Patent Publication Number: US-2023151737-A1

Title: Airfoil with axial cooling slot having diverging ramp

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     The present disclosure claims the benefit of U.S. Provisional Application No. 63/281,041 filed Nov. 18, 2021 and U.S. Provisional Application No. 63/315,674 filed Mar. 2, 2022. 
    
    
     BACKGROUND 
     A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines. 
     The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the low inner shaft. A direct drive gas turbine engine includes a fan section driven by the low spool such that the low pressure compressor, low pressure turbine and fan section rotate at a common speed in a common direction. 
     SUMMARY 
     An airfoil according to an example of the present disclosure includes a platform, and an airfoil section that extends from the platform and defines leading and trailing ends and suction and pressure sides. The airfoil section has a transition region through which the airfoil section blends into the platform. The trailing end of the airfoil section has an axial cooling slot that opens through the transition region and defines an circumferentially diverging ramp in the transition region. 
     In a further embodiment of any of the foregoing embodiments, the circumferentially diverging ramp is radially open. 
     In a further embodiment of any of the foregoing embodiments, the circumferentially diverging ramp is radially sloped. 
     In a further embodiment of any of the foregoing embodiments, the circumferentially diverging ramp has first and second circumferential edges and a radial floor there between that is bowed from the first circumferential edge to the second circumferential edge. 
     In a further embodiment of any of the foregoing embodiments, the platform defines a core gas path surface, and the radial floor is radially offset from the core gas path surface. 
     In a further embodiment of any of the foregoing embodiments, the platform defines a platform trailing end, and the circumferentially diverging ramp has first and second circumferential edges that diverge from each other toward the platform trailing end. 
     In a further embodiment of any of the foregoing embodiments, the airfoil section includes a barrier coating on the circumferentially diverging ramp, and wherein the axial cooling slot with the barrier coating is non-metering for cooling flow there from. 
     In a further embodiment of any of the foregoing embodiments, the circumferentially diverging ramp is radially open and radially sloped. 
     In a further embodiment of any of the foregoing embodiments, the platform defines a platform trailing end, and the circumferentially diverging ramp has first and second circumferential edges that diverge from each other toward the platform trailing end. 
     In a further embodiment of any of the foregoing embodiments, the circumferentially diverging ramp has a radial floor between the first and second circumferential edges that is bowed from the first circumferential edge to the second circumferential edge. 
     In a further embodiment of any of the foregoing embodiments, the transition region does not have any inflection points, is tangent to both the surface of the airfoil section, and is also tangent to a gas path surface on the platform. 
     A gas turbine engine according to an example of the present disclosure includes a compressor section, a combustor in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor. The turbine section has airfoils as in any of the foregoing embodiments. 
     In a further embodiment of any of the foregoing embodiments, the circumferentially diverging ramp is radially open. 
     In a further embodiment of any of the foregoing embodiments, the circumferentially diverging ramp is radially sloped. 
     In a further embodiment of any of the foregoing embodiments, the circumferentially diverging ramp has first and second circumferential edges and a radial floor there between that is bowed from the first circumferential edge to the second circumferential edge. 
     In a further embodiment of any of the foregoing embodiments, the platform defines a core gas path surface, and the radial floor is radially inboard of the core gas path surface. 
     In a further embodiment of any of the foregoing embodiments, the platform defines a platform trailing end, and the circumferentially diverging ramp has first and second circumferential edges that diverge from each other toward the platform trailing end. 
     In a further embodiment of any of the foregoing embodiments, the airfoil section includes a barrier coating on the circumferentially diverging ramp, and wherein the axial cooling slot with the barrier coating is non-metering for cooling flow there from. 
     An investment casting core according to an example of the present disclosure includes a refractory core body for investment casting an air airfoil that has a platform and airfoil section that extends from the platform and defines leading and trailing ends and suction and pressure sides. The airfoil section has a transition region through which the airfoil section blends into the platform. The trailing end of the airfoil section has an axial cooling slot that opens through the transition region and defines a circumferentially diverging ramp in the transition region. The core body includes a core body trailing edge that has an enlarged lobe that is configured to form the circumferentially diverging ramp. 
