Patent Publication Number: US-2015075180-A1

Title: Systems and methods for providing one or more cooling holes in a slash face of a turbine bucket

Description:
FIELD 
     Embodiments of the disclosure relate generally to a gas turbine engine and more particularly relate to systems and methods for providing one or more cooling holes in a slash face of a turbine bucket. 
     BACKGROUND 
     A gas turbine engine typically includes a compressor, a combustor, and a turbine. The efficiency of the turbine depends in part on the amount of cooling air flow from the compressor that is used to cool components in the hot gas path in the turbine section. The cooling air flow may be introduced into the wheel space of the turbine to limit (or purge) high-temperature gases from entering into the wheel space. Excess purge flow to the wheel space may decrease turbine efficiency since the cooling air flow may not be available for work production. 
     BRIEF DESCRIPTION 
     Some or all of the above needs and/or problems may be addressed by certain embodiments of the disclosure. According to one embodiment, there is disclosed a turbine bucket. The turbine bucket may include a platform and a shank portion extending radially inward from the platform. The shank portion may include a slash face, a radial seal pin groove formed in the slash face, and at least one cooling hole disposed in the slash face about the radial seal pin groove. 
     According to another embodiment, there is disclosed a gas turbine engine system. The system may include a compressor, a combustor in communication with the compressor, and a turbine in communication with the combustor. The turbine bucket may include a platform and a shank portion extending radially inward from the platform. The shank portion may include a slash face, a radial seal pin groove formed in the slash face, and at least one cooling hole disposed in the slash face about the radial seal pin groove. 
     Further, according to another embodiment, there is disclosed a shank portion of a turbine bucket. The shank portion may include a slash face, a radial seal pin groove formed in the slash face, and at least one cooling hole disposed in the slash face about the radial seal pin groove. 
     Other embodiments, aspects, and features of the invention will become apparent to those skilled in the art from the following detailed description, the accompanying drawings, and the appended claims. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Reference will now be made to the accompanying drawings, which are not necessarily drawn to scale. 
         FIG. 1  schematically depicts an example view of a gas turbine engine. 
         FIG. 2  schematically depicts an example cross-sectional view of a turbine bucket. 
         FIG. 3  schematically depicts an example perspective cross-sectional view of a turbine bucket, according to an embodiment of the disclosure. 
         FIG. 4  schematically depicts an example perspective cross-sectional view of a turbine bucket, according to an embodiment of the disclosure. 
     
    
    
