FILM STARTERS IN COMBUSTORS OF GAS TURBINE ENGINES

Abstract

A combustor liner, a combustor, and a gas turbine engine are provided. The combustor liner includes a liner portion having a plurality of dilution holes extending through the liner portion. Each dilution hole includes a downstream edge in a direction of a flow of a first fluid through the combustion liner. The downstream edge includes a film cooling starter device. The film cooling starter device includes a plurality of cooling holes extending through the liner portion and spaced apart from the downstream edge in a downstream direction. The film cooling starter device also includes a support flange extending from the downstream edge into the combustion zone, and a lip extending away from the support flange in the downstream direction and spaced a predetermined distance from a surface of the liner portion.

Description

BACKGROUND

The field of the disclosure relates generally to gas turbine engines and, more particularly, to an apparatus for restarting a cooling film along a combustor liner surface for cooling the combustor liner of the gas turbine engine.

At least some known gas turbine engines include a compressor, a combustor, and a turbine in a serial flow arrangement. The compressor compresses air that is channeled into the combustor. The combustor ignites a mixture of the air and a flow of fuel to create high temperature gases which are channeled to the turbine and out of the gas turbine engine to create thrust. The combustor includes a liner to protect the combustor from the high temperature gases. Additionally, a film of cooling air around the liner creates a barrier that protects the combustor from the high temperature gases. The combustor also includes a plurality of dilution holes that provide additional air for combustion. However, the flow exiting the dilution holes disrupts the film of cooling air. Disrupting the film of cooling air may reduce the life span of the combustor.

BRIEF DESCRIPTION

In one aspect, a combustor liner is provided. The combustor liner includes a liner portion and a plurality of dilution holes extending through the liner portion. The dilution holes each include a downstream edge in a direction of a flow of a first fluid through the combustor liner. The downstream edge includes a film cooling starter device. The film cooling starter device includes a plurality of cooling holes extending through the liner portion and spaced apart from the downstream edge in a downstream direction. The film cooling starter device also includes a support flange extending from the downstream edge into the combustion zone, and a lip extending away from the support flange in the downstream direction and spaced a predetermined distance from a surface of the liner portion.

Optionally, the lip extends in the downstream direction over the plurality of cooling holes. The plurality of dilution holes are optionally configured to channel a flow of a second fluid in a second direction substantially perpendicular to the downstream direction. The plurality of cooling holes are optionally configured to channel a flow of a cooling fluid in the second direction. Optionally, the lip is configured to channel the flow of the cooling fluid in a third direction substantially parallel to the downstream direction. Optionally, the lip, the support flange, and combustor liner portion define a film initiation gap configured to channel the flow of the cooling fluid in the third direction. In some embodiments, the flow of the first fluid disrupts a flow of a cooling film along the surface of the liner portion, and the film cooling starter device is optionally configured to restart the cooling film downstream of a respective dilution hole.

In another aspect, a combustor includes a combustor liner including a plurality of dilution holes arranged in one or more fields of dilution holes. Each dilution hole extends through the combustor liner and includes a downstream edge in a downstream direction of a flow of a first fluid through the combustor liner. A support flange extends from the downstream edge into the combustion zone, and a lip extends away from the support flange in the downstream direction and is spaced a predetermined distance from the combustor liner.

Optionally, a plurality of cooling holes extends through the combustor liner and is positioned downstream of a respective dilution hole, and the lip extends in the downstream direction over the plurality of cooling holes. Optionally, the cooling holes may be non-circular cooling holes or circular cooling holes. Also optionally, the dilution holes are configured to channel a flow of a second fluid in a second direction substantially perpendicular to the downstream direction. Also optionally, the plurality of cooling holes is configured to channel a flow of a cooling fluid in the second direction. The lip may be configured to channel the flow of the cooling fluid in the downstream direction. Optionally, the lip, the support flange, and a surface of the combustor liner define a film initiation gap configured to channel the flow of the cooling fluid in the downstream direction along the surface of the combustor liner. In some embodiments, the flow of the second fluid disrupts a flow of a cooling film flowing along the surface of the combustor liner, and the cooling fluid may be directed in the downstream direction along the surface of the combustor liner portion to restart the cooling film downstream of a respective dilution hole. Also optionally, the plurality of cooling holes are oriented at an oblique angle with respect to the downstream direction.

