Patent Publication Number: US-9416666-B2

Title: Turbine blade platform cooling systems

Description:
TECHNICAL FIELD 
     The present application relates generally to gas turbine engines and more particularly relates to turbine blade platform cooling systems so as to cool the suction side of adjacent blade platforms. 
     BACKGROUND OF THE INVENTION 
     Known turbine assemblies generally include rows of circumferentially spaced turbine blades. Generally described, each turbine blade includes an airfoil extending outwardly from a platform and a shank with a dovetail extending inwardly therefrom. The dovetail is used to mount the turbine blade to a rotor disc for rotation therewith. Known turbine blades generally are hollow such that an internal cooling cavity may be defined through at least portions of the airfoil, the platform, the shank, and the dovetail. 
     Temperature mismatches may develop at the interface between the airfoil and the platform and/or between the shank and the platform because the airfoil portions of the blades are exposed to higher temperatures than the shank and the dovetail portions. Over time, such temperature differences and associated thermal strains may induce large compressive thermal stresses to the blade platform. Moreover, the increased operating temperatures of the turbine as a whole may cause oxidation, fatigue, cracking, and/or creep deflection and, hence, a shorten useful life for the turbine blade. The potential stresses to the overall turbine blade and the bucket platform in particular generally increase with higher turbine combustion temperatures. 
     There is thus a desire for a turbine blade with improved cooling, particularly about the suction side of the platform. Such an improved turbine blade design would allow for the use of higher combustion temperatures and, hence, higher overall system efficiency with increased component lifetime. 
     SUMMARY OF THE INVENTION 
     The present application thus provides a turbine blade cooling system. The turbine blade cooling system may include a first turbine blade with a first turbine blade platform having a cooling cavity in communication with a pressure side passage and a second turbine blade with a second turbine blade platform having a platform cooling cavity with a suction side passage. The pressure side passage of the first turbine blade platform is in communication with the suction side passage of the second turbine blade platform. 
     The present application further provides a method of cooling a turbine blade platform. The method may include the steps of flowing a cooling medium through a pressure side passage of a first turbine blade platform, flowing the cooling medium through a suction side passage of a second turbine blade platform, flowing the cooling medium through a platform cooling cavity in the second turbine blade platform, and cooling the second turbine blade platform. 
     The present application further provides a turbine blade platform. The turbine blade platform may include a pressure side passage, a cooling circuit in communication with the pressure side passage, a suction side passage, and a platform cooling cavity in communication with the suction side passage. 
     These and other features and improvements of the present application will become apparent to one of ordinary skill in the art upon review of the following detailed description when taken in conjunction with the several drawings and the appended claims. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic view of the components of a known gas turbine engine. 
         FIG. 2  is a perspective view of a known turbine blade. 
         FIG. 3  is a top plan view of a pair of turbine blades of the turbine blade platform cooling system as may be described herein. 
         FIG. 4  is a side cross-sectional view of the pair of turbine blades of the turbine blade platform cooling system of  FIG. 3 . 
         FIG. 5  is a partial side perspective view of the pair of turbine blades of the turbine blade platform cooling system of  FIG. 3  as separated. 
     
    
    
