Patent Publication Number: US-2016245097-A1

Title: Airfoil and method for manufacturing an airfoil

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     The present application claims filing benefit of U.S. patent application Ser. No. 13/600,717 having a filing date of Aug. 31, 2012, which is incorporated by reference herein in its entirety. 
    
    
     FIELD OF THE INVENTION 
     The present invention generally involves an airfoil and a method for manufacturing an airfoil. 
     BACKGROUND OF THE INVENTION 
     Turbines are widely used in industrial and commercial operations. A typical commercial steam or gas turbine used to generate electrical power includes alternating stages of stationary and rotating airfoils. For example, stationary vanes may be attached to a stationary component such as a casing that surrounds the turbine, and rotating blades may be attached to a rotor located along an axial centerline of the turbine. A compressed working fluid, such as but not limited to steam, combustion gases, or air, flows through the turbine, and the stationary vanes accelerate and direct the compressed working fluid onto the subsequent stage of rotating blades to impart motion to the rotating blades, thus turning the rotor and performing work or generating thrust. 
     The efficiency of the turbine generally increases with increased temperatures of the compressed working fluid. However, excessive temperatures within the turbine may reduce the longevity of the airfoils in the turbine and thus increase repairs, maintenance, and outages associated with the turbine. As a result, various designs and methods have been developed to provide cooling to the airfoils. For example, a cooling media may be supplied to a cavity inside the airfoil to convectively and/or conductively remove heat from the airfoil. In particular embodiments, the cooling media may flow out of the cavity through cooling passages in the airfoil to provide film cooling over the outer surface of the airfoil. 
     The cavity and cooling passages in the airfoil may be manufactured using an investment casting process commonly referred to as a lost wax process. The lost wax process uses a ceramic core to define the cavity inside the airfoil. A wax is applied over the ceramic core, and the wax surface is shaped into the desired curvature for the airfoil. The wax-covered ceramic core is then repeatedly dipped into a liquid ceramic solution to create a ceramic shell over the wax surface. The wax may then be heated to remove the wax from between the ceramic core and the ceramic shell, creating a void between the ceramic core and the ceramic shell that serves as a mold for the airfoil. Molten metal may then be poured into the mold to form the airfoil. After the metal cools and solidifies, the ceramic shell may be broken and removed, exposing the metal that has taken the shape of the void created by the removal of the wax. The ceramic core may then be dissolved to produce the airfoil with the cavity and cooling passages. 
     Various efforts have been attempted to reduce the amount of cooling media flowing through the airfoil. For example, reducing the size and/or width of the cooling passages may enhance heat transfer to the cooling media while also reducing the amount of cooling media flowing through the airfoil. However, the smaller cooling passages require correspondingly smaller projections from the ceramic core that are sensitive to damage during the casting process. In particular, the projections from the ceramic core near either end of the ceramic core are susceptible to breaking off during casting. In an effort to strengthen the ceramic core while still providing smaller cooling passages, the projections from the ceramic core may be larger at either end and narrower in the middle. However, the larger projections may result in uneven cooling media flow through the correspondingly larger cooling passages, depriving the smaller cooling passages in the middle of the airfoil of sufficient cooling media flow. Accordingly, an airfoil and method for manufacturing an airfoil that produces a desired cooling media flow profile through cooling passages in the airfoil would be useful. 
     BRIEF DESCRIPTION OF THE INVENTION 
     Aspects and advantages of the invention are set forth below in the following description, or may be obvious from the description, or may be learned through practice of the invention. 
     One embodiment of the present invention is an airfoil that includes a pressure side, a suction side opposed to the pressure side, a cavity inside the airfoil between the pressure and suction sides, and a trailing edge downstream from the cavity between the pressure and suction sides. A first set of cooling passages through the trailing edge provide fluid communication from the cavity through the trailing edge. A first divider across each cooling passage in the first set of cooling passages extends from the pressure side to the suction side at the trailing edge. 