In yet another aspect, a gas turbine engine is provided. The gas turbine engine includes a core engine including a high pressure compressor, a combustor, and a high pressure turbine in a serial flow arrangement. The high pressure compressor and the high pressure turbine are arranged coaxially about an axis of rotation of the gas turbine engine. The combustor includes a combustor liner including a plurality of dilution holes. Each dilution hole extends through the combustor liner and includes a downstream edge in a direction of a flow of a first fluid through the combustor. The downstream edge of at least some of the plurality of dilution holes includes a film cooling starter device. The film cooling starter device includes a support flange extending from the downstream edge into the combustor, and a lip extending away from the support flange in the downstream direction and spaced a predetermined distance from the combustor liner. The film cooling starter device also includes a plurality of cooling holes extending through the combustor liner, and the lip extends over the plurality of cooling holes. Optionally, the plurality of cooling holes are oriented approximately perpendicular to the axis of rotation, or at an oblique angle with respect to the axis of rotation.

BRIEF DESCRIPTION OF THE DRAWINGS

These and other features, aspects, and advantages of the present disclosure 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:

FIGS. 1-8 show example embodiments of the method and apparatus described herein.

FIG. 1 is a perspective view of an aircraft.

FIG. 2 is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure that may be used with the aircraft shown in FIG. 1.

FIG. 3 is a schematic cross-sectional view of a combustor in accordance with an exemplary embodiment of the present disclosure that may be used with the gas turbine engine shown in FIG. 2.

FIG. 4 is a schematic diagram of an aft portion of an outer liner or an inner liner in accordance with an exemplary embodiment of the present disclosure that may be used with the combustor shown in FIG. 3.

FIG. 5 is a schematic diagram of a dilution hole with a first exemplary embodiment of a film cooling starter device for use in the gas turbine engine shown in FIG. 2.

FIG. 6 is a top view of the dilution hole with film cooling starter device shown in FIG. 5.

FIG. 7 is a top view of a dilution hole with a second exemplary embodiment of a film cooling starter device for use in the gas turbine engine shown in FIG. 2.

FIG. 8 is another schematic diagram of a dilution hole with a third exemplary embodiment of a film cooling starter device for use in the gas turbine engine shown in FIG. 2.

Although specific features of various embodiments may be shown in some drawings and not in others, this is for convenience only. Any feature of any drawing may be referenced and/or claimed in combination with any feature of any other drawing.

Unless otherwise indicated, the drawings provided herein are meant to illustrate features of embodiments of the disclosure. These features are believed to be applicable in a wide variety of systems comprising one or more embodiments of the disclosure. As such, the drawings are not meant to include all conventional features known by those of ordinary skill in the art to be required for the practice of the embodiments disclosed herein.

DETAILED DESCRIPTION

In the following specification and the claims, reference will be made to a number of terms, which shall be defined to have the following meanings.

The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise.

“Optional” or “optionally” means that the subsequently described event or circumstance may or may not occur, and that the description includes instances where the event occurs and instances where it does not.

Approximating language, as used herein throughout the specification and claims, may be applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about”, “approximately”, and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value. Here and throughout the specification and claims, range limitations may be combined and/or interchanged; such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise.

The following detailed description illustrates embodiments of the disclosure by way of example and not by way of limitation. It is contemplated that the disclosure has general application to a system for cooling fluids in an aircraft engine.

Embodiments of a film cooling starter device described herein create a cooling film in a combustor in a gas turbine engine. The combustor channels combustion gases in a generally axial direction through the combustor. The combustor includes a liner to protect the combustor from the combustion gases and to extend the life of the combustor. Additionally, an air film is created along a surface of the liner that protects the liner. Dilution holes channel additional air in a direction away from the liner for combustion. However, the flow of air flowing from the dilution holes disrupts the air film along the surface of the liner. In one embodiment, the film cooling starter device is positioned within the dilution hole and restarts the air film downstream of the dilution hole in the direction of flow. In another embodiment, the film cooling starter device includes a plurality of cooling holes positioned downstream of the dilution hole in the direction of flow. The plurality of cooling holes channels a flow of air into the combustor to restart the air film downstream of the dilution hole. However, the plurality of cooling holes channel the flow of air in a direction perpendicular to flow of the air film. Thus, the film cooling starter device also includes a support flange and a lip. The support flange is positioned within the dilution hole and the lip is coupled to the support flange. The lip extends from the support flange in the axial direction and channels the flow of air in the direction of flow of the air film. The film cooling starter device restarts the air film downstream of the dilution hole after the flow from the dilution hole has disrupted the air film. Thus, the film cooling starter device extends the life of the combustor by extending the life of the liner within the combustor.

