Patent Application
20160237850
This disclosure relates to a gas turbine engine, and more particularly, to cooling hole arrangements on an airfoil such as a vane.
Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.
The efficiency of the engine can be increased by passing a higher temperature gas flow through the turbine. However, the turbine inlet temperature is limited to the vane and blade (airfoils) material properties and the cooling capabilities of these airfoils. The first stage airfoils are exposed to the highest temperature gas flow since these airfoils are located immediately downstream from the combustor.
An airfoil is disclosed comprising a distal edge, a proximal edge, a pressure side surface extending between the distal edge and the proximal edge and between a leading edge and a trailing edge, the leading edge forward of the trailing edge, wherein a first cooling hole and a second cooling hole are disposed in the pressure side surface, wherein the first cooling hole is oriented at a first angle relative to the distal edge and the second cooling hole is oriented at a second angle relative to the proximal edge, wherein the first angle and the second angle have bilateral symmetry about a plane, wherein the plane is located at a point at least one of more than or less than the midpoint between the distal edge and proximal edge.
An airfoil is disclosed comprising a distal edge, a proximal edge, a pressure side surface extending between the distal edge and the proximal edge and between a leading edge and a trailing edge, the leading edge forward of the trailing edge, wherein a first cooling hole and a second cooling hole are disposed in the pressure side surface, wherein the first cooling hole is oriented at a first angle relative to the distal edge and the second cooling hole is oriented at a second angle relative to the proximal edge, and a third hole is disposed on the pressure side surface at a third angle which is different from the first angle and second angle.
The detailed description of exemplary embodiments herein makes reference to the accompanying drawings, which show exemplary embodiments by way of illustration. While these exemplary embodiments are described in sufficient detail to enable those skilled in the art to practice the inventions, it should be understood that other embodiments may be realized and that logical changes and adaptations in design and construction may be made in accordance with this invention and the teachings herein. Thus, the detailed description herein is presented for purposes of illustration only and not of limitation. The scope of the invention is defined by the appended claims. For example, the steps recited in any of the method or process descriptions may be executed in any order and are not necessarily limited to the order presented. Furthermore, any reference to singular includes plural embodiments, and any reference to more than one component or step may include a singular embodiment or step. Also, any reference to attached, fixed, connected or the like may include permanent, removable, temporary, partial, full and/or any other possible attachment option. Additionally, any reference to without contact (or similar phrases) may also include reduced contact or minimal contact. Surface shading lines may be used throughout the figures to denote different parts but not necessarily to denote the same or different materials.
As used herein, “aft” refers to the direction associated with the tail (e.g., the back end) of an aircraft, or generally, to the direction of exhaust of the gas turbine. As used herein, “forward” refers to the direction associated with the nose (e.g., the front end) of an aircraft, or generally, to the direction of flight or motion.
As used herein, “distal” refers to the direction radially outward, or generally, away from the axis of rotation of a turbine engine. As used herein, “proximal” refers to a direction radially inward, or generally, towards the axis of rotation of a turbine engine. As used herein, “airfoil” and “vane” are used interchangeably.
In various embodiments and with reference to
The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A-A′ relative to an engine static structure 36 via one or more bearing systems 38 (shown as bearing system 38-1 and bearing system 38-2 in
The low speed spool 30 may generally comprise an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine (“HPT”) 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the HPT 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the HPT 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A-A′, which is collinear with their longitudinal axes. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine.
The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the HPT 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 50 may be varied. For example, gear system 50 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
Gas turbine engine 20 may be, for example, a high-bypass geared aircraft engine. In various embodiments, the bypass ratio of gas turbine engine 20 may be greater than about six (6). In various embodiments, the bypass ratio of gas turbine engine 20 may be greater than ten (10). In various embodiments, geared architecture 48 may be an epicyclic gear train, such as a star gear system (sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear) or other gear system. Geared architecture 48 may have a gear reduction ratio of greater than about 2.3 and low pressure turbine 46 may have a pressure ratio that is greater than about 5. In various embodiments, the bypass ratio of gas turbine engine 20 is greater than about ten (10:1). In various embodiments, the diameter of fan 42 may be significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 may have a pressure ratio that is greater than about (5:1). Low pressure turbine 46 pressure ratio may be measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of low pressure turbine 46 prior to an exhaust nozzle. It should be understood, however, that the above parameters are exemplary of various embodiments of a suitable geared architecture engine and that the present disclosure contemplates other gas turbine engines including direct drive turbofans.
The turbine section 28 includes at least one array of stator vane 260 arranged circumferentially about an engine axis A-A′ to define an outer radial flow path boundary for a core flow path C. Cooling air 212 and 214 may be directed from, for example, the compressor section 24 through bypass ducts and into vane configuration 200. A pressure gradient may force cooling air 212 and cooling air 214 through cooling hole set 216 and cooling hole set 218. Cooling air 212 may enter airfoil 200 via distal edge 224. Cooling air 214 may enter airfoil 200 via proximal edge 226. The cooling air 212 and 214 may exit vane 200 through cooling holes 216 and 218 and enter core flow path C. This cooling air may provide film cooling and convective cooling to reduce the vane operating temperature. Pressure side surface 238 extends from distal edge 224 to proximal edge 222. Leading edge 236 is shown on the pressure side surface 238 forward of trailing edge 242.