     Another embodiment of the present invention is an airfoil that includes a pressure side, a suction side opposed to the pressure side, a cavity inside the airfoil between the pressure and suction sides, and a trailing edge downstream from the cavity between the pressure and suction sides. A first set of cooling passages through the trailing edge provide fluid communication from the cavity through the trailing edge. A first set of pins extend across each cooling passage in the first set of cooling passages upstream from the trailing edge. 
     The present invention may also include an airfoil having a pressure side, a suction side opposed to the pressure side, a cavity inside the airfoil between the pressure and suction sides, and a trailing edge downstream from the cavity between the pressure and suction sides. A first set of cooling passages through the trailing edge provide fluid communication from the cavity through the trailing edge. A second set of cooling passages through the trailing edge provide fluid communication from the cavity through the trailing edge, and the first set of cooling passages are wider than the second set of cooling passages. The airfoil further includes first means for reducing flow through the first set of cooling passages. 
     Those of ordinary skill in the art will better appreciate the features and aspects of such embodiments, and others, upon review of the specification. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A full and enabling disclosure of the present invention, including the best mode thereof to one skilled in the art, is set forth more particularly in the remainder of the specification, including reference to the accompanying figures, in which: 
         FIG. 1  is a perspective view of an airfoil according to a first embodiment of the present invention; 
         FIG. 2  is a plan view of a core for manufacturing the airfoil shown in  FIG. 1 ; 
         FIG. 3  is a perspective view of an airfoil according to a second embodiment of the present invention; 
         FIG. 4  is a plan view of a core for manufacturing the airfoil shown in  FIG. 3 ; 
         FIG. 5  is a perspective view of an airfoil according to a third embodiment of the present invention; 
         FIG. 6  is a plan view of a core for manufacturing the airfoil shown in  FIG. 5 ; 
         FIG. 7  is a perspective view of an airfoil according to a fourth embodiment of the present invention; 
         FIG. 8  is a plan view of a core for manufacturing the airfoil shown in  FIG. 7 ; and 
         FIG. 9  is an exemplary graph of stresses in the core shown in  FIG. 8 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Reference will now be made in detail to present embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the invention. As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. In addition, the terms “upstream” and “downstream” refer to the relative location of components in a fluid pathway. For example, component A is upstream from component B if a fluid flows from component A to component B. Conversely, component B is downstream from component A if component B receives a fluid flow from component A. 
     Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that modifications and variations can be made in the present invention without departing from the scope or spirit thereof. For instance, features illustrated or described as part of one embodiment may be used on another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents. 
     Various embodiments of the present invention include an airfoil and a method for manufacturing an airfoil. The airfoil generally includes a pressure side having a concave curvature, a suction side having a convex curvature and opposed to the pressure side, a cavity inside the airfoil between the pressure and suction sides, and a trailing edge downstream from the cavity between the pressure and suction sides. The airfoil further includes one or more sets of cooling passages through the trailing edge that provide fluid communication from the cavity through the trailing edge. One or more of the sets of cooling passages may include various means for reducing flow through the cooling passages. In particular embodiments, for example, the means may include one or more dividers across some of the cooling passages at the trailing edge. In other particular embodiments, the means may include a set of pins that extend across some of the cooling passages. Although exemplary embodiments of the present invention will be described generally in the context of an airfoil incorporated into a turbine, one of ordinary skill in the art will readily appreciate from the teachings herein that embodiments of the present invention are not limited to a turbine unless specifically recited in the claims. 
     Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures,  FIG. 1  provides a perspective view of an airfoil  10  according to a first embodiment of the present invention. As shown in  FIG. 1 , the airfoil  10  generally includes a pressure side  12  having a concave curvature and a suction side  14  having a convex curvature and opposed to the pressure side  12 . The pressure and suction sides  12 ,  14  are separated from one another to define a cavity  16  inside the airfoil  10  between the pressure and suction sides  12 ,  14 . The cavity  16  may provide a serpentine or tortuous path for a cooling media to flow inside the airfoil  10  to remove heat from the airfoil  10 . The airfoil  10  further includes a trailing edge  18  downstream from the cavity  16  between the pressure and suction sides  12 ,  14 , and a plurality of cooling passages  20  through the trailing edge  18  provide fluid communication from the cavity  16  through the trailing edge  18 . As used herein, the term “trailing edge” is not limited to the most downstream portion of the airfoil  10  and may instead also include portions of the airfoil  10  on the pressure and/or suction sides  12 ,  14  that are downstream from the cavity  16 . 