FIG. 1 is a perspective view of an aircraft 100. In the example embodiment, aircraft 100 includes a fuselage 102 that includes a nose 104, a tail 106, and a hollow, elongate body 108 extending therebetween. Aircraft 100 also includes a wing 110 extending away from fuselage 102 in a lateral direction 112. Wing 110 includes a forward leading edge 114 in a direction 116 of motion of aircraft 100 during normal flight and an aft trailing edge 118 on an opposing edge of wing 110. Aircraft 100 further includes at least one gas turbine engine 120 configured to drive a bladed rotatable member or fan to generate thrust. Gas turbine engine 120 is coupled to at least one of wing 110 and fuselage 102. At least one gas turbine engine 120 is connected to an engine pylon 124, which may connect the at least one gas turbine engine 120 to aircraft 100. Engine pylon 124, for example, may couple at least one gas turbine engine 120 to at least one of wing 110 and fuselage 102, for example, in a pusher configuration proximate tail 106.

FIG. 2 is a schematic cross-sectional view of gas turbine engine 120 in accordance with an exemplary embodiment of the present disclosure. In the example embodiment, gas turbine engine 120 is embodied in a high bypass turbofan jet engine. As shown in FIG. 2, gas turbine engine 120 defines an axial direction A (extending parallel to a longitudinal axis 202 provided for reference) and a radial direction R. In general, gas turbine engine 120 includes a fan assembly 204 and a core engine 206 positioned downstream from fan assembly 204.

In the example embodiment, core engine 206 includes an approximately cylindrical inner casing 208 that defines an annular core engine inlet 220. Inner casing 208 encases, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor 222 and a high pressure (HP) compressor 224; a combustor 226; a turbine section including a high pressure (HP) turbine 228 and a low pressure (LP) turbine 230; and a jet exhaust nozzle 232. A high pressure (HP) spool or shaft 234 drivingly connects HP turbine 228 to HP compressor 224. A low pressure (LP) spool or shaft 236 drivingly connects LP turbine 230 to LP compressor 222. The LP compressor 222, HP compressor 224, combustor 226, HP turbine 228, LP turbine 230, and jet exhaust nozzle 232 together define a core air flowpath 237.

In the example embodiment, fan assembly 204 includes a fan 238, which in some embodiments, has variable pitch features, as shown in FIG. 2. Fan 238 includes a plurality of fan blades 240 coupled to a fan disk 242 in a spaced apart relationship. Although fan assembly 204 is described as including a variable pitch fan 238, fan assembly 204 could include a conventional fixed pitch fan. Fan blades 240 extend radially outwardly from fan disk 242. Each variable-pitch fan blade 240 is rotatable relative to fan disk 242 about a pitch axis P by virtue of fan blades 240 being operatively coupled to a suitable pitch change mechanism (PCM) 244 configured to vary the pitch of fan blades 240. In other embodiments, PCM 244 is configured to collectively vary the pitch of fan blades 240 in unison. Fan blades 240, fan disk 242, PCM 244, and LP compressor 222 are together rotatable about longitudinal axis 202 by LP shaft 236 across a power gear box 246.

Fan disk 242 is covered by rotatable front hub 248, which is aerodynamically contoured to promote an airflow through the plurality of fan blades 240. Additionally, fan assembly 204 and at least a portion of core engine 206 are surrounded by a nacelle assembly 249. Nacelle assembly 249 is a system of components or structures attached to gas turbine engine 120 and/or engine pylon 124, and provides aerodynamic surfaces around gas turbine engine 120. Nacelle assembly 249 may include an annular fan casing or outer nacelle 250 and a core engine cowl or inner nacelle 259 generally separated by a bypass duct 256. An undercowl space 263 of core engine 206 is defined by the volume between inner nacelle 259 and inner casing 208.

