Patent Application
20180274370
Turbine engines, and particularly gas or combustion turbine engines, are rotary engines that extract energy from a flow of combusted gases passing through the engine onto a multitude of turbine blades. Gas turbine engines have been used for land and nautical locomotion and power generation, but are most commonly used for aeronautical applications such as for aircraft, including helicopters. In aircraft, gas turbine engines are used for propulsion of the aircraft. In terrestrial applications, turbine engines are often used for power generation.
Gas turbine engines for aircraft are designed to operate at high temperatures to maximize engine efficiency, so cooling of certain engine components, such as the high pressure turbine and the low pressure turbine, may be necessary. Typically, cooling is accomplished by ducting cooler air from the high and/or low pressure compressors to the engine components which require cooling. Temperatures in the high pressure turbine are around 1000° C. to 2000° C. and the cooling air from the compressor is around 500° C. to 700° C. While the compressor air is a high temperature, it is cooler relative to the turbine air, and may be used to cool the turbine.
Particles, such as dirt, dust, sand, and other environmental contaminants, in the cooling air can cause a loss of cooling and reduced operational time or “time-on-wing” for the aircraft environment. For example, particles supplied to the turbine components can clog, obstruct, or coat the flow passages and surfaces of the components, which can reduce the lifespan of the turbine. In particular, particles can coat and block the film holes present in components. This problem is exacerbated in certain operating environments around the globe where turbine engines are exposed to significant amounts of airborne particles.
In one aspect, the technology described herein relates to an engine component for a gas turbine engine generating hot combustion gas, the engine component having a wall at least partially defining an interior cavity and separating the hot combustion gas from a cooling fluid flow supplied to the interior cavity and having a hot surface facing the hot combustion gas and a cooling surface facing the cooling fluid flow, and a film hole having an inlet provided on the cooling surface, an outlet provided on the hot surface, and a passage connecting the inlet and the outlet, with the passage defining a metering section, wherein the inlet comprises a flared portion flaring inwardly from the cooling surface and about the entire circumference of the inlet.
In another aspect, the technology described herein relates to an engine component for a gas turbine engine generating hot combustion gas, the engine component having a wall at least partially defining an interior cavity and separating the hot combustion gas from a cooling fluid flow supplied to the interior cavity and having a hot surface facing the hot combustion gas and a cooling surface facing the cooling fluid flow, and a film hole having an inlet provided on the cooling surface, an outlet provided on the hot surface, and a passage connecting the inlet and the outlet, with the passage defining a metering section, wherein the inlet comprises at least one flute extending inwardly from the cooling surface to the passage.
In yet another aspect, the technology described herein relates to an engine component for a gas turbine engine generating hot combustion gas. The engine component includes a wall separating the hot combustion gas from a cooling fluid flow and having a hot surface facing the hot combustion gas and a cooling surface facing the cooling fluid flow, multiple film holes having an inlet provided on the cooling surface, an outlet provided on the hot surface, and a passage connecting the inlet and the outlet, with the passage defining a metering section, and a contoured portion provided in the cooling surface and encompassing the inlets for at least two of the film holes.
In the drawings:
The described embodiments of the technology described herein are directed to a film-cooled engine component, particularly in a gas turbine engine. For purposes of illustration, the technology described herein will be described with respect to an aircraft gas turbine engine. It will be understood, however, that the technology described herein is not so limited and may have general applicability in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications.
As used herein, the terms “axial” or “axially” refer to a dimension along a longitudinal axis of an engine. The term “forward” used in conjunction with “axial” or “axially” refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component. The term “aft” used in conjunction with “axial” or “axially” refers to a direction toward the rear or outlet of the engine relative to the engine centerline.
As used herein, the terms “radial” or “radially” refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference. The use of the terms “proximal” or “proximally,” either by themselves or in conjunction with the terms “radial” or “radially,” refers to moving in a direction toward the center longitudinal axis, or a component being relatively closer to the center longitudinal axis as compared to another component. The use of the terms “distal” or “distally,” either by themselves or in conjunction with the terms “radial” or “radially,” refers to moving in a direction toward the outer engine circumference, or a component being relatively closer to the outer engine circumference as compared to another component.