     The cooling passages  20  may be arranged in multiple sets, with each set of cooling passages  20  having a different size, shape, and/or width. For example, a first set of cooling passages  22  located at the top and bottom of the trailing edge  18  may have a larger size and/or width than a second set of cooling passages  24  located in the middle of the trailing edge  18 . In the particular embodiment shown in  FIG. 1 , for example, the first set of cooling passages  22  may include three cooling passages  20  at the top and three cooling passages  20  at the bottom of the trailing edge  18 . As shown in  FIG. 1 , each cooling passage  20  may be tapered axially, and each cooling passage  20  in the first set of cooling passages  22  may have a size and/or width that is approximately three times as large as each cooling passage  20  in the second set of cooling passages  24 . One of ordinary skill in the art will readily appreciate from the teachings herein that the number of cooling passages  20  in the first set of cooling passages  22  may vary between 1 and 10 or more, and the present invention is not limited to any particular number of cooling passages  20  in any set of cooling passages  22 ,  24  unless specifically recited in the claims. Similarly, the difference in size, and/or width between the sets of cooling passages  20  may vary between approximately 1.1 times and 10 times or more, depending on the size of the airfoil  10  and number of different sets of cooling passages  22 ,  24  in the airfoil  10 , and the present invention is not limited to any particular difference in size and/or width of cooling passages  20  unless specifically recited in the claims. 
     The difference in size, shape, and/or width between the first and second sets of cooling passages  22 ,  24  would ordinarily create an undesirable disparity in cooling media flow along the length of the trailing edge  18 . Specifically, the larger size and/or width of the first set of cooling passages  22  would result in more cooling media flowing through the first set of cooling passages  22 , possibly resulting in insufficient cooling media flow through the second set of cooling passages  24 . To reduce this disparity, the first set of cooling passages  22  may further include means for reducing flow through the first set of cooling passages  22 . In the particular embodiment shown in  FIG. 1 , for example, the structure associated with the means may include a first divider  30  across each cooling passage  20  in the first set of cooling passages  22 . Each first divider  30  is essentially a post, tab, stub, pin, or similar structure that may extend from the pressure side  12  to the suction side  14  at the trailing edge  18 . As a result, each first divider  30  may partially obstruct cooling media flow through each cooling passage  20  in the first set of cooling passages  22  to reduce any disparity in cooling media flow along the length of the trailing edge  18 . Additionally, the first set of cooling passages  22  may be tapered more than the other cooling passages  20  to reduce this disparity even further. 
       FIG. 2  provides a plan view of a core  40  that may be used to manufacture the airfoil  10  shown in  FIG. 1 . As shown in  FIG. 2 , the core  40  may include a serpentine portion  42  with a number of long, thin branches or projections  44  that extend from the serpentine portion  42 . The serpentine portion  42  generally corresponds to the size and location for the cavity  16  in the airfoil  10 , and the projections  44  generally correspond to the size and location of the cooling passages  20  through the trailing edge  18 . For example, as shown in  FIG. 2 , the projections  44  may be grouped into a first set of projections  46  at the top and bottom of the core  40  that are approximately three times the size and/or width of the remaining projections  44  in a second set of projections  48  in the middle of the core  40 . In addition, the first set of projections  46  include tabs or notches  50  that generally correspond to the location of the first dividers  30  in the first set of cooling passages  22  described with respect to  FIG. 1 . The increased size and/or width of the first set of projections  46  enhances the durability and resistance to damage of the projections  44  during the subsequent casting operations. 