Outer nacelle 250 circumferentially surrounds fan 238 and/or at least a portion of core engine 206. More specifically, a downstream section 254 of outer nacelle 250 may extend over a forward portion 261 of inner nacelle 259 so as to define bypass duct 256 therebetween, with outer nacelle 250 providing a radially outer wall for bypass duct 256 and inner nacelle 259 providing a radially inner wall. In the example embodiment, outer nacelle 250 is configured to be supported relative to core engine 206 by a plurality of circumferentially-spaced outlet guide vanes 252.

Nacelle assembly 249 further defines an appropriate inlet opening 260 of fan assembly 204 and outer nacelle 250, defines an appropriate core engine inlet 220 for core air flowpath 237, defines appropriate nozzles for the exhaust of bypass duct 256 and a core exhaust 257, and houses or contains auxiliary devices for the engine and other components for the aircraft including various ducts, lines, pipes and wires.

During operation of gas turbine engine 120, a volume of air 258 enters gas turbine engine 120 through inlet opening 260 of nacelle 250 and/or fan assembly 204. As volume of air 258 passes across fan blades 240, a bypass portion 262 of volume of air 258 is directed or routed into bypass duct 256 and a core engine portion 264 of volume of air 258 is directed or routed into core air flowpath 237, or more specifically into LP compressor 222. A ratio between bypass portion 262 and core engine portion 264 is commonly referred to as a bypass ratio. The pressure of core engine portion 264 is then increased as it is routed through HP compressor 224 and into combustor 226, where it is mixed with fuel and burned to provide combustion gases 266.

Combustion gases 266 are routed through HP turbine 228 where a portion of thermal and/or kinetic energy from combustion gases 266 is extracted via sequential stages of HP turbine stator vanes 268 that are coupled to inner casing 208 and HP turbine rotor blades 270 that are coupled to HP shaft 234, thus causing HP shaft 234 to rotate, which then drives a rotation of HP compressor 224. Combustion gases 266 are then routed through LP turbine 230 where a second portion of thermal and kinetic energy is extracted from combustion gases 266 via sequential stages of LP turbine stator vanes 272 that are coupled to inner casing 208 and LP turbine rotor blades 274 that are coupled to LP shaft 236, which drives a rotation of LP shaft 236, LP compressor 222, and rotation of fan 238 across power gear box 246.

Combustion gases 266 are subsequently routed through jet exhaust nozzle 232 of core engine 206 to provide propulsive thrust. Simultaneously, the pressure of bypass portion 262 is substantially increased as bypass portion 262 is routed through bypass duct 256 before it is exhausted from a fan nozzle exhaust 276 of gas turbine engine 120, also providing propulsive thrust. HP turbine 228, LP turbine 230, and jet exhaust nozzle 232 at least partially define a hot gas path 278 for routing combustion gases 266 through core engine 206.

Exemplary gas turbine engine 120 depicted in FIG. 2 is by way of example only, and in other embodiments, gas turbine engine 120 may have any other suitable configuration. It should also be appreciated, that in still other embodiments, aspects of the present disclosure may be incorporated into any other suitable gas turbine engine. For example, in other embodiments, aspects of the present disclosure may be incorporated into, e.g., a turboprop engine, turboshaft engine, and military style engine.