All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise) are only used for identification purposes to aid the reader's understanding of the present invention, and do not create limitations, particularly as to the position, orientation, or use of the invention. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and may include intermediate members between a collection of elements and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and in fixed relation to each other. The exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the drawings attached hereto may vary.
The fan section 18 includes a fan casing 40 surrounding the fan 20. The fan 20 includes a plurality of fan blades 42 disposed radially about the centerline 12.
The HP compressor 26, the combustor 30, and the HP turbine 34 form a core 44 of the engine 10 which generates combustion gases. The core 44 is surrounded by core casing 46 which can be coupled with the fan casing 40.
A HP shaft or spool 48 disposed coaxially about the centerline 12 of the engine 10 drivingly connects the HP turbine 34 to the HP compressor 26. A LP shaft or spool 50, which is disposed coaxially about the centerline 12 of the engine 10 within the larger diameter annular HP spool 48, drivingly connects the LP turbine 36 to the LP compressor 24 and fan 20.
The LP compressor 24 and the HP compressor 26 respectively include a plurality of compressor stages 52, 54, in which a set of compressor blades 56, 58 rotate relative to a corresponding set of static compressor vanes 60, 62 (also called a nozzle) to compress or pressurize the stream of fluid passing through the stage. In a single compressor stage 52, 54, multiple compressor blades 56, 58 may be provided in a ring and may extend radially outwardly relative to the centerline 12, from a blade platform to a blade tip, while the corresponding static compressor vanes 60, 62 are positioned downstream of and adjacent to the rotating blades 56, 58. It is noted that the number of blades, vanes, and compressor stages shown in
The HP turbine 34 and the LP turbine 36 respectively include a plurality of turbine stages 64, 66, in which a set of turbine blades 68, 70 are rotated relative to a corresponding set of static turbine vanes 72, 74 (also called a nozzle) to extract energy from the stream of fluid passing through the stage. In a single turbine stage 64, 66, multiple turbine blades 68, 70 may be provided in a ring and may extend radially outwardly relative to the centerline 12, from a blade platform to a blade tip, while the corresponding static turbine vanes 72, 74 are positioned upstream of and adjacent to the rotating blades 68, 70. It is noted that the number of blades, vanes, and turbine stages shown in
In operation, the rotating fan 20 supplies ambient air to the LP compressor 24, which then supplies pressurized ambient air to the HP compressor 26, which further pressurizes the ambient air. The pressurized air from the HP compressor 26 is mixed with fuel in combustor 30 and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine 34, which drives the HP compressor 26. The combustion gases are discharged into the LP turbine 36, which extracts additional work to drive the LP compressor 24, and the exhaust gas is ultimately discharged from the engine 10 via the exhaust section 38. The driving of the LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP compressor 24.
Some of the ambient air supplied by the fan 20 may bypass the engine core 44 and be used for cooling of portions, especially hot portions, of the engine 10, and/or used to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are normally downstream of the combustor 30, especially the turbine section 32, with the HP turbine 34 being the hottest portion as it is directly downstream of the combustion section 28. Other sources of cooling fluid may be, but is not limited to, fluid discharged from the LP compressor 24 or the HP compressor 26.
One or more of the engine components of the engine 10 has a film-cooled wall in which various film hole embodiments disclosed further herein may be utilized. Some non-limiting examples of the engine component having a film-cooled wall can include the blades 68, 70, vanes or nozzles 72, 74, combustor deflector 76, combustor liner 77, or shroud assembly 78, described in
The engine component 80 includes at least one wall 82 having a hot surface 84 facing the hot combustion gas and a cooling surface 86 facing cooling fluid. In the case of a gas turbine engine, the hot surface 84 may be exposed to gases having temperatures in the range of 1000° C. to 2000° C. Suitable materials for the wall 82 include, but are not limited to, steel, refractory metals such as titanium, or super alloys based on nickel, cobalt, or iron, and ceramic matrix composites.