     The core  40  may be manufactured from any material having sufficient strength to withstand the high temperatures associated with the casting material (e.g., a high alloy metal) while maintaining tight positioning required for the core  40  during casting. For example, the core  40  may be cast from ceramic material, ceramic composite material, or other suitable materials. Once cast or otherwise manufactured, a laser, electron discharge machine, drill, water jet, or other suitable device may be used to refine or form the serpentine portion  42 , projections  44 , and/or notches  50  shown in  FIG. 2 . The core  40  may then be utilized in a lost wax process as is known in the art. For example, the core  40  may be coated with a wax or other suitable material readily shaped to the desired thickness and curvature for the airfoil  10 . The wax covered core  40  may then be repeatedly dipped into a liquid ceramic solution to create a ceramic shell over the wax surface. The wax may then be heated to remove the wax from between the core  40  and the ceramic shell, creating a void between the core  40  and the ceramic shell that serves as a mold for the airfoil  10 . Molten metal may then be poured into the mold to form the airfoil  10 . After the metal cools and solidifies, the ceramic shell may be broken and removed, exposing the metal that has taken the shape of the void created by the removal of the wax. The core  40  may then be dissolved to produce the airfoil  10  with the cavity  16 , cooling passages  20 , and first dividers  30  shown in  FIG. 1 . 
       FIG. 3  provides a perspective view of the airfoil  10  according to a second embodiment of the present invention. As shown in  FIG. 3 , the airfoil  10  generally includes the pressure side  12 , suction side  14 , cavity  16 , trailing edge  18 , and cooling passages  20  as previously discussed with respect to  FIG. 1 . In this particular embodiment, the cooling passages  20  are arranged in first, second, and third sets of cooling passages  22 ,  24 ,  26 , with each set of cooling passages  20  having a different size and/or width. As shown in  FIG. 3 , for example, the first set of cooling passages  22  includes a single cooling passage  20  located at the top and bottom of the trailing edge  18 , the second set of cooling passages  24  includes a single cooling passage  20  located next to each cooling passage  20  in the first set of cooling passages  22 , and the third set of cooling passages  26  includes the remaining cooling passages  20  located in the middle of the trailing edge  18 . In the particular embodiment shown in  FIG. 3 , each cooling passage  20  may be tapered axially. Each cooling passage  20  in the first set of cooling passages  22  may have a size and/or width that is approximately five times as large as each cooling passage  20  in the third set of cooling passages  26 , and each cooling passage  20  in the second set of cooling passages  24  may have a size and/or width that is approximately three times as large as each cooling passage  20  in the third set of cooling passages  26 . One of ordinary skill in the art will readily appreciate from the teachings herein that the number of cooling passages  20  in the first and second sets of cooling passages  22 ,  24  may vary between 1 and 10 or more, and the present invention is not limited to any particular number of cooling passages  20  in any set of cooling passages  22 ,  24 ,  26  unless specifically recited in the claims. Similarly, the difference in size and/or width between the sets of cooling passages  22 ,  24 ,  26  may vary between approximately 1.1 times and 10 times or more, depending on the size of the airfoil  10  and number of different sets of cooling passages  22 ,  24 ,  26  in the airfoil  10 , and the present invention is not limited to any particular difference in size, shape, and/or width of cooling passages  20  unless specifically recited in the claims. 