FIG. 3 is an exemplary embodiment of combustor 226. In the example embodiment, combustor 226 includes a combustion zone 318 defined between and by an annular radially outer liner 320 and an annular radially inner liner 322, respectively circumscribed about an engine centerline 352. In the exemplary embodiment, combustor 226 includes a rich burn combustor. However, combustor 226 additionally or alternatively includes any type of combustor that enables gas turbine engine 120 to operate as described herein, including, without limitation, TAPS combustors. The outer and inner liners 320, 322 are located radially inwardly of an annular combustor casing 326 which extends circumferentially around outer and inner liners 320, 322. The combustor 226 also includes an annular dome 334 mounted upstream of combustion zone 318 and attached to the outer and inner liners 320, 322. Outer and inner liners 320, 322 are supported at an upstream end by dome 334 or dome support structure, and a downstream end of outer and inner liners 320, 322 is supported by an aft outer liner support 321 or by an aft inner liner support 323. Dome 334 defines an upstream end 336 of combustion zone 318 and a plurality of fuel/air injectors 340, such as, but not limited to fuel nozzles or mixer assemblies (only one is illustrated) are spaced circumferentially around dome 334. In some embodiments, fuel/air injectors 340 are embodied as mixer assemblies that may include a main mixer 394 mounted in the dome 334 and a pilot mixer 396. In other embodiments, fuel/air injectors 340 are embodied as primary and secondary swirlers, such as, in a rich-burn combustor. Outer and inner liners 320, 322 include a plurality of dilution holes 412 (shown in FIG. 4) that extend through outer and inner liners 320, 322.

Combustor 226 receives an annular stream of pressurized compressor discharge air 314 from a high pressure compressor discharge outlet 369, referred to as CDP air (compressor discharge pressure air). A premix portion 319 of t CDP air 314 flows into fuel/air injector 340, where fuel is also injected to mix with the air and form a fuel-air mixture 365 that is provided to combustion zone 318 for combustion. A fuel injector 310 includes a nozzle mount or flange 330 adapted to be fixed and sealed to the combustor casing 326. A hollow stem 332 of fuel injector 310 is integral with or fixed to flange 330 (such as by brazing or welding) and includes a fuel nozzle assembly 312. Fuel and air are provided to fuel/air injectors 340 so that a primary combustion zone 398 is maintained within a central portion of combustion zone 318.

In the exemplary embodiment, fuel-air mixture 365 is a rich mixture of fuel and air. That is, fuel-air mixture 365 includes more fuel than air. In the exemplary embodiment, fuel-air mixture 365 includes a ratio of fuel to air that is approximately 1.8. However, fuel-air mixture 365 may include any ratio of fuel to air that enables combustor 226 to operate as described herein. Fuel/air injector 340 atomizes fuel within CDP air 314. Ignition of the fuel-air mixture 365 is accomplished by an ignitor 370, and the resulting rich burn combustion gases 364 are rich in carbon monoxide and partially oxidized hydrocarbon species. Rich burn combustion gases 364 flow in axial direction A toward an aft portion of combustor 226.

The arrows in FIG. 3 illustrate the directions in which compressor discharge air flows within combustor 226. An outer flow portion 324 of CDP air 314 flows around the outer liner 320 and inner flow portion 325 of CDP air 314 flows around the inner liner 322. Dilution holes 412 (shown in FIG. 4) channel outer flow portion 324 and inner flow portion 325 into combustion zone 318. The additional CDP air 314 from dilution holes 412 reduces the ratio of fuel to air within combustion zone 318. The resultant mixture is combusted to produce a flow of resulting combustion gases 360 that includes a reduced content of carbon monoxide and partially oxidized hydrocarbon species. Resulting combustion gases 360 flow in axial direction A toward and into an annular, first stage turbine nozzle 372. The first stage turbine nozzle 372 is defined by an annular flow channel that includes a plurality of radially extending, circularly-spaced nozzle vanes 374 that turn the gases so that they flow angularly and impinge upon the first stage turbine blades of a first turbine, such as rotor blades 270 of HP turbine 228 (shown in FIG. 2).

FIG. 4 is a schematic diagram of a liner portion 400 of outer liner 320 or inner liner 322 (shown in FIG. 3) in accordance with an example embodiment of the present disclosure. With reference to FIGS. 3 and 4, in the example embodiment, liner portion 400 includes a surface 401 that is exposed to combustion zone 318 during operation. Liner portion 400 also includes a length 402 in axial direction A and a width 404 in a circumferential direction 405. Because of the changing cross-section of liner portion 400 from upstream end 336 to first stage turbine nozzle 372, width 404 may be of different values at an upstream end 406 of liner portion 400 and a downstream end 408 of liner portion 400. In the example embodiment, upstream end 406 is couplable to dome 334 or to structure coupled to dome 334. In other embodiments, liner portion 400 is couplable to an upstream portion of outer liner 320 or inner liner 322. Also in the example embodiment, downstream end 408 is couplable to a downstream portion of outer liner 320 or inner liner 322 and in other embodiments, downstream end 408 is couplable to aft outer liner support 321 or to aft inner liner support 323. Liner portion 400 is joined to other similar liner portions to form outer liner 320 or inner liner 322.