In the illustrated embodiment, a second wall 87 of the engine component 80 is shown, which, together with the first wall 82, defines at least one interior cavity 88, which comprises the cooling surface 86. The hot surface 84 may be an exterior surface of the engine component 80.
The engine component 80 further includes multiple film holes 90 that provide fluid communication between the interior cavity 88 and the hot surface 84 of the engine component 80. During operation, cooling air C is supplied to the interior cavity 88 and out of the film holes 90 to create a thin layer or film of cool air on the hot surface 84, protecting it from the hot combustion gas H.
Each film hole 90 can have an inlet 92 provided on the cooling surface 86 of the wall 82, an outlet 94 provided on the hot surface 84, and a passage 96 connecting the inlet 92 and outlet 94. Cooling fluid C enters the film hole 90 through the inlet 92 and passes through the passage 96 before exiting the film hole 90 at the outlet 94 along the hot surface 84.
The passage 96 can further define a metering section 98 for metering of the mass flow rate of the cooling fluid C. The metering section 98 can be a portion of the passage 96 with the smallest cross-sectional area, and may be a discrete location or an elongated section of the passage 96.
The present invention provides for a shaping or contouring of the film hole 90 by providing the inlet 92 with a flared portion 100 that flares inwardly from the cooling surface 86 about the entire circumference of the inlet 92. As used herein, the term “flared” and variations thereof, is defined as gradually becoming wider at one end. Here, the flared portion 100 is wider at the cooling surface 86 and narrows gradually in the downstream direction of the passage 96. The metering section 98 can be provided at or near the downstream end of the flared portion 100. In operation, cooling fluid C enters the film hole 90 through the inlet 92 and passes sequentially through the flared portion 100 and the metering section 98 before exiting the film hole 90 at the outlet 94 along the hot surface 84.
The flared portion 100 can be continuous about the circumference of the inlet 92 and can converge with the cooling surface 86 to create a radiused edge 102 at the cooling surface 86. The radiused edge 102 can define a maximum cross-sectional area of the flared portion 100, and can comprise a series of curved and/or linear segments defining the shape of the inlet 92 in the cooling surface 86.
The flared portion 100 can comprise one or more discrete lobes or flutes 104 around the otherwise circular or oblong cross section of the inlet 92. The flutes 104 may vary in shape and size along their length, but are generally largest near the inlet 92 and taper down to disappear at some point interior to the inlet 92 within the passage 96. The flutes 104 may be shaped in various ways, including being arcuate, multiply curved, straight, or piecewise linear.
The shaped inlets 92 can be configured to mitigate the effect that particles within the cooling fluid flow C have on cooling of the engine component 80, either through improved particle collection or improved particle flow. For example, the inlets 92 can be shaped to allow for expanded local flow areas in which particles may collect without affecting the metering section 98 of the film hole 90. In another example, when placed in particular locations around the periphery of the inlet 92, the flutes 104 provide a more gradual transition of the coolant flow turning into the film hole 90. This will allow the particles to be better retained in the fluid, rather than depositing on the engine component 80.
It is noted that a streamline of the cooling fluid flow C may be generally collinear with the centerline 106 of the film hole 90 in areas where the passage 96 is circular or otherwise symmetrical. In areas where the passage 96 is irregular or asymmetrical, the streamline may diverge from the centerline 106.
The flared portion 100 converges toward the centerline 106 in the downstream direction. The cross-sectional area A of the metering section 98, defined with respect to a plane perpendicular to the centerline 106, may remain substantially constant between the flared portion 100 and the outlet 94. In other embodiments, the metering section 98 can have a cross-sectional area A that decreases toward the outlet 94, with the cross section being defined by a plane perpendicular to the centerline 106. In the illustrated embodiment, the metering section 98 further has the same cross-sectional shape as the outlet 94, which is circular in a plane perpendicular to the centerline 106. As the film hole 90 is inclined however, the outlet 94 can have an oval or elliptical plan form when viewed from the hot surface 84. Further, the inlet 92 and the flared portion 100 are greater in cross-sectional area than the metering section 98 and outlet 94.