     The difference in size, shape, and/or width between the first, second, and third sets of cooling passages  22 ,  24 ,  26  would ordinarily create an undesirable disparity in cooling media flow along the length of the trailing edge  18 . Specifically, the larger size and/or width of the first and second sets of cooling passages  22 ,  24  would result in more cooling media flowing through the first and second sets of cooling passages  22 ,  24 , possibly resulting in insufficient cooling media flow through the third set of cooling passages  26 . To reduce this disparity, the first and/or second sets of cooling passages  22 ,  24  may further include means for reducing flow through the respective cooling passages  20 . In the particular embodiment shown in  FIG. 3 , for example, the structure associated with the means in the first set of cooling passages  22  may include multiple first dividers  30  across each cooling passage  20  in the first set of cooling passages  22 . Each first divider  30  is essentially a post, tab, stub, pin, or similar structure that may extend from the pressure side  12  to the suction side  14  at the trailing edge  18 . As a result, the multiple first dividers  30  may partially obstruct cooling media flow through each cooling passage  20  in the first set of cooling passages  22 . The structure associated with the means in the second set of cooling passages  24  may similarly include one or more second dividers  32  across each cooling passage  20  in the second set of cooling passages  24 . The combination of the means for reducing flow through the first and second sets of cooling passages  22 ,  24  reduces any disparity in cooling media flow along the length of the trailing edge  18 . Additionally, the first and/or second set of cooling passages  22 ,  24  may be tapered more than the third set of cooling passages  26  to reduce this disparity even further. 
       FIG. 4  provides a plan view of the core  40  that may be used to manufacture the airfoil  10  shown in  FIG. 3 . As shown in  FIG. 4 , the core  40  may again include the serpentine portion  42 , projections  44 , and notches  50  as previously described with respect to  FIG. 2 . In the particular embodiment shown in  FIG. 4 , the projections  44  are arranged in first, second, and third sets of projections  46 ,  48 ,  49  that correspond to the location and sizes of the cooling passages  20  in the first, second, and third sets of cooling passages  22 ,  24 ,  26 , respectively. Specifically, the first set of projections  46  include the projections  44  at the top and bottom of the core  40  that are approximately five times the size and/or width of the projections  44  in the third set of projections  49 . Similarly, the second set of projections  48  include the projections  44  adjacent to the first set of projections  46  that are approximately three times the size and/or width of the projections  44  in the third set of projections  49 . Lastly, the third set of projections  49  are located in the middle of the core  40 . In addition, the first and second sets of projections  46 ,  48  include the tabs or notches  50  that generally correspond to the location of the first and second dividers  30 ,  32  in the first and second sets of cooling passages  22 ,  24  described with respect to  FIG. 3 . The increased size and/or width of the first and second sets of projections  46 ,  48  enhances the durability and resistance to damage of the projections  44  during the subsequent casting operations. 
       FIG. 5  provides a perspective view of the airfoil  10  according to a third embodiment of the present invention, and  FIG. 6  provides a plan view of the core  40  for manufacturing the airfoil  10  shown in  FIG. 5 . As shown in  FIGS. 5 and 6 , the airfoil  10  and core  40  generally include the same components as previously described with respect to the embodiments shown in  FIGS. 1-4 . In this particular embodiment, each cooling passage  20  may be tapered axially. Each cooling passage  20  in the first set of cooling passages  22  may have a size and/or width that is approximately four times as large as each cooling passage  20  in the third set of cooling passages  26 , and each cooling passage  20  in the second set of cooling passages  24  may have a size and/or width that is approximately three times as large as each cooling passage  20  in the third set of cooling passages  26 . The structure associated with the means for reducing flow through the cooling passages  20  in the first and second sets of cooling passages  22 ,  24  again includes first and second dividers  30 ,  32 , as previously described with respect to  FIG. 3 . However, in the particular embodiment shown in  FIG. 5 , each first divider  30  is wider than each second divider  32 . Specifically, each first divider  30  may be wider than each second divider  32  by approximately 1.1 to 5 times or more, depending on the particular embodiment. As a result, the wider first dividers  30  may combine with the wider cooling passages  20  in the first set of cooling passages  22  to reduce any disparity in cooling media flow along the length of the trailing edge  18 . Additionally, the first and/or second set of cooling passages  22 ,  24  may be tapered more than the third set of cooling passages  26  to reduce this disparity even further. 
