Patent Publication Number: US-2016237850-A1

Title: Systems and methods for vane cooling

Description:
FIELD 
     This disclosure relates to a gas turbine engine, and more particularly, to cooling hole arrangements on an airfoil such as a vane. 
     BACKGROUND 
     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. 
     SUMMARY 
     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. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  illustrates an example turbine engine, in accordance with various embodiments. 
         FIG. 2  illustrates a vane configuration with cooling holes, in accordance with various embodiments; 
         FIG. 3  illustrates a vane cooling hole configuration of prior art; 
         FIGS. 4 a  and 4 b    illustrate an exemplary vane cooling hole configuration, in accordance with various embodiments; 
         FIG. 5  illustrates a vane exit temperature profile of prior art 
         FIGS. 6 a  and 6 b    illustrate an exemplary vane exit temperature profile, in accordance with various embodiments; and 
         FIG. 7  illustrates an airfoil section from leading edge to half axial chord length, in accordance with various embodiments. 
     
    
    
     DETAILED DESCRIPTION 
     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  FIG. 1 , a gas turbine engine  20  is provided. Gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle  15 , while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     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  FIG. 1 ). It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided, including for example, bearing system  38 , bearing system  38 - 1 , and bearing system  38 - 2 . 
     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. 
       FIG. 2  schematically illustrates an isometric view of a HPT  54  vane configuration  200 . Vane configuration  200  illustrates first stage stator vane  260  for a turbine section  28  (with momentary reference to  FIG. 1 ). However, it is to be understood that other types of airfoils arranged downstream from a combustor and/or other parts of the gas turbine engine  20  may benefit from the examples disclosed herein which are not limited to the design shown. 
     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. 
     In various embodiments, with reference to  FIG. 4B  and continued reference to  FIG. 2 , cooling air may exit cooling holes  216  and flow in a direction closely aligned with arrows  416 B forming cooling flow streamlines as exemplified by arrow  420 B. Arrows  416 B may also be defined as vectors. Similarly, cooling air may exit cooling holes  218  and flow in a direction closely aligned with arrows  418 B forming cooling flow streamlines as exemplified by arrow  422 B. In this manner, it can be said that plane  430 B is a plane across which the orientation of cooling holes  216  and  218  are mirrored such that cooling flow streamlines  420 B and  422 B converge in close proximity to plane  430 B. Cooling hole angles theta (θ) and phi (φ) are closely aligned with arrows  418 A and arrows  418 B respectively. 
     In various embodiments, with further reference to  FIG. 4A , the plane about which cooling holes are mirrored may not be a tangible plane, but may be a location on the airfoil about which the orientation of the cooling holes is mirrored. In further embodiments, plane  430 A and  430 B may not be a plane about which the cooling hole angles are mirrored, but rather a plane about which the cooling hole angles change from theta (θ) to phi (φ). For example, if theta (θ) and phi (φ) are not exactly the same, then by definition the angles are not mirrored. In various embodiments, a cooling hole which is oriented at an angle which is not mirrored about a plane may be referred to as a cooling hole at a third angle. In further embodiments, angles theta (θ) and phi (φ) may be different in order to tailor for different design purposes. 
     With reference to  FIG. 3  and  FIG. 5 , current cooling hole orientations are configured such that cooling flow streamlines  320  and  322  converge in a location in close proximity to plane  330  on vane  300 . The location of plane  330  may be defined as being situated at a percentage of the span  335  of airfoil  360 . Airfoil  360  spans from proximal edge  326  to distal edge  324  across span  335 . Proximal edge  326  may be defined as being located at zero percent (0%) span of airfoil  360 . Distal edge  324  may be defined as being located at one hundred percent (100%) span of airfoil  360 . Plane  330  is located at the mid-span of airfoil  300 . The mid-span may be defined as being located at fifty percent (50%) span. 
     With reference to  FIG. 5 , a graph of operating temperature (on the x axis) and percent span (on the y axis) is shown for vane  300 . Because cooling flow streamlines  320  and  322  converge at the mid-span of airfoil  300 , a cool spot  566  is noted in the vane exit temperature profile as shown in  FIG. 5 . Line  564  represents a vane exit temperature profile of a vane without converging cooling flow streamlines. Line  562  represents a vane exit temperature profile of a vane with cooling flow streamlines converging at the mid-span of the vane. 
     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. 
     In various embodiments, with reference to  FIG. 4A  and  FIG. 6A , the location of plane  430 A may be located at a point other than mid-span (i.e., 50% span). The exemplary illustration of  FIG. 4A  shows plane  430 A being located at approximately seventy-five percent (75%) span. By disposing plane  430 A at approximately seventy-five percent (75%) span, the region of convergence of cooling air shifts to the approximate location of plane  430 A. Shifting the region of convergence of cooling air may allow the vane exit temperature profile  662 A in  FIG. 6A  to be tailored such that cool spot  666 A is located at seventy-five percent (75%) span as viewed on the y-axis. 
     In various embodiments, with reference to  FIG. 4B  and  FIG. 6B , the location of plane  430 B may be located at a point other than mid-span (i.e., 50% span). The exemplary illustration of  FIG. 4B  shows plane  430 B being located at approximately twenty-five percent (25%) span. By disposing plane  430 B at twenty-five percent (25%) span, the region of convergence of cooling air shifts to the approximate location of plane  430 B. Shifting the region of convergence of cooling air may allow the vane exit temperature profile  662 B in  FIG. 6B  to be tailored such that cool spot  666 B is located at twenty-five percent (25%) span as viewed on the y-axis. 
     In various embodiments, with renewed reference to  FIGS. 4A and 4B , the size and shape of cooling holes  416 A,  416 B,  418 A, and  418 B may vary. Cooling holes  416 A,  416 B,  418 A, and  418 B can be any variation of circular, ovular, rectangular, triangular, or any other shape in cross section. 
     In various embodiments, with reference to  FIG. 2 , plane  230  in each vane in an array of stator vanes  260  may include the same cooling hole configuration as other vanes in the array. For example, each vane in an array may include cooling holes which are mirrored about a plane located in the same radial location as the neighboring vane. In various embodiments, the angle of cooling holes  216  and cooling holes  218  in each vane in an array of stator vanes may be the same as other vanes in the array. For example, each vane in an array may include cooling holes oriented at the same angle as the neighboring vane. 
     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  FIG. 2  may be referred to as neighboring vanes because they are located next to each other in the circumferential direction. 
     In various embodiments, cooling hole angles theta (θ) and phi (φ) may be tailored to affect cool spot  666 A and  666 B, with brief reference to  FIGS. 6A and 6B . If increased cooling is desired, angles theta (θ) and phi (φ) may be increased in order to concentrate cooling air flow on a particular location to increase cooling. Similarly, angles theta (θ) and phi (φ) may be decreased in order to disperse the cooling air flow for decreased localized convergence of cooling air. 
     In various embodiments, with reference to  FIG. 7 , airfoil  700  may be defined as having a leading edge  736  on the forward side, a trailing edge  742  on the aft side, and a pressure side  738 . Cooling holes may be located on the leading edge and pressure side of airfoil  700  and also within the first half of the axial chord length  710  of airfoil  700 . Cooling holes located in the first half of the axial chord length  710  may have the greatest effect on the airfoil exit temperature profile. 
     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.