US20080271457A1 - Cooling Holes For Gas Turbine Combustor Having A Non-Uniform Diameter Therethrough - Google Patents

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.

Description

BACKGROUND OF THE INVENTION
• 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.
• 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.
• 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.
• 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.
• 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
• 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 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.
• 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
• 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;
• FIG. 2 is a partial sectional view of the outer liner for the combustor depicted inFIG. 1, wherein cooling holes in accordance with the present invention are shown;
• FIG. 3 is a partial top perspective view of a portion of the combustor outer liner depicted inFIG. 2;
• FIG. 4 is a partial bottom perspective view of a portion of the combustor outer liner depicted inFIGS. 2 and 3;
• FIG. 5 is an enlarged partial sectional view of the combustor outer liner depicted inFIGS. 1-4 taken in the axial-radial plane;
• FIG. 6 is an enlarged partial section view of the combustor outer liner depicted inFIG. 1-4 taken in the circumferential-radial plane;
• FIG. 7 is an enlarged partial sectional view of the combustor outer liner depicted inFIG. 2 taken in the axial-radial plane, where the cooling hole has an alternate configuration;
• FIG. 8 is an enlarged partial top view of the combustor outer liner depicted inFIGS. 1-4; and,
• FIG. 9 is an enlarged partial bottom view of the combustor outer liner depicted inFIGS. 1-4.
DETAILED DESCRIPTION OF THE INVENTION
• Referring now to the drawings in detail, wherein identical numerals indicate the same elements throughout the figures,FIG. 1 depicts acombustor10 of the type suitable for use in a gas turbine engine. Combustor10 includes anouter liner12 and aninner liner14 disposed between anouter combustor casing16 and aninner combustor casing18. Outer andinner liners12 and14 are radially spaced from each other to define acombustion chamber20.Outer liner12 andouter casing16 form anouter passage22 therebetween, andinner liner14 andinner casing18 form aninner passage24 therebetween. Acowl assembly26 is mounted to the upstream ends of outer andinner liners12 and14. Anannular opening28 is formed incowl assembly26 for the introduction of compressed air intocombustor10. The compressed air is supplied from a compressor (not shown) in a direction generally indicated byarrow25 ofFIG. 1. The compressed air passes principally throughannular opening28 to support combustion and partially into outer andinner passages22 and24 where it is used to coolliners12 and14.
• Disposed between and interconnecting outer andinner liners12 and14 near their upstream ends is anannular dome plate30. A plurality of circumferentially spacedswirler assemblies32 are mounted indome plate30. Eachswirler assembly32 receives compressed air fromannular opening28 and fuel from acorresponding fuel tube34. The fuel and air are swirled and mixed byswirler assemblies32, and the resulting fuel/air mixture is discharged intocombustion chamber20. It is noted that althoughFIG. 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.
• Outer andinner liners12 and14 each comprise a single wall, metal shell having a generally annular and axially extending configuration.Outer liner12 includes afirst end13 adjacent to an upstream end ofcombustor10 and asecond end15 adjacent to a downstream end ofcombustor10. Likewise,inner liner14 includes afirst end17 adjacent to an upstream end ofcombustor10 and asecond end19 adjacent to a downstream end ofcombustor10.Outer liner12 has ahot side36 facing the hot combustion gases incombustion chamber20 and acold side38 in contact with the relatively cool air inouter passage22. Similarly,inner liner14 has ahot side40 facing the hot combustion gases incombustion chamber20 and acold side42 in contact with the relatively cool air ininner passage24. Bothliners12 and14 include a plurality of small, closely-spacedfilm cooling holes44 formed therein through which air flows for providing a cooling film alonghot sides36 and40 of outer andinner liners12 and14, respectively.