     As shown most clearly in  FIG. 6 , the first set of projections  46  include the projections  44  at the top and bottom of the core  40  that are approximately four times the size and/or width of the projections  44  in the third set of projections  49 , and the second set of projections  48  include the projections  44  adjacent to the first set of projections  46  that are approximately three times the size and/or width of the projections  44  in the third set of projections  49 . The increased size and/or width of the first and second sets of projections  46 ,  48  enhances the durability and resistance to damage of the projections  44  during the subsequent casting operations. 
       FIG. 7  provides a perspective view of the airfoil  10  according to a fourth embodiment of the present invention, and  FIG. 8  provides a plan view of the core  40  for manufacturing the airfoil  10  shown in  FIG. 7 . Specifically, the airfoil  10  generally includes the pressure side  12 , suction side  14 , cavity  16 , trailing edge  18 , and cooling passages  20  as previously discussed with respect to  FIG. 1 . In this particular embodiment, the first set of cooling passages  22  includes two cooling passages  20  at the top and bottom of the trailing edge  18 , and the second set of cooling passages  24  includes the cooling passages  20  located in the middle of the trailing edge  18 . Each cooling passage  20  may be tapered axially, and each cooling passage  20  in the first set of cooling passages  22  may have a size and/or width that is approximately three times as large as each cooling passage  20  in the second set of cooling passages  24 . The structure associated with the means for reducing flow through the cooling passages  20  in the first set of cooling passages  22  may include a first set of pins  60  that extend across each cooling passage  20  in the first set of cooling passages  22  upstream from the trailing edge  18 . The pins  60  may disrupt the cooling media flow through the cooling passages  20  to reduce the amount of cooling media flowing through the first set of cooling passages  22  while also enhancing heat exchange between the airfoil  10  and the cooling media. As shown in the particular embodiment shown in  FIG. 7 , one or more of the pins  60  may be axially staggered inside the cooling passages  20  to further enhance heat exchange and control cooling media flow through the cooling passages  20 . As a result, the first set of pins  60  may combine with the wider cooling passages  20  in the first set of cooling passages  22  to reduce any disparity in cooling media flow along the length of the trailing edge  18 . Additionally, the first set of cooling passages  22  may be tapered more than the second set of cooling passages  24  to reduce this disparity even further. One of ordinary skill in the art will readily appreciate from the teachings herein that in still further embodiments, the means for reducing flow through the second set of cooling passages  24  shown in  FIGS. 3 and 5  may include a second set of pins in each cooling passage  20  in the second set of cooling passages  24  upstream from the trailing edge  18 , and further illustration of this alternate structure is not necessary. 
     As shown most clearly in  FIG. 8 , the core  40  may again include the serpentine portion  42  and projections  44  as previously described with respect to  FIG. 2 . In the particular embodiment shown in  FIG. 8 , the projections  44  are arranged in first and second sets of projections  46 ,  48  that correspond to the location and sizes of the cooling passages  20  in the first and second sets of cooling passages  22 ,  24 , respectively. Specifically, the first set of projections  46  include the two projections  44  at the top and bottom of the core  40  that are approximately three times the size and/or width of the projections  44  in the second set of projections  48 . In addition, the first set of projections  46  include multiple holes  62  that generally correspond to the location of the first set of pins  60  in the first set of cooling passages  22  described with respect to  FIG. 7 . The increased size and/or width of the first set of projections  46  enhances the durability and resistance to damage of the projections  44  during the subsequent casting operations. 
       FIG. 9  provides an exemplary graph of stresses in the core  40  shown in  FIG. 8 . Specifically, the horizontal axis represents the ratio of the width of the projections  44  with and without the pins  60 , and the vertical axis represents the ratio of the stress on the projections with and without pins  60 . As shown in  FIG. 9 , doubling the width of the projections  44  and adding pins  60  to the projections reduces the stress across by projections  44  by more than 50%. For the particular embodiment shown in  FIG. 8  in which the projections  44  in the first set of projections  46  are approximately three times larger and/or wider than the projections  44  in the second set of projections  48 , the stress across the projections  44  in the first set of projections  46  are calculated to be less than 20% of the stress across the projections  44  in the second set of projections  48 . 
     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 include 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 language of the claims.