In various embodiments, the orientation of cooling holes 216 and cooling holes 218 may be angled about a plane that intersects the z axis at only a single point. Plane 230 passes through the center of each respective vane. In other words, each vane has its own respective plane 230 defined as a plane that intersects the z axis at only a single point. Each plane 230 may also be described as a plane which is normal to a line extending in the span-wise direction of the vane. In various embodiments, cooling holes 216 may be angled so that a trailing point of cooling holes 216 points towards the negative z-direction. Stated another way, cooling holes 216 may be angled in the positive theta (θ) direction. A cooling hole which is part of cooling holes 216 may be referred to as a first cooling hole. In various embodiments, cooling holes 218 may be angled so that a trailing point of cooling holes 218 points towards plane 230 in the positive z-direction. In other words, cooling holes 218 may be angled in the positive phi (φ) direction. A cooling hole which is part of cooling holes 218 may be referred to as a second cooling hole. In various embodiments, optimal angles theta (θ) and phi (φ) may be between fifteen (15) degrees and thirty-five (35) degrees, though as presently illustrated, angles theta (θ) and phi (φ) are approximately twenty-five (25) degrees. In various embodiments, angle theta (θ) may be referred to as a first angle. In various embodiments, angle phi (φ) may be referred to as a second angle.
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In various embodiments, cooling holes may introduced into an airfoil in any suitable manner. For example, cooling holes may be formed by electric discharge machining (EDM) methods. The use of EDM allows the orientation of cooling holes 316 and cooling holes 318 to be closely tailored. EDM is a manufacturing process whereby a desired shape is obtained using electrical discharges. Tailoring first stage turbine vane exit temperature profiles may be desirable for different design considerations. For example, combustor exit temperature profiles may benefit from a higher rate of cooling in certain areas on turbine vanes and blades. For example, components downstream of turbine vanes, such as other airfoils, may benefit from localized cooling which may be achieved by tailoring a turbine vane exit temperature profile.
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In various embodiments, exhaust flow temperature profiles may benefit from more complex cooling configurations. For example, certain components may benefit from higher rates of cooling than other areas. In various embodiments, the angle of cooling holes 216 and cooling holes 218 in each vane in an array of stator vanes 260 may be different than other vanes in the array. For example, each vane in an array may include cooling holes oriented at a different angle than the neighboring vane. In various embodiments, each vane in an array may include cooling holes which are mirrored about a plane located in a different radial location (e.g., along the z axis) than the neighboring vane. The two airfoils in
In various embodiments, cooling hole angles theta (θ) and phi (φ) may be tailored to affect cool spot 666A and 666B, with brief reference to
In various embodiments, with reference to
The method of creating the cooling holes may include any method of drilling, boring, or cutting as well as any other method known to persons of ordinary skill in the art. According to various embodiments, cooling holes are formed via EDM. According to various embodiments, cooling holes are formed via additive manufacturing processes. According to various embodiments, cooling holes are formed via subtractive manufacturing processes.
Benefits, other advantages, and solutions to problems have been described herein with regard to specific embodiments. Furthermore, the connecting lines shown in the various figures contained herein are intended to represent exemplary functional relationships and/or physical couplings between the various elements. It should be noted that many alternative or additional functional relationships or physical connections may be present in a practical system. However, the benefits, advantages, solutions to problems, and any elements that may cause any benefit, advantage, or solution to occur or become more pronounced are not to be construed as critical, required, or essential features or elements of the inventions. The scope of the inventions is accordingly to be limited by nothing other than the appended claims, in which reference to an element in the singular is not intended to mean “one and only one” unless explicitly so stated, but rather “one or more.” Moreover, where a phrase similar to “at least one of A, B, or C” is used in the claims, it is intended that the phrase be interpreted to mean that A alone may be present in an embodiment, B alone may be present in an embodiment, C alone may be present in an embodiment, or that any combination of the elements A, B and C may be present in a single embodiment; for example, A and B, A and C, B and C, or A and B and C. Systems, methods and apparatus are provided herein. In the detailed description herein, references to “one embodiment”, “an embodiment”, “various embodiments”, etc., indicate that the embodiment described may include a particular feature, structure, or characteristic, but every embodiment may not necessarily include the particular feature, structure, or characteristic. Moreover, such phrases are not necessarily referring to the same embodiment. Further, when a particular feature, structure, or characteristic is described in connection with an embodiment, it is submitted that it is within the knowledge of one skilled in the art to affect such feature, structure, or characteristic in connection with other embodiments whether or not explicitly described. After reading the description, it will be apparent to one skilled in the relevant art(s) how to implement the disclosure in alternative embodiments.
Furthermore, no element, component, or method step in the present disclosure is intended to be dedicated to the public regardless of whether the element, component, or method step is explicitly recited in the claims. No claim element herein is to be construed under the provisions of 35 U.S.C. 112(f), unless the element is expressly recited using the phrase “means for.” As used herein, the terms “comprises”, “comprising”, or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus.
