5Ø
More specifically, it will be seen that each cooling hole 44 preferably includes a first opening 521ocated at cold side 38 (for outer liner 12) having a first diameter 54 and a second opening 561ocated 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 3.0-5Ø
[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 10 to approximately 15 . A more preferred range for diffusion angle 60 is approximately 3 to approximately 100, while an optimal range for diffusion angle 60 is approximately 5 to approximately 9 . 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 numera176 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.
[0025] 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.
[0026] 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.
[0027] 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.
[0028] 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 1461ocated 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 801ocated 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.
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.
[0029] 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 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.
Secondly, first portion 146 of cooling hole 144 is formed through second portion 148 so that first portion 146 has a substantially uniform diameter.
[0030] 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.
[00311 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.
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.
[00311 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.
Claims (23)
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.
(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.
(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 3.0-5Ø
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° to approximately 15°.
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° to approximately 10°.
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° to approximately 9°.
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° to approximately 35°.
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° to approximately 60°.
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.
(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.
(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.
Applications Claiming Priority (3)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US11/743,126 US20080271457A1 (en) | 2007-05-01 | 2007-05-01 | Cooling Holes For Gas Turbine Combustor Having A Non-Uniform Diameter Therethrough |
| US11/743,126 | 2007-05-01 | ||
| PCT/US2008/055758 WO2008137201A1 (en) | 2007-05-01 | 2008-03-04 | Cooling holes for gas turbine combustor liner having a non-uniform diameter therethrough |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA2685342A1 true CA2685342A1 (en) | 2008-11-13 |
Family
ID=39645559
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA002685342A Abandoned CA2685342A1 (en) | 2007-05-01 | 2008-03-04 | Cooling holes for gas turbine combustor liner having a non-uniform diameter therethrough |
Country Status (5)
| Country | Link |
|---|---|
| US (1) | US20080271457A1 (en) |
| EP (1) | EP2153133A1 (en) |
| JP (1) | JP2010526274A (en) |
| CA (1) | CA2685342A1 (en) |
| WO (1) | WO2008137201A1 (en) |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE102009007164A1 (en) | 2009-02-03 | 2010-08-12 | Rolls-Royce Deutschland Ltd & Co Kg | A method of forming a cooling air opening in a wall of a gas turbine combustor and combustor wall made by the method |
| US8307657B2 (en) * | 2009-03-10 | 2012-11-13 | General Electric Company | Combustor liner cooling system |
| US20100263384A1 (en) * | 2009-04-17 | 2010-10-21 | Ronald James Chila | Combustor cap with shaped effusion cooling holes |
| US8371814B2 (en) * | 2009-06-24 | 2013-02-12 | Honeywell International Inc. | Turbine engine components |
| US8529193B2 (en) * | 2009-11-25 | 2013-09-10 | Honeywell International Inc. | Gas turbine engine components with improved film cooling |
| US8628293B2 (en) | 2010-06-17 | 2014-01-14 | Honeywell International Inc. | Gas turbine engine components with cooling hole trenches |
| US9052113B1 (en) * | 2011-06-06 | 2015-06-09 | General Electric Company | Combustor nozzle and method for modifying the combustor nozzle |
| US9650900B2 (en) | 2012-05-07 | 2017-05-16 | Honeywell International Inc. | Gas turbine engine components with film cooling holes having cylindrical to multi-lobe configurations |
| US10113433B2 (en) | 2012-10-04 | 2018-10-30 | Honeywell International Inc. | Gas turbine engine components with lateral and forward sweep film cooling holes |
| US9194585B2 (en) | 2012-10-04 | 2015-11-24 | United Technologies Corporation | Cooling for combustor liners with accelerating channels |
| US9309809B2 (en) * | 2013-01-23 | 2016-04-12 | General Electric Company | Effusion plate using additive manufacturing methods |
| US9939154B2 (en) * | 2013-02-14 | 2018-04-10 | United Technologies Corporation | Combustor liners with U-shaped cooling channels |