     In a further embodiment of any of the foregoing embodiments, the enlarged lobe is enlarged in a circumferential direction. 
     The present disclosure may include any one or more of the individual features disclosed above and/or below alone or in any combination thereof. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The various features and advantages of the present disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows. 
         FIG.  1    illustrates a gas turbine engine. 
         FIG.  2    illustrates an airfoil of the engine. 
         FIG.  3    illustrates a view of a portion of the airfoil of  FIG.  2   . 
         FIG.  4    illustrates an investment casting core for manufacturing the airfoil of  FIG.  2   . 
         FIG.  5    illustrates a relative position of the core embedded in wax. 
         FIG.  6    illustrates another relative position of the core embedded in wax. 
         FIG.  7    illustrates the airfoil geometry that results from casting around the core as in  FIG.  5   . 
         FIG.  8    illustrates the airfoil geometry that results from casting around the core as in  FIG.  6   . 
         FIG.  9    illustrates another example airfoil that has an axially diverging ramp. 
         FIG.  10    illustrates an investment casting core for manufacturing the airfoil of  FIG.  9   . 
         FIG.  11    illustrates the core of  FIG.  10    embedded in wax. 
         FIG.  12    illustrates a tool for manufacturing the investment casting core of  FIG.  10   . 
     
    
    
     DETAILED DESCRIPTION 
       FIG.  1    schematically illustrates a gas turbine engine  20 . The 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 . The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a housing  15  such as a fan case or nacelle, and also drives air along a core gas 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 relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided, and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects, a first (or low) pressure compressor  44  and a first (or 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 a 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 second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in the exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  may be arranged generally between the high pressure turbine  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 which is collinear with their longitudinal axes. 
     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 through the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  that are in the core gas 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  48  may be varied. For example, gear system  48  may be located aft of the low pressure compressor, or aft of the combustor section  26  or even aft of turbine section  28 , and fan  42  may be positioned forward or aft of the location of gear system  48 . 
     The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), and can be less than or equal to about 18.0, or more narrowly can be less than or equal to 16.0. The geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3. The gear reduction ratio may be less than or equal to 4.0. The low pressure turbine  46  has a pressure ratio that is greater than about five. The low pressure turbine pressure ratio can be less than or equal to 13.0, or more narrowly less than or equal to 12.0. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to an inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1 and less than about 5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. The engine parameters described above and those in this paragraph are measured at this condition unless otherwise specified. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45, or more narrowly greater than or equal to 1.25. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/(518.7°R)] 0.5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150.0 ft/second (350.5 meters/second), and can be greater than or equal to 1000.0 ft/second (304.8 meters/second). 
       FIG.  2    illustrates an example airfoil  60  from the turbine section  28  of the engine  20 . As will be evident from the following description, the airfoil  60  provides examples of several potential manufacturing scenarios that are related to, and form a basis of, this disclosure. The airfoil  60  is depicted as a blade, but it is to be understood that the examples herein are also applicable to vanes. The airfoil includes a platform  62  and an airfoil section  64  that extends from the platform. The platform  62  defines a core gas path surface  62   a  that serves to bound a portion of the core gas path C of the engine  20 . The airfoil section  64  includes leading and trailing ends  64   a/   64   b  and suction and pressure sides  64   c/   64   d.  An internal cooling passage  66  in the airfoil section  64  receives cooling (bleed) air from the compressor section  24  of the engine  20 . The trailing end  64   b  of the airfoil section  64  includes axial cooling slots  68  that are connected with the cooling passage  66 . The axial cooling slots  68  extend through the trailing end  64   b  and serve as exits for the cooling air to be discharged from the cooling passage  66  into the core gas path C. For example, the axial cooling slot  68  may i) extend through the circumferential mid-line of the trailing end  64   b  and open at the trailing edge, ii) extend circumferentially and open at the pressure side  64   d  near the trailing edge, or iii) open in a conjoined matter at the pressure side  64   d  and trailing edge. 
