Patent Publication Number: US-8974182-B2

Title: Turbine bucket with a core cavity having a contoured turn

Description:
TECHNICAL FIELD 
     The present application and the resultant patent relate generally to gas turbine engines and more particularly relate to a gas turbine engine with a turbine bucket having an airfoil with a core cavity having a contoured turn about a platform so as to reduce stress therein due to thermal expansion. 
     BACKGROUND OF THE INVENTION 
     Known gas turbine engines generally include rows of circumferentially spaced nozzles and buckets. A turbine bucket generally includes an airfoil having a pressure side and a suction side and extending radially upward from a platform. A hollow shank portion may extend radially downward from the platform and may include a dovetail and the like so as to secure the turbine bucket to a turbine wheel. The platform generally defines an inner boundary for the hot combustion gases flowing through a gas path. As such, the platform may be an area of high stress concentration due to the hot combustion gases and the mechanical loading thereon. 
     More specifically, there is often a large amount of thermally induced strain at the intersection of an airfoil and a platform. This thermally induced strain may be due to the temperature differential between the airfoil and the platform. The thermally induced strain may combine with geometric discontinuities in the region so as to create areas of very high stress that may limit component lifetime. To date, these issues have been addressed by attempting to keep geometric discontinuities such as root turns, internal ribs, and the like, away from the intersection. Further, attempts have been made to control the temperature about the intersection. Temperature control, however, generally requires additional cooling flows at the expense of overall engine efficiency. These known cooling arrangements, however, thus may be difficult and expensive to manufacture and may require the use of an excessive amount of air or other types of cooling flows. 
     There is thus a desire for an improved turbine bucket for use with a gas turbine engine. Preferably such a turbine bucket may limit the stresses at the intersection of an airfoil and a platform without excessive manufacturing and operating costs and without excessive cooling medium losses for efficient operation and an extended component lifetime. 
     SUMMARY OF THE INVENTION 
     The present application and the resultant patent thus provide a turbine bucket. The turbine bucket may include a platform, an airfoil extending from the platform at an intersection thereof, and a core cavity extending within the platform and the airfoil. The core cavity may include a contoured turn about the intersection so as to reduce thermal stress therein. 
     The present application and the resultant patent further provide a turbine bucket. The turbine bucket may include a platform, an airfoil extending from the platform at an intersection thereof, and a trailing edge core cavity extending within the platform and the airfoil. The trailing edge core cavity may include a cooling conduit with a contoured turn about the intersection so as to reduce thermal stress therein. 
     The present application and the resultant patent further provide a turbine bucket. The turbine bucket may include a platform, an airfoil extending from the platform at an intersection thereof, a trailing edge core cavity extending within the platform and the airfoil, and a cooling medium flowing therethrough. The trailing edge core cavity may include a contoured turn about the intersection with an area of reduced thickness so as to reduce thermal stresses therein. 
     These and other features and improvement of the present application and the resultant patent 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 diagram of a gas turbine engine with a compressor, a combustor, and a turbine. 
         FIG. 2  is a perspective view of a known turbine bucket. 
         FIG. 3  is a side plan view of a core body of a turbine bucket as may be described herein. 
         FIG. 4  is an expanded view of a trailing edge core cavity as may be described herein. 
         FIG. 5  is a sectional view of a portion of the trailing edge core cavity of  FIG. 4 . 
         FIG. 6  is a further sectional view of a portion of the trailing edge core cavity of  FIG. 4 . 
     
    
    