The flutes 104 comprise concave recesses 112 with a convexly bowed ends 114 defined at the cooling surface 86. The concave recesses 112 can extend in a downstream direction to converge with or disappears into an inner surface 116 of the passage 96 at a distal end 118. Adjacent flutes 104 are contiguous with each other and are separated by ridges 120, which also converge with or disappear into the passage 96 at the downstream end. The recesses 112 can have a generally circular, ovoid, elliptical in shape, or a combination thereof, when viewed in cross-section. The bowed ends 114 can be generally circular, ovoid, or elliptical in shape, or combination thereof. The ridges 120 shown are created by the meeting of adjacent concave recesses 112, and may be a sharp edge or a smoothly radiused structure such as a convex edge.
The flared portion 130 is continuous about the circumference of the inlet 92 and can converge with the cooling surface 86 to create a radiused edge 132 at the cooling surface 86. The radiused edge 132 shown is non-constant about the circumference of the inlet 92, and generally includes an upstream edge segment 134 on the upstream side of the inlet 92 and a downstream edge segment 136 on the downstream side of the inlet 92. The segments 134. 136 each generally define a radius of curvature 138, 140, respectively, with the radius of curvature 138, 140 for the edge segments 134, 136 being measured for a circular arc running along the segment 134, 136. In the illustrated embodiment, the radius of curvature 140 for the downstream edge segment 136 is less than the radius of curvature 138 of the upstream edge segment 134. In general terms, the inlet 92 is more tightly curved along its downstream side than its upstream side.
Furthermore, in any of the above embodiments, a protective coating, such as a thermal barrier coating or multi-layer coating system, can be applied to the hot surface 84 of the engine component 80. Also, the present invention may be combined with shaping or contouring of the passage or outlet of the film holes. For example, the passage 96 can further define a diffusing section in which the cooling fluid C may expand to form a wider cooling film. The diffusion section can be downstream of the metering section 98 and defined at or near the outlet 94. The present invention may also apply to slot-type film cooling, in which case the outlets 94 are provided within a slot on the hot surface 84.
The various embodiments of systems, methods, and other devices related to the invention disclosed herein provide improved cooling for engine structures, particularly in a turbine component having film holes. One advantage that may be realized in the practice of some embodiments of the described systems is that the film hole can be shaped to include a flared or fluted inlet. Conventional film hole design utilizes a passage with a circular inlet region, a metering section, and a shaped outlet region to help diffuse the cooling fluid. By shaping the film hole to include a fluted inlet, improved cooling performance and mitigation of particle buildup in the engine component is achievable, which can lead to longer service life of the engine component.
Another advantage that may be realized in the practice of some embodiments of the described systems and methods is that the flutes allow for expanded local flow areas in which particulates may collect without affecting the metering section of the film hole. When placed in particular locations around the inlet periphery, such as upstream, downstream, or between these, the flutes provide a more gradual transition of the coolant flow turning into the film hole. This will allow particles to be better retained in the fluid flow, rather than depositing on the surfaces of the engine component.
The engine component 80 includes a wall 82 having a hot surface 84 facing the hot combustion gas and a cooling surface 86 facing cooling fluid. In the case of a gas turbine engine, the hot surface 84 may be exposed to gases having temperatures in the range of 1000° C. to 2000° C. Suitable materials for the wall 82 include, but are not limited to, steel, refractory metals such as titanium, or super alloys based on nickel, cobalt, or iron, and ceramic matrix composites.
The engine component 80 can define at least one interior cavity 88 comprising the cooling surface 86. The hot surface 84 may be an exterior surface of the engine component 80.