Liner portion 400 includes a field 410 of a plurality of dilution holes 412 and a plurality of cooling holes 414 that extend through liner portion 400 at specific predetermined locations in a multihole pattern. In FIG. 4, only two rows of field 410 are shown. In the exemplary embodiment, a plurality of cooling holes 414 are positioned downstream of each dilution hole 412 in axial direction A. However, any number of cooling holes 414 may be positioned aft of each dilution hole 412 which enable combustor 226 to operate as described herein. Dilution holes 412 and cooling holes 414 extend through liner portion 400 so that outer flow portion 324 or inner flow portion 325 of CDP air 314 flows through dilution holes 412 and cooling holes 414. Dilution holes 412 and cooling holes 414 are shown as approximately circular holes, however, in other embodiments, dilution holes 412 and cooling holes 414 can be any shape and may include diffusers formed in an outlet of the dilution holes 412 and cooling holes 414. The dilution holes 412 and cooling holes 414 include a downstream edge 416 on an aft or downstream side of dilution holes 412 and cooling holes 414. In the exemplary embodiment, dilution holes 412 channel outer flow portion 324 or inner flow portion 325 of the CDP air 314 into combustor 226 to provide additional air for combustion in a rich-burn combustor. Cooling holes 414 channel outer flow portion 324 or inner flow portion 325 of the CDP air 314 into combustor 226 to form a cooling film along surface 401. The arrangement of cooling holes 414 is configured to restart the cooling film disrupted by dilution holes 414. The cooling film is a laminar flow of cooling fluid that is generated by outer flow portion 324 or inner flow portion 325 of the CDP air 314 exiting cooling holes 414 and being directed by a film cooling starter device 500, such as film cooling starter device 500, 700, and 800 (see FIGS. 5-8), at a predetermined angle with respect to surface 401. In some locations, dilution holes 412 cause a disruption in the laminar flow of the cooling film. In some cases the cooling film may be disrupted such that surface 401 is exposed to the harsh environment of combustion zone 318. To restart the cooling film, cooling holes 414 are positioned in a predetermined location to restart the cooling film downstream of dilution holes 412. Cooling holes 414 are positioned proximate to dilution holes 412 and downstream of dilution holes 412. As described below in greater detail, dilution holes 412 are shaped differently than cooling holes 414. For example, dilution holes 412 are larger than cooling holes 414 to allow more flow of outer flow portion 324 or inner flow portion 325 of the CDP air 314.

FIG. 5 is a schematic diagram of dilution hole 412 with a first exemplary embodiment of a film cooling starter device 500. FIG. 6 is a top view of dilution hole 412 with film cooling starter device 500. Film cooling starter device 500 includes a support flange 502 and a lip 504. Support flange 502 is positioned within dilution hole 412 and coupled to a rim 506 of dilution hole 412. Support flange 502 extends past rim 506 into combustion zone 318. Lip 504 is coupled to support flange 502 and extends from support flange 502 in axial direction A within combustion zone 318, such that lip 504 is spaced a predetermined distance 503 from surface 401. Lip 504 extends in axial direction A over cooling holes 414. Support flange 502, lip 504, and liner portion 400 define a film initiation gap 508 aft of dilution hole 412 in axial direction A.

During operations, a flow of a first fluid is channeled through combustion zone 318 or hot gas path 278 (shown in FIG. 2) in a downstream direction 510. In the exemplary embodiment, the first fluid is combustion gases 266 (shown in FIG. 2) and downstream direction 510 is through combustion zone 318 in axial direction A. A flow of a second fluid is channeled through dilution hole 412 in a second direction 512. In the exemplary embodiment, the second fluid is outer flow portion 324 or inner flow portion 325 of the CDP air 314 and second direction 512 is into combustion zone 318 in radial direction R substantially perpendicular to downstream direction 510. A flow of a third fluid is channeled through cooling holes 414 in a third direction 514. In the exemplary embodiment, the third fluid is outer flow portion 324 or inner flow portion 325 of the CDP air 314 and third direction 514 is into combustion zone 318 in radial direction R substantially perpendicular to downstream direction 510.

