Patent Application
20180283689
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.
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.
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:
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.
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.
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
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
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
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
During operations, a flow of a first fluid is channeled through combustion zone 318 or hot gas path 278 (shown in
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.
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.