The engine component 80 further includes multiple film holes 90 that provide fluid communication between the interior cavity 88 and the hot surface 84 of the engine component 80. During operation, cooling air C is supplied to the interior cavity 88 and out of the film holes 90 to create a thin layer or film of cool air on the hot surface 84, protecting it from the hot combustion gas H.
Each film hole 90 can have an inlet 92 provided on the cooling surface 86 of the wall 82, an outlet 94 provided on the hot surface 84, and a passage 96 connecting the inlet 92 and outlet 94. Cooling fluid C enters the film hole 90 through the inlet 92 and passes through the passage 96 before exiting the film hole 90 at the outlet 94 along the hot surface 84.
The passage 96 can define a metering section for metering of the mass flow rate of the cooling fluid C. The metering section can be a portion of the passage 96 with the smallest cross-sectional area, and may be a discrete location or an elongated section of the passage 96. The passage 96 can further define a diffusing section in which the cooling fluid C may expand to form a wider cooling film. The diffusion section has a larger cross-sectional area of than the metering section. The metering section can be provided at or near the inlet 92, while the diffusion section can be defined at or near the outlet 94.
The present invention provides for a shaping or contouring of the cooling surface 86 of the engine component 80 by providing the cooling surface 86 with a contoured portion 98 that encompasses the inlets 92 of two or more film holes 90. Rather than shaping the entry region to each individual film hole 90, which may introduce undesirable stress increases locally, two or more film holes 90 adjacent one another have their entry regions, or inlets 92, encompassed within a broader surface contoured portion 98 to transition the flow more smoothly into the group of film holes 90. Likewise, such contouring may also serve other desirable local purposes, such as the provision of flow deflection as the fluid approaches the inlets 92 to divert particles from entering the film holes 90, or to prevent impact of particles on inlet surfaces, or to provide a more beneficial flow entry angle to the film holes 90.
The contoured portion 98 can encompass the inlets 92 of a partial row of film holes 90, or an entire row of film holes 90, whether that row be considered in a radial or axial direction, or otherwise oriented on the engine component 80. As shown in
The contoured portion 98 of
A single engine component 80 can be provided with one or more of the profiles shown in
The ledges 106 upstream of the inlets 92 can define particle deflectors that prevent or at least reduce the number of particles that enter the film holes 90 by deflecting the particles away from the inlets 92. The height of the ledges 106 relative to the inlet 92 and/or the distance from the ledges 106 to the inlet 92 can be configured based on the expected size and speed of the particles in the cooling fluid flow C. Since the cooling fluid flow C is generally along a channel direction or has a main local direction and momentum as indicated by the arrows in
The contoured portion 98 may further include smaller discrete features around the inlets 92, in addition to the broader feature of the channel. For example,
The curved portion 116 can meet the film hole 90 at a radiused edge 118. The radiused edge 118 can define at least one flute 120 forming the inlet 92. The at least one flute 120 may be arcuate, multiply curved so as to provide local recessed bowls, straight, or piecewise linear. As illustrated, the one flute 120 is arcuate and tapers into the passage 96. Thus, the ramp 108, curved portion 116 and flute 120 all taper in the direction the direction of the cooling fluid flow C to define a sequentially narrow path for cooling fluid into the film holes 90.
The inlets 92 are aligned with the ledges 122, 124, with the ramps 126, 128 positioned between adjacent inlets 92. The ledges 122, 124 and ramps 126, 128 may have corners and edges as shown in
The ramps 126 on the upstream side wall 102 decline in the direction of the cooling fluid flow C, such that the upstream or top edge 130 of the ramp 126 is coincident with the cooling surface 86 and the downstream or bottom edge 132 of the ramp 126 is coincident with the bottom wall 100 of the channel. The declined ramps 126 further taper inwardly in the downstream direction, such that the ramp 126 is wider at the top edge 130 and narrower at the bottom edge 132.
The ramps 128 on the downstream side wall 104 incline in the direction of the cooling fluid flow C, such that the upstream or bottom edge 134 of the ramp 128 is coincident with the bottom wall 100 and the downstream or top edge 136 of the ramp 128 is coincident with the cooling surface 86. The inclined ramps 128 further taper outwardly in the downstream direction, such that the ramp 128 is narrower at the bottom edge 134 and wider at the top edge 136.