A flow of a cooling film 516 is developed within combustor 226 substantially parallel to liner portion 400. Dilution hole 412 and flow of second fluid 512 disrupt cooling film 516. To restart cooling film 516, the flow of the third fluid is channeled into cooling holes 414 and into film initiation gap 508. Lip 504 turns flow of third fluid in a fourth direction 518 parallel to downstream direction 510. Fourth direction 518 is substantially parallel to surface 401 and restarts cooling film 516 aft of dilution hole 412.

In the exemplary embodiment, seven cooling holes 414 are positioned aft of each dilution hole 412 in axial direction A. However, any number of cooling holes 414 may be positioned aft of each dilution hole 412 which enable combustor 226 to operate as described herein. Cooling holes 414 include a circular shape and extend perpendicularly through liner portion 400.

FIG. 7 is a top view of dilution hole 412 with another embodiment of a film cooling starter device, designated film cooling starter device 700. Film cooling starter device 700 is the same as film cooling starter device 500 except film cooling starter device 700 includes a plurality of cooling holes 714. Cooling holes 714 include a rectangular shape to define a slot extending perpendicularly through liner portion 400. In the exemplary embodiment, three cooling holes 714 are positioned aft of each dilution hole 412 in axial direction A. However, any number of cooling holes 714 may be positioned aft of each dilution hole 412 which enable combustor 226 to operate as described herein.

FIG. 8 is a schematic diagram of dilution hole 412 with another embodiment of a film cooling starter device, designated film cooling starter device 800. Film cooling starter device 800 is the same as film cooling starter device 500 except film cooling starter device 800 includes a plurality of cooling holes 814. Cooling holes 814 include a circular shape and extend at an oblique angle through liner portion 400 such that a third fluid is channeled through cooling holes 814 in a third direction 802. In the exemplary embodiment, the third fluid is outer flow portion 324 or inner flow portion 325 of CDP air 314 and the third direction is into combustion zone 318 at an oblique angle relative to radial direction R and downstream direction 510. Obliquely angled cooling holes 814 provide additional impingement back side cooling to the corner of film cooling starter device 800. Also, obliquely angled cooling holes 814 expand air quicker circumferentially, which facilitates reducing hot gas ingested into the lip flow path.

Embodiments of the above-described film cooling starter device provide an efficient method for restarting a cooling film within a combustor. The film cooling starter device is positioned within a dilution hole and restarts the cooling film downstream of the dilution hole. The film cooling starter device includes a plurality of cooling holes positioned downstream of the dilution hole. The cooling holes channel a flow of air into the combustor to restart the air film downstream of the dilution hole. The film cooling starter device includes a lip that directs the flow of air from the cooling holes parallel to the film cooling flow, restarting the air film downstream of the dilution hole after the dilution hole has disrupted the air film. Thus, the film cooling starter device extends the life of the combustor by extending the life of the liner within the combustor.

Exemplary embodiments of a film cooling starter device are described above in detail. The film cooling starter device, and methods of operating such systems and devices, are not limited to the specific embodiments described herein, but rather, components of systems and/or steps of the methods may be utilized independently and separately from other components and/or steps described herein. For example, the methods may also be used in combination with other systems requiring a cooling film, and are not limited to practice with only the systems and methods as described herein. Rather, the exemplary embodiment can be implemented and utilized in connection with many other machinery applications that are currently configured to receive and accept a cooling film.

Example methods and apparatus for starting a film of cooling air are described above in detail. The apparatus illustrated is not limited to the specific embodiments described herein, but rather, components of each may be utilized independently and separately from other components described herein. Each system component can also be used in combination with other system components.

This written description uses examples to describe the disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the disclosure 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.