       FIG.  3    shows a portion of the airfoil  60  from the region indicated in  FIG.  2   . The airfoil  64  has a transition region  70  through which the airfoil  64  blends into the platform  62 . The transition region  70  may be or may include what is known as the fillet region. For example, the transition region  70  is a curved-surface region that does not have any inflection points (points where there is an inversion in direction of curvature), is tangent on one end to the aerodynamic surface of the airfoil  64 , and is also tangent on its other end to the gas path surface  62   a.  Rather than a continuance of the aerodynamic surface of the airfoil  64  and the gas path surface  62   a  to meet at a corner, the transition region  70  provides a soft blend between the aerodynamic surface of the airfoil  64  and the gas path surface  62   a.  As shown, at least one of the axial cooling slots  68  opens through the transition region  70 . 
     The airfoil  60  is formed of a metal alloy, such as a Ni- or Co-based superalloy, and is fabricated by investment casting. For instance, in investment casting, a refractory metal or ceramic investment core is arranged in a mold and coated with a wax material, which ultimately defines the geometry of the airfoil  60 . The wax material is then coated with another material, such as a metallic or ceramic slurry that is hardened into a shell. The wax is melted out of the shell and molten metal is then poured into the cavity that remains between the shell and the core. The metal solidifies around the core to form the airfoil. The core is then removed in a known manner, leaving internal passages within the airfoil. 
       FIG.  4    illustrates a representative portion of an example investment core  72  for forming the internal cooling passage  66  and axial cooling slots  68  in the airfoil  60 . The investment core  72  represents a negative of the cooling passage  66  and axial cooling slots  68  in which solid structures of the investment core  72  produce void structures in the airfoil  60  and void structures of the investment core  72  produce solid structures in the airfoil  60 . The investment core  72  includes a refractory metal or ceramic core body  74 . In this example, the core body  74  includes pedestal openings  74   a  for forming pedestals in the internal cooling passage  66  and teardrop openings  74   b  for forming teardrop features that radially bound and define the axial cooling slots  68 . As will be appreciated, the investment core  72  may alternatively or additionally include other openings and geometries depending on the particular design of the airfoil  60  for its desired end use performance Aft of the teardrop openings  74   b  the trailing end portion of the investment core  72  includes an oversize region  74   c,  the function of which will be described in more detail below. 
     In the investment casting process, the relative position of the investment core  72  in the wax  76  (and thus also the relative position of the cooling passage  66  and axial cooling slots  68  in the airfoil  60 ) is subject to tolerances of the tooling and techniques that are used. For instance, the actual position of the investment core  72  and the actual position of the wax  76  can differ from their nominal design positions. As a result, in some instances as represented in  FIG.  5   , the investment core  72  may be radially-outwardly displaced such that it is a radial distance D 1  above the surface  76   a  of the wax  76  (which corresponds to the gas path surface  62   a  in the final airfoil  60 ). In other instances, as represented in  FIG.  6   , the investment core  72  may be radially-inwardly displaced such that it is a radial distance D 2  below the surface  76   a  of the wax  76 . 
     After the shell is molded around the wax  76  and the molten metal is poured to form the airfoil  60 , the metal encases the trailing end of the refractory casting core  72 , i.e., the oversize region  74   c.  Therefore, the axial cooling slots  68  of the airfoil  60  are initially closed-off after casting. A portion of the metal that encases the oversize region  74   c  is then machined back in order to open the axial cooling slots  68 . 