In any of the above embodiments, it is understood that while the drawings may show the contoured portion having sharp corners, edges, and transitions for purposes of illustration, is may be more practical for the corners, edges, and/or transitions to be smoothly radiused or filleted to avoid the formation of stagnation points within the engine component 80.
While many of the embodiments show the contoured portion 98 extending across the cooling surface 86 in a channel-like manner, this need not be the case. The cooling surface 86 can comprise discrete and multiple contoured portion 98 that do or do not extend entirely across the cooling surface 86 of the component 80. The embodiments also show the inlets 92 primarily located in the geometric center of the contoured portion 98, which is also not necessary for the invention. The inlets 92 can be located anywhere within the contoured portion 98. With respect to the cooling fluid flow C, the inlets 92 can be located at either an upstream or downstream edge of the contoured portion 98. The inlets 92 can even be partially located outside the contoured portion 98. The inlets 92 may be located on any area of the contoured portion 98, be it a flat area or curved area.
The ellipsoidal concavity 138 extends only partially across the cooling surface 86 and includes an incurvate recessed surface 140 that meets the cooling surface 86 at a perimeter edge 142. The perimeter edge 142 can be smoothly radiused or filleted to avoid the formation of stagnation points within the engine component 80.
In the above embodiments, the cooling fluid flow C is shown as being in a direction generally across the cooling surface 86 of the engine component 80, with the film holes 90 being arranged in a row extending generally transverse to the direction of the cooling fluid flow C. However, other row orientations with respect to the main direction of the cooling fluid flow C are possible. For example, for some engine components, most notably blades, the film holes 90 may be arranged in a row having an orientation parallel to that of the cooling fluid flow C. Is it noted that the cooling fluid flow C is turbulent, and is composed of directional components or vectors, particularly on a local scale with respect to the film holes, but that the main or bulk flow direction can be transverse to, parallel to, or some combination thereof, the row of film holes.
In any of the above embodiments, a protective coating, such as a thermal barrier coating, or multi-layer protective coating system can be applied to the hot surface 84 of the engine component 80. It is also understood that the film holes 90 and inlets 92 may have various orientations, not just the axial orientations shown in the figures. Furthermore, the present invention may be combined with shaping or contouring of the outlet 94 and passage 96 of the film holes 90. The present invention may also apply to slot-type film cooling, in which case the outlets 94 are provided within a slot on the hot surface 84.
The various embodiments of systems, methods, and other devices related to the invention disclosed herein provide improved cooling for engine components, particularly in an engine component having film holes. One advantage that may be realized in the practice of some embodiments of the described systems is that the cooling surface of the engine component can be shaped to include a contoured portion encompassing the inlets of multiple film holes. Conventional film hole design utilizes a passage with a circular inlet region, a metering section, and a shaped outlet region to help diffuse the cooling fluid. However, shaping of the inlet region has been limited. By shaping the film hole to include a contoured inlet region, improved cooling performance and mitigation of particle buildup in the engine component is achievable, which can lead to longer service life of the engine component.
Another advantage that may be realized in the practice of some embodiments of the described systems and methods is that multiple film holes may be encompassed within a regional contoured portion. Conventionally, surface contouring of film hole inlets requires local shaping around or into each individual film hole. By encompassing multiple inlets within a common contour, local design needs may be met, including protection against particles impacting the inlet surfaces, preconditioning the cooling fluid flow with additional pressure loss to obtain a better film exit condition, re-directing the cooling fluid flow to provide a more beneficial entry vector into the film holes, or eliminating the typical entry flow separation and consequent high turbulence and/or shock inside the film holes.
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention 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.
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/US15/57718 | 10/28/2015 | WO | 00 |
| Number | Date | Country | |
|---|---|---|---|
| 62073429 | Oct 2014 | US | |
| 62073455 | Oct 2014 | US |