     DETAILED DESCRIPTION 
     Illustrative embodiments will now be described more fully hereinafter with reference to the accompanying drawings, in which some, but not all embodiments are shown. The disclosure may be embodied in many different forms and should not be construed as limited to the embodiments set forth herein. Like numbers refer to like elements throughout. 
       FIG. 1  depicts a schematic view of gas turbine engine  10  as may be used herein. The gas turbine engine  10  may include a compressor  15 . The compressor  15  compresses an incoming flow of air  20 . The compressor delivers the compressed flow of air  20  to a combustor  25 . The combustor  25  mixes the compressed flow of air  20  with a compressed flow of fuel  30  and ignites the mixture to create a flow of combustion gases  35 . Although only a single combustor  25  is shown, the gas turbine engine  10  may include any number of combustors  25 . The flow of combustion gases  35  is in turn delivered to a downstream turbine  40 . The flow of combustion gases  35  drives the turbine  40  to produce mechanical work. The mechanical work produced in the turbine  40  drives the compressor  15  via a shaft  45  and an external load  50 , such as an electrical generator or the like. 
     The gas turbine engine  10  may use natural gas, various types of syngas, and/or other types of fuels. The gas turbine engine  10  may be anyone of a number of different gas turbine engines such as those offered by General Electric Company of Schenectady, New York and the like. The gas turbine engine  10  may have different configurations and may use other types of components. Other types of gas turbine engines also may be used herein. Multiple gas turbine engines, other types of turbines, and other types of power generation equipment also may be used herein together. 
       FIG. 2  schematically depicts one example embodiment of a portion of the turbine  40 . The turbine  40  may include a rotor  52  positioned about a longitudinal axis. A number of buckets  54  may be mounted to the rotor  52 . For example, the buckets  54  may be circumferentially position adjacent to one another and extend radially outward from the rotor  52 . The buckets  54  may form one or more stages in the turbine  40 . For example, the buckets  54  may form a first stage, a last stage, or any stage therebetween. The buckets  54  may include a platform  56 , a shank portion  58 , an airfoil  60 , and a dovetail  62 . The dovetail  62  may be configured to mate with a corresponding dovetail  64  of the rotor  52 . 
     The shank portion  56  may include a slash face  66 . The slash face  66  is the circumferential edge of the shank portion  58 . In some instances, the leading edge of the shank portion  58  may include a forward trench cavity  68 . The forward trench cavity  68  may be formed between an angle wing seal  70  and a leading edge  72  of the platform  56 . The forward trench cavity  68  may provide an area where purge air from the wheelspace interfaces with the hot combustion gases. Other components and other configurations may be used herein. 
       FIGS. 3 and 4  depict an example embodiment of a turbine bucket  100  as may be used herein. The turbine bucket  100  may include may include a platform  102 , a shank portion  104 , an airfoil  106 , and a dovetail  108 . The shank portion  104  may extend radially inward from the platform  102 , and the airfoil  106  may extend radially outward from the platform  102 . The shank portion  104  may include a slash face  110 . The slash face  110  is the circumferential edge of the shank portion  104 . Depending on the orientation of the airfoil  106 , the slash face  110  may be a pressure side slash face (as depicted in  FIG. 3 ) or a suction side slash face (as depicted in  FIG. 4 ). That is, the slash face  110  positioned about the pressure side of the airfoil  106  is the pressure side slash face, and the slash face  110  positioned about the suction side of the airfoil  106  is the suction side slash face. 
     In some instances, the leading edge of the shank portion  104  may include a forward trench cavity  112 . The forward trench cavity  112  may be formed between an angle wing seal  114  and a leading edge  116  of the platform  102 . The forward trench cavity  112  may provide an area where purge air from the wheelspace interfaces with the hot combustion gases. 
     In certain embodiments, the turbine bucket  102  may include a radial seal pin groove  118  formed in the slash face  110 . The radial seal pin groove  118  may extend at least partially from the platform  102  to the dovetail  108 . In some instances, a radial seal pin  120  (depicted in dashed lined for clarity) may be positioned within the radial seal pin groove  118 . That is, each radial seal pin groove  118  may be sized and shaped to receive a radial seal pin  120  therein to facilitate sealing between adjacent shanks portions  104  when a number of turbine buckets  100  are coupled to the rotor. U.S. Patent Pub. No. 2011/0081245 and U.S. Pat. No. 7,600,972 both describe example embodiments of a radial seal pin groove and a radial seal pin and are both hereby incorporated by reference. In some instances, only the pressure side slash face and/or the suction side slash face may include the radial seal pin groove  118  and/or the radial seal pin  120 . In this manner, a slash face that does not include the radial seal pin groove  118  and/or the radial seal pin  120  may still form a seal with an adjacent turbine bucket  100  that does include the radial seal pin groove  118  and/or the radial seal pin  120 . 
     The turbine bucket  100  may include at least one cooling hole  122  disposed in the slash face  110  about the radial seal pin groove  118 . The cooling hole  122  may be disposed within a pressure side slash face and/or a suction side slash face. The cooling hole  122  may be configured to provide a flow of cooling fluid (e.g., air) to the area about the radial seal pin groove  118  and/or the radial seal pin  120 . For example, the cooling hole  122  may be in communication with a flow of diverted air from the compressor by way of a cooling circuit  124 . Other sources of air may be used. In some instances, the cooling circuit  124  may include a number of channels  126  or the like disposed within the turbine bucket  100 . In this manner, the cooling hole  122  may be in fluid communication with any one of the channels  126 . The orientation, configuration, and number of cooling circuits  124  and/or channels  126  may vary. 
     In certain embodiments, the cooling hole  122  may be disposed in the slash face  110  about the forward trench cavity  112 . That is, the cooling hole  122  may be disposed in the slash face  110  between the angle wing seal  114  and the leading edge  116  of the platform  102 . Alternatively, or in addition, the cooling hole  122  may be positioned about a radial outer portion of the radial seal pin groove  118 . In another instance, the cooling hole  122  may be positioned about an upstream portion of the radial seal pin groove  118  and/or a downstream portion of the radial seal pin groove  118 . The cooling hole  122  may be positioned at any location about the radial seal pin groove  118 . Furthermore, in some instances, the cooling hole  122  may include a number of cooling holes  122 . That is, a number of cooling holes  122  may be disposed in the slash face  110  at various locations about the radial seal pin groove  118 . 
     The location of the cooling holes  122  facilitates cooling of the area about the radial seal pin groove  118  and/or the radial seal pin  120 . In turn, the forward trench cavity  112  may require less purge air, resulting in greater efficiency of the gas turbine engine. 
     Although embodiments have been described in language specific to structural features and/or methodological acts, it is to be understood that the disclosure is not necessarily limited to the specific features or acts described. Rather, the specific features and acts are disclosed as illustrative forms of implementing the embodiments.