     DETAILED DESCRIPTION 
     Referring now to the drawings, in which like numerals refer to like elements throughout the several views,  FIG. 1  shows a schematic view of the components of a known gas turbine engine  10 . The gas turbine engine  10  may include a compressor  15 . The compressor  15  compresses an incoming flow of air  20 . The compressor  15  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  are in turn delivered to a turbine  40 . The flow of combustion gases  35  drives the turbine  40  so as to produce mechanical work. The mechanical work produced in the turbine  40  drives the compressor  15  and an external load  45  such as an electrical generator and the like. 
     The gas turbine engine  10  may use natural gas, various types of syngas, and other types of fuels. The gas turbine engine  20  may be one of any number of different gas turbines offered by General Electric Company of Schenectady, N.Y. or otherwise. The gas turbine engine  10  may have other configuration and may use other types of components. Other types of gas turbine engines also may be used herein. Multiple gas turbine engines  10 , other types of turbines, and other types of power generation equipment may be used herein together. 
       FIG. 2  shows a perspective view of a known turbine blade  50 . The turbine blade  50  may be used in the turbine  40  as described above and the like. Any number of the blades  50  may be arranged adjacent to each other in a circumferentially spaced array. Each turbine blade  50  generally includes an airfoil  55  extending from a platform  60 . The airfoil  55  may be convex in shape with a suction side  65  and a pressure side  70 . Each airfoil  55  also may have a leading edge  75  and a trailing edge  80 . Other airfoil configurations also may be used herein. 
     The turbine blade  50  also may include a shank  85  and a dovetail  90  extending inwardly from the platform  60 . A number of angel wings  86  may be attached to the shank  85 . The dovetail  90  may attach the turbine blade  50  to a disc (not shown) for rotation therewith. The shank  85  may be substantially hollow with a shank cavity  95  therein. The shank cavity  95  may be in communication with a cooling medium such compressor discharge air. Other types of cooling circuits and cooling mediums also may be used herein. The cooling medium may circulate through at least portions of the dovetail  90 , the shank  85 , the platform  60 , and into the airfoil  55 . Other configurations may be used herein. 
       FIGS. 3-5  show a turbine blade platform cooling system  100  as may be described herein. The turbine blade platform cooling system  100  may include any number of turbine blades  110  although only a first turbine blade  120  and a second turbine blade  130  are shown. As described above, any number of the turbine blades  110  may be circumferentially positioned adjacent to each other about a rotor disc (not shown). Each pair of the turbine blades  110  may define a gap  140  therebetween. The first turbine blade  120  and the second turbine blade  130  may be substantially identical. 
     Each turbine blade  110  may include a platform  150  with an airfoil  160  extending outwardly therefrom and a shank  170  extending inwardly therefrom. The platform  150  may have a forward side  152 , an aft side  154 , a suction side  156 , and a pressure side  158 . 
     The turbine blade  110  may include a cooling cavity  180  extending therethrough. The cooling cavity  180  may be in communication with a cooling medium  190  such as compressor discharge air and the like. The cooling cavity  180  may extend at least in part through the shank  170  and into the airfoil  160 . A portion of the cooling cavity  180  also may extend into the platform  150  such that at least a portion of the cooling medium  190  may pass therethrough, either instead of or after passing through the airfoil  160 . Specifically, the cooling cavity  180  may extend into the aft portion  154  of the platform  150  about the pressure side  158  thereof. The portion of the cooling cavity  180  may end about a pressure side passage  200  of the platform  150 . Other configurations may be used herein. 
     The platform  150  also may include a platform cooling cavity  210 . The platform cooling cavity  210  may extend from the suction side  156  of the platform  150  towards the aft side  154 . The platform cooling cavity  210  may begin about a suction side passage  220 . The suction side passage  220  may align with the pressure side passage  200  of the adjoining turbine blade  110  so as to pass the cooling medium  190  therethrough. The platform cooling cavity  210  also may include an aft side passage  230  so as to discharge the cooling medium  190  once it passes therethrough. The platform cooling cavity  210  also may include a pin bank or other types of turbulators  240  therein so as to provide turbulence for enhanced heat transfer. Other types of internal configurations may be used herein. 
     In use, the cooling medium  190  passes through the cooling channel  180  of the first turbine blade  120 . At least a portion of the cooling medium  190  passes through the platform  150  and exits via the pressure side passage  200 . The cooling medium  190  then passes through the gap  140  and into the platform cooling cavity  210  of the second turbine blade  130 . Specifically, the cooling medium  190  passes into the suction side passage  220  of the platform cooling cavity  210  positioned on the suction side  156  of the platform  150  along the aft end  154  thereof. The cooling medium  190  then may exit the platform  150  along the aft side passage  230 . 
     The turbine blade platform cooling system  100  thus provides cooling on the suction side  156  of the platform  150  of the second turbine blade  130  via the cooling medium  190  from the first turbine blade  120 . The pin bank or other types of turbulators  240  within the platform cooling cavity  210  also provide enhanced heat transfer therein. This cooling also provides some lateral flexibility between the cooler shank side and the hot gas side of the platform  150  so as to reduce thermal stresses therein. Surface film holes and the like also may be used herein in communication with the platform cooling cavity  210 . Various types of seals also may be used about the gap  140  to reduce leakage and ingestion therethrough. 
     The turbine blade platform cooling system  100  thus provides platform cooling to enable higher turbine operating temperatures so as to provide higher efficiencies and lower customer operating costs with less impact on component durability. Using the cooling medium  190  from the first blade  120  so as to cool the second blade  130  further increases such overall efficiency. Transfer of the cooling medium  190  also may be made from the suction side  156  to the pressure side  158  in a similar manner. Any type of platform to platform cooling schemes in any direction may be used herein. 
     It should be apparent that the foregoing relates only to certain embodiments of the present application and that numerous changes and modifications may be made herein by one of ordinary skill in the art without departing from the general spirit and scope of the invention as defined by the following claims and the equivalents thereof.