| CA2907048A1 (en) | 2013-03-15 | 2014-09-25 | President And Fellows Of Harvard College | Low porosity auxetic sheet |
| WO2015030927A1 (en) * | 2013-08-30 | 2015-03-05 | United Technologies Corporation | Contoured dilution passages for a gas turbine engine combustor |
| EP3066388B1 (en) * | 2013-11-04 | 2024-04-10 | RTX Corporation | Turbine engine combustor heat shield with multi-angled cooling apertures |
| US10190773B2 (en) * | 2013-11-05 | 2019-01-29 | United Technologies Corporation | Attachment stud on a combustor floatwall panel with internal cooling holes |
| JP6456481B2 (en) * | 2014-08-26 | 2019-01-23 | シーメンス エナジー インコーポレイテッド | Film cooling hole array for an acoustic resonator in a gas turbine engine |
| GB201419327D0 (en) | 2014-10-30 | 2014-12-17 | Rolls Royce Plc | A cooled component |
| US20160258623A1 (en) * | 2015-03-05 | 2016-09-08 | United Technologies Corporation | Combustor and heat shield configurations for a gas turbine engine |
| US11021965B2 (en) | 2016-05-19 | 2021-06-01 | Honeywell International Inc. | Engine components with cooling holes having tailored metering and diffuser portions |
| US20180266687A1 (en) * | 2017-03-16 | 2018-09-20 | General Electric Company | Reducing film scrubbing in a combustor |
| US11149948B2 (en) * | 2017-08-21 | 2021-10-19 | General Electric Company | Fuel nozzle with angled main injection ports and radial main injection ports |
| US11009230B2 (en) | 2018-02-06 | 2021-05-18 | Raytheon Technologies Corporation | Undercut combustor panel rail |
| US10830435B2 (en) | 2018-02-06 | 2020-11-10 | Raytheon Technologies Corporation | Diffusing hole for rail effusion |
| US11248791B2 (en) | 2018-02-06 | 2022-02-15 | Raytheon Technologies Corporation | Pull-plane effusion combustor panel |
| US11022307B2 (en) | 2018-02-22 | 2021-06-01 | Raytheon Technology Corporation | Gas turbine combustor heat shield panel having multi-direction hole for rail effusion cooling |
| US11306659B2 (en) | 2019-05-28 | 2022-04-19 | Honeywell International Inc. | Plug resistant effusion holes for gas turbine engine |
| US11371701B1 (en) | 2021-02-03 | 2022-06-28 | General Electric Company | Combustor for a gas turbine engine |
| US11959643B2 (en) | 2021-06-07 | 2024-04-16 | General Electric Company | Combustor for a gas turbine engine |
| US11885495B2 (en) | 2021-06-07 | 2024-01-30 | General Electric Company | Combustor for a gas turbine engine including a liner having a looped feature |
| US11774098B2 (en) | 2021-06-07 | 2023-10-03 | General Electric Company | Combustor for a gas turbine engine |
| CN115493163B (en) * | 2022-09-06 | 2024-02-20 | 清华大学 | Combustion chamber flame tube and high-efficiency cooling method thereof |
Family Cites Families (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2221979B (en) * | 1988-08-17 | 1992-03-25 | Rolls Royce Plc | A combustion chamber for a gas turbine engine |
| US5660525A (en) * | 1992-10-29 | 1997-08-26 | General Electric Company | Film cooled slotted wall |
| US5528904A (en) * | 1994-02-28 | 1996-06-25 | Jones; Charles R. | Coated hot gas duct liner |
| JPH09119322A (en) * | 1995-10-27 | 1997-05-06 | Ishikawajima Harima Heavy Ind Co Ltd | Cooling liner of aircraft engine |
| US6383602B1 (en) * | 1996-12-23 | 2002-05-07 | General Electric Company | Method for improving the cooling effectiveness of a gaseous coolant stream which flows through a substrate, and related articles of manufacture |
| DE59808269D1 (en) * | 1998-03-23 | 2003-06-12 | Alstom Switzerland Ltd | Film cooling hole |
| CA2288557C (en) * | 1998-11-12 | 2007-02-06 | Mitsubishi Heavy Industries, Ltd. | Gas turbine combustor cooling structure |
| US6408629B1 (en) * | 2000-10-03 | 2002-06-25 | General Electric Company | Combustor liner having preferentially angled cooling holes |
| US7093439B2 (en) * | 2002-05-16 | 2006-08-22 | United Technologies Corporation | Heat shield panels for use in a combustor for a gas turbine engine |
| US7328580B2 (en) * | 2004-06-23 | 2008-02-12 | General Electric Company | Chevron film cooled wall |
| JP4898253B2 (en) * | 2005-03-30 | 2012-03-14 | 三菱重工業株式会社 | High temperature components for gas turbines |
| EP1712739A1 (en) * | 2005-04-12 | 2006-10-18 | Siemens Aktiengesellschaft | Component with film cooling hole |
-
2007
- 2007-05-01 US US11/743,126 patent/US20080271457A1/en not_active Abandoned
-
2008
- 2008-03-04 EP EP08731322A patent/EP2153133A1/en not_active Withdrawn
- 2008-03-04 JP JP2010506353A patent/JP2010526274A/en active Pending
- 2008-03-04 WO PCT/US2008/055758 patent/WO2008137201A1/en active Application Filing
- 2008-03-04 CA CA002685342A patent/CA2685342A1/en not_active Abandoned
Also Published As
| Publication number | Publication date |
|---|---|
| JP2010526274A (en) | 2010-07-29 |
| EP2153133A1 (en) | 2010-02-17 |
| WO2008137201A1 (en) | 2008-11-13 |
| US20080271457A1 (en) | 2008-11-06 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| EEER | Examination request |
Effective date: 20130117 |
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| FZDE | Discontinued |
Effective date: 20150427 |