     In the instance of  FIG.  5    where the investment core  72  is radially-outwardly displaced, the result is that the radially inner-most axial cooling slot  68  is displaced radially outwardly from the gas path surface  62   a,  which is shown in  FIG.  7   . In the instance of  FIG.  6    where the investment core  72  is radially-inwardly displaced, the result is that the radially inner-most axial cooling slot  68  is displaced radially inwardly from the gas path surface  62   a,  which is shown in  FIG.  8   . In either instance, there may be a detriment to performance and/or manufacturing. For example, in the instance of  FIG.  7   , there is a region R 1  that is a relatively large mass of solid metal. Such a mass, which is not the design intent, may influence thermal transfer and thus effect airfoil durability. In the instance of  FIG.  8   , the truncated axial cooling slot  68  leaves a dead-ended pocket  68   a  in the transition region  70 . The dead-ended pocket  68   a  i) may collect grind chips during manufacturing that are difficult to remove and ii) has relatively sharp edges that may collect debris during operation. Moreover, even if a portion of the pocket  68   a  is initially open, the opening may be of such small size that it is filled by coating material in subsequent processes to deposit a barrier coating on the airfoil  60 . Accordingly, in either of the instances above, the relative displacement of the investment core  72  may lead to undesired features in the final airfoil  60 . 
     The machining of the airfoil  60  may be limited by the grind wheel sizing and accessibility to the area. In general, relatively small grind wheels have lower life, produce rougher finished surfaces, require higher operator skill, increase “touch” time, and challenge consistent repeatable finished geometry. To mitigate at least the scenario of  FIG.  5   , a buffer distance is maintained between the nominal design position of the refractory casting core  72  and the gas path surface  62   a  so that even if the core  72  is displaced inwardly it is still radially above the core gas path surface  62   a.  The buffer distance provides space to envelope the core  72  in wax with sufficient transition to allow for proper molten metal to fill in the trailing edge core slip envelope, represented by region  74   c,  without the need to hand finish the transition region  70  and thus enable a constant repeatable product. 
     The mitigation may be more effective on relatively large airfoils where the tolerances are a relatively small fraction of the trailing edge radial length. However, as the parts become smaller the same tolerances and the minimum wax thickness and transition requirements for proper metal flow become a significant percentage of the trailing edge length. This may be exacerbated on dual gas-path vanes where the reduction in cooled trailing edge exit is doubled. Thus, on smaller parts it is common to run the wax slip envelopment into the gas paths and accept that the area will have major geometric compromises. Such compromises may take two forms. The first involves hand finishing the entire interface with a grinding belt, resulting in a higher geometric variability that may affect geometric accuracy of the region. The second involves CNC grinding or electric discharge machining to consistently cut the trailing edge within tolerances just above the gas path and apply the standard break edges. Although each of these compromises can meet design requirements for the resultant finished airfoil  60 , there are inevitably debits to performance and manufacturing from the geometry variations that must be balanced in a given design. 
       FIG.  9    illustrates an airfoil  160  that, in comparison to the airfoil  60 , has a geometry that accounts for core position tolerance during casting and thus enables greater repeatability in manufacturing. As shown, the airfoil  160  includes axial cooling slots  168 . The radially inner-most one of the axial cooling slots  168  opens through the transition region  70  and defines a circumferentially diverging ramp  180  in the transition region  70 . The ramp  180  is defined by first and second circumferential edges  180   a/   180   b  that diverge from each other toward a platform trailing end  62   b  before converging to a rounded ramp trailing edge  180   c.  The ramp  180  has a radial floor  180   d  that is bowed from the first circumferential edge  180   a  to the second circumferential edge  180   b.  The radial floor  180   d  is radially sloped toward the platform trailing end  62   b,  and at least a portion of the radial floor  180   d  is radially inboard of the core gas path surface  62   a.  The top, i.e. the radially outer plane from the first circumferential edge  180   a  to the second circumferential edge  180   b,  is radially open. The airfoil  160  may further include a barrier coating  181 , such as but not limited to a stabilized zirconia coating. However, the axial cooling slot  168  and ramp  180  are of sufficient size to prevent plugging by the coating  181  and thereby ensure that the slot  168  remains non-metering. That is, if a substantial amount of coating material plugs a slot, the flow area of the slot may be reduced to an extent at which it restricts outflow. 