• As seen inFIGS. 2-6 and8-9,cooling holes44 disposed through at least a portion ofouter liner12 are shown in more detail. Althoughcooling holes44 are depicted inouter liner12, it should be understood that the configuration of cooling holes ofinner liner14 is substantially identical to that ofouter liner12. As such, the following description will also apply toinner liner14.FIGS. 3 and 4 include a frame ofreference having axes35,37 and39, whereinaxis35 is in the axial direction throughcombustor10,axis37 is in the circumferential direction, andaxis39 is in the radial direction. As best seen inFIG. 5, cooling holes44 are preferably axially slanted fromcold side38 tohot side36 at adownstream angle45, which is preferably in the range of approximately 15° to approximately 35°. Cooling holes44 may also be circumferentially slanted or clocked at aclock angle55, as shown inFIG. 6.Clock angle55 preferably corresponds to the swirl of flow throughcombustor chamber20, which is typically in the range of approximately 30° to approximately 65°. It will further be seen fromFIGS. 3 and 4 that cooling holes44 are preferably arranged in a series of circumferentially extendingrows46.Such rows46 are also preferably staggered as they extend downstream in an axial direction.
• Contrary to the cooling holes of the prior art, cooling holes44 are configured so as to have anon-uniform diameter50 throughouter liner12. More specifically, it will be seen that each coolinghole44 preferably includes afirst opening52 located at cold side38 (for outer liner12) having afirst diameter54 and asecond opening56 located athot side36 ofouter liner12 having asecond diameter58. It will be appreciated thatdiameter58 ofsecond opening56 is preferably larger thandiameter54 offirst opening54. In particular, a ratio ofsecond diameter58 tofirst diameter54 preferably is approximately 3.0-5.0.
• It will further be seen fromFIGS. 5 and 6 thatdiameter50 of coolinghole44 preferably gets progressively larger fromcold side38 ofouter liner12 tohot side36 ofouter liner12. Thus, it will be understood that an angle of diffusion (or included angle)60 exists with respect to anaxis65 extending through each coolinghole44. Angle ofdiffusion60 is defined as an angle extending omni-directionally from a focal point onaxis65 and preferably is in a range of approximately 1° to approximately 15°. A more preferred range fordiffusion angle60 is approximately 3° to approximately 10°, while an optimal range fordiffusion angle60 is approximately 5° to approximately 9°. In any event, it will be appreciated that each coolinghole44 will have a substantially frusto-conical shape.
• It will be appreciated that spacing (represented byreference numeral62 inFIG. 8) between adjacentfirst openings64 and66 of adjacent cooling holes68 and70 is approximately 3.0-6.0 timesfirst diameter54 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 andinner liners12 and14, respectively. Spacing between adjacentsecond openings72 and74 is represented byreference numeral76 inFIG. 9 and preferably is approximately 0.2-0.7 timessecond diameter58 thereof. In order to provide some means of comparison tofirst openings52 oncold sides38 and42 of outer andinner liners12 and14, it will be understood that spacing76 between adjacentsecond openings72 and74 is preferably approximately 2.0-5.0times diameter54 offirst openings64 and66. Because of the shorter spacing between adjacentsecond openings56 of cooling holes44, it will be appreciated that less metal of outer andinner liners12 and14 is provided onhot sides40 and36 thereof is exposed to the harsh environment ofcombustion chamber20. Also, by minimizing the spacing betweensecond openings56 of cooling holes44, the air flowing through cooling holes44 is better able to work in concert to eliminate or minimize hot streaks onhot sides40 and36.
• It will be appreciated that no dilution holes are shown within outer andinner liners12 and14. Nevertheless, dilution air may be introduced intocombustor chamber20 through a plurality of circumferentially spaced dilution holes disposed in each of outer andinner liners12 and14 to promote additional combustion when desired. Such dilution holes would generally be far smaller in number than cooling holes44, with a cross-sectional area that is substantially greater than the cross-sectional area of one of cooling holes44. It will be understood that cooling holes44 will serve to admit some dilution air intocombustor chamber20. Additionally, the disclosed configuration of cooling holes44 is able to enhance bore cooling of outer andinner liners12 and14 since the overall volume thereof has increased.