Claims

1. A combustor liner at least partially surrounding a combustion zone, said combustor liner comprising:

a liner portion;

a plurality of dilution holes extending through said liner portion, each of said dilution holes including a downstream edge in a direction of a flow of a first fluid through said combustor liner, said downstream edge including a film cooling starter device comprising:

a plurality of cooling holes extending through said liner portion and spaced apart from said downstream edge in a downstream direction;

a support flange extending from said downstream edge into the combustion zone; and

a lip extending away from said support flange in the downstream direction and spaced a predetermined distance from a surface of said liner portion.

2. The combustor liner of claim 1, wherein said lip extends in the downstream direction over said plurality of cooling holes.

3. The combustor liner of claim 2, wherein said plurality of dilution holes are configured to channel a flow of a second fluid in a second direction substantially perpendicular to the downstream direction.

4. The combustor liner of claim 3, wherein said plurality of cooling holes is configured to channel a flow of a cooling fluid in the second direction.

5. The combustor liner of claim 4, wherein said lip is configured to channel the flow of the cooling fluid in a third direction substantially parallel to the downstream direction.

6. The combustor liner of claim 5, wherein said lip, said support flange, and combustor liner portion define a film initiation gap configured to channel the flow of the cooling fluid in the third direction.

7. The combustor liner of claim 6, wherein the flow of the second fluid disrupts a flow of a cooling film along said surface of said liner portion, said film cooling starter device is configured to restart the flow of the cooling film downstream of a respective dilution hole.

8. A combustor comprising:

a combustor liner surrounding a combustion zone and comprising a plurality of dilution holes arranged in one or more fields of said dilution holes, each of said dilution holes extending through said combustor liner, each said dilution hole including a downstream edge in a downstream direction of a flow of a first fluid through said combustor liner; a support flange extending from said downstream edge into said combustion zone; and a lip extending away from said support flange in the downstream direction and spaced a predetermined distance from said combustor liner.

9. The combustor of claim 8, further comprising a plurality of cooling holes extending through said combustor liner, said plurality of cooling holes positioned downstream of a respective dilution hole of said dilution holes, wherein said lip extends in the downstream direction over said plurality of cooling holes.

10. The combustor of claim 9, wherein said plurality of cooling holes comprises a plurality of non-circular cooling holes.

11. The combustor of claim 9, wherein said plurality of cooling holes comprises a plurality of circular cooling holes.

12. The combustor of claim 9, wherein said dilution holes are configured to channel a flow of a second fluid in a second direction substantially perpendicular to the downstream direction.

13. The combustor of claim 12, wherein said plurality of cooling holes is configured to channel a flow of a cooling fluid in the second direction.

14. The combustor of claim 13, wherein said lip is configured to channel the flow of the cooling fluid in the downstream direction.

15. The combustor of claim 14, wherein said lip, said support flange, and a surface of said combustor liner define a film initiation gap configured to channel the flow of the cooling fluid in the downstream direction along said surface of said combustor liner.

16. The combustor of claim 15, wherein the flow of the second fluid disrupts a flow of a cooling film flowing along said surface of said combustor liner, wherein said cooling fluid directed in the downstream direction along said surface of said combustor liner is configured to restart the flow of the cooling film downstream of a respective dilution hole.

17. The combustor of claim 9, wherein said plurality of cooling holes are oriented at an oblique angle with respect to the downstream direction.

18. A gas turbine engine comprising:

a core engine comprising a high pressure compressor, a combustor, and a high pressure turbine in a serial flow arrangement, said high pressure compressor and said high pressure turbine coaxial about an axis of rotation, said combustor comprising: a combustor liner comprising a plurality of dilution holes, each of said dilution holes extending through said combustor liner, each of said plurality of dilution holes including a downstream edge in a direction of a flow of a first fluid through said combustor, said downstream edge of at least some of said plurality of dilution holes including a film cooling starter device comprising: a support flange extending from said downstream edge into said combustor; a lip extending away from said support flange in the downstream direction and spaced a predetermined distance from said combustor liner; and a plurality of cooling holes extending through said combustor liner, said lip extending in the downstream direction over said plurality of cooling holes.

19. The gas turbine engine of claim 18, said plurality of cooling holes are oriented approximately perpendicular to the axis of rotation.

20. The gas turbine engine of claim 18, said plurality of cooling holes are oriented at an oblique angle with respect to the axis of rotation.