     As will be discussed below, the ramp  180  mitigates the effects of variations in the relative position of the axial cooling slots  168  relative to the gas path surface  62   a.  The ramp  180  may also facilitate a smooth flow of cooling air from the axial cooling slot  168  into the core gas path C due to the bowing and floor slope, a reduction in metal mass in the transition region  70 , an increase in surface area (e.g., via the radial floor  180   d ), and an increase in the area of the axial cooling slot  168  to reduce the possibility that a later-applied coating plugs the slot  168  and to allow ejection of debris particles from the airfoil  160 . Additionally, during machining, the divergent geometry of the ramp  180  may also facilitate access to clean out any grinding chips that may settle into the slot  168 . 
       FIG.  10    illustrates the investment core  172  that is configured to form the axial cooling slot  168  and ramp  180 . The investment core  172  is similar to the above-described core  72  in that it also includes a core body  174  with pedestal openings  174   a  and teardrop openings  174   b.  However, the radially inner section of the oversize region  174   c  includes an enlarged lobe  174   d  that is configured to form the circumferentially diverging ramp  180 . For instance, in comparison to the radially outer section of the oversize region  174   c  (e.g., at location RO), the enlarged lobe  174   d  is enlarged in the circumferential direction and extends farther aft in the axial direction (both relative to engine central longitudinal axis A). Thus, in the example shown, the cross-sectional area of the oversize region  174   c  through the lobe  174   d  taken in a plane parallel to the engine central longitudinal axis A increases from the location RO along the radially inward direction to a location L 1 . The cross-sectional area of the lobe  174   d  is at its maximum at L 1  and then decreases a remainder of the radial inner distance of the oversize region  174   d  and through an elbow location L 2 , at which point the edge of the core  172  curves forward beyond the radial plane of the location RO. 
       FIG.  11    illustrates the investment core  172  enveloped in wax  76 , with the trailing end  64   b  and transition region  70  of the final airfoil  160  superimposed thereon. When the metal that encases the oversize region  174   c  is machined back in order to open the axial cooling slots  168 , the plane of the machining will cut through the region that corresponds to the lobe  174   d,  thereby also opening the circumferentially diverging ramp  180 . The geometry of the ramp  180  thus correspond to the profile of the lobe  174   d.    
     The size of the lobe  174   d  in terms of the locations RO and L 2  is at least equal to, but is preferably larger than, the maximum tolerance that the position of the core  172  can vary with respect to the gas path surface  162   a.  Thus, at any position within the maximum tolerance envelope, the machined plane will cut through the region that corresponds to the lobe  174   d.  This assures that the circumferentially diverging ramp  180 , with its desirable geometry, will be opened during machining, rather than the highly variable geometry that results from core position variation for airfoil  60 . 
     The lobe  174   d  begins to radially decrease and circumferentially widen within the prescribed diffusion area flow area changes of the other axial cooling slots  168 . This is a change in geometry which maintains a consistency of flow area that begins just after the aft most metering location of the axial cooling slot  168 . The metering area is defined by the minimum cross-sectional area between trailing edge features, and the diffusion area is defined as the region between the metering location and the trailing edge of the trailing edge internal features. Beyond the diffusion region of the trailing edge slot, the lobe  174   d  continues to widen and ramp toward and beyond the gas path to a degree that will account for any possible coating material that is deposited onto the ramp  180  and ensure that any radial movement of the investment core relative to the wax will produce an acceptable and repeatable trailing edge exit geometry. 
       FIG.  12    schematically illustrates an investment casting tool  100  that may be used in a molding process to form the refractory casting core  172  disclosed herein. For example, the casting tool  100  is formed of a metal alloy and may contain a hard-facing or other protective surface to reduce wear and erosion. The tool  100  includes a molding cavity  101 . In this example, the molding cavity  101  is shown as an inverse view, e.g., the refractory casting core as represented and described in  FIG.  10   . The molding cavity  101  thus has the attributes of the core as described herein. 
     Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments. 
     The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.