     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 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  15  delivers the compressed flow of air  20  to a combustor  25 . The combustor  25  mixes the compressed flow of air  20  with a pressurized 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 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  via a shaft  45  and an external load  50  such as an electrical generator and 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 any one of a number of different gas turbine engines offered by General Electric Company of Schenectady, N.Y., including, but not limited to, those such as a 7 or a 9 series heavy duty gas turbine engine 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  shows an example of a turbine bucket  55  that may be used with the turbine  40 . Generally described, the turbine bucket  55  includes an airfoil  60 , a shank portion  65 , and a platform  70  disposed between the airfoil  60  and the shank portion  65 . The airfoil  60  generally extends radially upward from the platform  70  and includes a leading edge  72  and a trailing edge  74 . The airfoil  60  also may include a concave wall defining a pressure side  76  and a convex wall defining a suction side  78 . The platform  70  may be substantially horizontal and planar. Likewise, the platform  70  may include a top surface  80 , a pressure face  82 , a suction face  84 , a forward face  86 , and an aft face  88 . The top surface  80  of the platform  70  may be exposed to the flow of the hot combustion gases  35 . The shank portion  65  may extend radially downward from the platform  70  such that the platform  70  generally defines an interface between the airfoil  60  and the shank portion  65 . The shank portion  65  may include a shank cavity  90  therein. The shank portion  65  also may include one or more angle wings  92  and a root structure  94  such as a dovetail and the like. The root structure  94  may be configured to secure the turbine bucket  55  to the shaft  45 . Other components and other configurations may be used herein. 
     The turbine bucket  55  may include one or more cooling circuits  96  extending therethrough for flowing a cooling medium  98  such as air from the compressor  15  or from another source. The cooling circuits  96  and the cooling medium  98  may circulate at least through portions of the airfoil  60 , the shank portion  65 , and the platform  70  in any order, direction, or route. Many different types of cooling circuits and cooling mediums may be used herein. Other components and other configurations also may be used herein. 
       FIGS. 3-6  show an example of a turbine bucket  100  as may be described herein. The turbine bucket  100  may include an airfoil  110 , a platform  120 , and a shank portion  130 . Similar to that described above, the airfoil  110  extends radially upward from the platform  120  and includes a leading edge  140  and a trailing edge  150 . Within the turbine bucket  100  there may be a number of core cavities  160 . The core cavities  160  supply a cooling medium  170  to the components thereof so as to cool the overall turbine bucket  100 . The cooling medium  170  may be air, steam, and the like from any source. In this example, a leading edge core cavity  180 , a central core cavity  190 , and a trailing edge core cavity  200  are shown. A number of the core cavities  160  may be used herein. Other components and other configurations may be used. 
     Generally described, the trailing edge core cavity  200  may be in the form of a cooling conduit  210 . The cooling conduit  210  may define a cooling passage  220  extending therethrough for the cooling medium  170 . The cooling conduit  210  may extend from a cooling input  230  about the shank portion  130  towards the platform  120  and the airfoil  110 . At about an intersection  240  between the platform  120  and the airfoil  110 , the cooling conduit  210  may expand at a contoured turn  250 . The contoured turn  250  thus may have an area of an increased edge radius  260 . The cooling passage  220  therein likewise expands through the contoured turn  250  so as to reduce the thickness of the material thereabout. Specifically, the contoured turn  250  may have an area of a reduced wall thickness  255 . 
     The cooling conduit  210  continues through a series of pins  270  or other types of turbulators through the airfoil  110 . Likewise, a number of cooling tubes  280  leading to a number of cooling holes  290  may extend towards the trailing edge  150  so as to provide film cooling to the airfoil  110 .  FIG. 5  shows the contoured turn  250  of the cooling conduit  210  about the intersection  240 . Likewise,  FIG. 6  shows the expanded cooling section  220  about the intersection  240 . Other components and other configurations also may be used herein. 
     The use of the contoured turn  250  in the cooling conduit  210  about the intersection  240  between the airfoil  110  and the platform  120  reduces the stiffness at the intersection  240  via the reduced wall thickness  255 . The reduced stiffness thus reduces stress therein due to temperature differences between the airfoil  110  and the platform  120 . The reduced wall thickness  255  about the contoured turn  250  also allows for the larger edge radius  260 . The larger edge radius  260  also reduces the peak stresses therein. Reducing stress at the intersection  240  should provide increased overall lifetime with reduced maintenance and maintenance costs. Moreover, the reduced wall thickness  255  and increased edge radius  260  may make the overall trailing edge core cavity  200  stronger so as to prevent core breakage during manufacture and thus decreasing overall casting costs. Further, excessive amounts of the cooling medium  170  may not be required herein. The overall impact of thermal expansion to the turbine bucket  100  thus may be reduced. 
     It should be apparent that the foregoing relates only to certain embodiments of the present application and the resultant patent. 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.