• As indicated by an arrow75 (seeFIG. 3), it is preferred that cooling air enterfirst opening54 of each coolinghole44 with a predetermined jet velocity on the order of approximately 200-300 feet per second. Due todiffusion angle60 of coolinghole44, whereinsecond opening56 has alarger diameter58 thandiameter54 offirst opening52, cooling air (indicated by arrow85) athot side38 ofouter liner12 has a jet velocity that is approximately 75-100 feet per second. Accordingly, the jet velocity of coolingair85 is less than that for a conventional straight (i.e., uniform diameter) cooling hole. By comparison, the jet velocity of coolingair85 atsecond opening56 is approximately 30%-50% less than the jet velocity of coolingair75 atfirst opening52. This reduction in the jet velocity of coolingair85 alonghot side38 ofouter liner12 assists to promote more effective film cooling and is less apt to penetrate therethrough.
• As shown inFIG. 7, an alternate configuration for cooling holes44 is provided forouter liner12. In this embodiment, each coolinghole144 includes afirst portion146 located adjacentcold side38 ofouter liner12 and asecond portion148 located adjacenthot side36 ofouter liner12. It will be seen thatfirst portion146, which includes afirst opening152, has a substantiallyuniform diameter154 and extends apredetermined length78 fromcold side38 to asecond end80 located within athickness82 ofouter liner12.Second portion148, for its part, extends fromsecond end80 offirst portion146 tosecond opening156 onhot side36 ofouter liner12 so as to have a desiredlength84 and preferably anon-uniform diameter158. While not shown, it will be understood thatsecond portion156 havingnon-uniform diameter158 may be located adjacent tocold side38 ofouter liner12 andfirst portion146 having substantiallyuniform diameter154 may be located adjacent tohot side36 ofouter liner12.
• By configuring the cooling holes in outer andinner liners12 and14 like that described for coolingholes144, the manufacturing of such cooling holes is made less complex. In accordance therewith, a method of forming acooling hole144 in outer andinner liners12 and14 ofcombustor10, where coolinghole144 has a non-uniform diameter therethrough is hereby disclosed. In a first step,second portion148 ofcooling hole144 is formed fromhot side36 ofouter liner12. It will be understood thatsecond portion148 has a diameter150 that progressively decreases fromhot side36 of outer liner and extends a desiredlength84 throughthickness82 ofouter liner12. Thus,second portion148 is substantially conical in shape. Secondly,first portion146 ofcooling hole144 is formed throughsecond portion148 so thatfirst portion146 has a substantially uniform diameter.
• While it is primarily intended for coolingholes44 and/orcooling holes144 to be provided over essentially an entire axial length and circumference of outer andinner liners12 and14, 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 andinner liners12 and14 where hot streaks are known to occur. Exemplary locations for such cooling holes may include adjacent to dilution holes48, adjacent to cooling nuggets present in the liners, immediately downstream of aswirler assembly32, upstream ends13 and17 of the liners, or downstream ends15 and19 of the liners.
• 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.

2. The combustor liner ofclaim 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 ofclaim 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 ofclaim 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 ofclaim 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 ofclaim 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 ofclaim 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 ofclaim 7, wherein each said cooling hole has a substantially frusto-conical shape.

9. The combustor liner ofclaim 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 ofclaim 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 ofclaim 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 ofclaim 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 ofclaim 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 ofclaim 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 ofclaim 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 ofclaim 15, wherein said first portion of said cooling hole is located adjacent said cold side of said shell.

17. The combustor liner ofclaim 15, wherein said first portion of said cooling hole is located adjacent said hot side of said shell.

18. The combustor liner ofclaim 1, wherein said cooling holes are provided over essentially an entire length of said shell.

19. The combustor liner ofclaim 1, wherein said cooling holes are provided at certain designated locations of said shell.

20. The combustor liner ofclaim 1, wherein said cooling holes are provided at an upstream portion of said shell.

21. The combustor liner ofclaim 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 ofclaim 22, wherein said first portion of said cooling hole has a diameter which progressively decreases from said hot side of said liner.

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