Turbine bucket with a core cavity having a contoured turn

The present application thus provides 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.

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.

DETAILED DESCRIPTION

Referring now to the drawings, in which like numerals refer to like elements throughout the several views,FIG. 1shows a schematic view of gas turbine engine10as may be used herein. The gas turbine engine10may include a compressor15. The compressor15compresses an incoming flow of air20. The compressor15delivers the compressed flow of air20to a combustor25. The combustor25mixes the compressed flow of air20with a pressurized flow of fuel30and ignites the mixture to create a flow of combustion gases35. Although only a single combustor25is shown, the gas turbine engine10may include any number of combustors25. The flow of combustion gases35is in turn delivered to a turbine40. The flow of combustion gases35drives the turbine40so as to produce mechanical work. The mechanical work produced in the turbine40drives the compressor15via a shaft45and an external load50such as an electrical generator and the like.

The gas turbine engine10may use natural gas, various types of syngas, and/or other types of fuels. The gas turbine engine10may 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 engine10may 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. 2shows an example of a turbine bucket55that may be used with the turbine40. Generally described, the turbine bucket55includes an airfoil60, a shank portion65, and a platform70disposed between the airfoil60and the shank portion65. The airfoil60generally extends radially upward from the platform70and includes a leading edge72and a trailing edge74. The airfoil60also may include a concave wall defining a pressure side76and a convex wall defining a suction side78. The platform70may be substantially horizontal and planar. Likewise, the platform70may include a top surface80, a pressure face82, a suction face84, a forward face86, and an aft face88. The top surface80of the platform70may be exposed to the flow of the hot combustion gases35. The shank portion65may extend radially downward from the platform70such that the platform70generally defines an interface between the airfoil60and the shank portion65. The shank portion65may include a shank cavity90therein. The shank portion65also may include one or more angle wings92and a root structure94such as a dovetail and the like. The root structure94may be configured to secure the turbine bucket55to the shaft45. Other components and other configurations may be used herein.

The turbine bucket55may include one or more cooling circuits96extending therethrough for flowing a cooling medium98such as air from the compressor15or from another source. The cooling circuits96and the cooling medium98may circulate at least through portions of the airfoil60, the shank portion65, and the platform70in 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. 3-6show an example of a turbine bucket100as may be described herein. The turbine bucket100may include an airfoil110, a platform120, and a shank portion130. Similar to that described above, the airfoil110extends radially upward from the platform120and includes a leading edge140and a trailing edge150. Within the turbine bucket100there may be a number of core cavities160. The core cavities160supply a cooling medium170to the components thereof so as to cool the overall turbine bucket100. The cooling medium170may be air, steam, and the like from any source. In this example, a leading edge core cavity180, a central core cavity190, and a trailing edge core cavity200are shown. A number of the core cavities160may be used herein. Other components and other configurations may be used.

Generally described, the trailing edge core cavity200may be in the form of a cooling conduit210. The cooling conduit210may define a cooling passage220extending therethrough for the cooling medium170. The cooling conduit210may extend from a cooling input230about the shank portion130towards the platform120and the airfoil110. At about an intersection240between the platform120and the airfoil110, the cooling conduit210may expand at a contoured turn250. The contoured turn250thus may have an area of an increased edge radius260. The cooling passage220therein likewise expands through the contoured turn250so as to reduce the thickness of the material thereabout. Specifically, the contoured turn250may have an area of a reduced wall thickness255.

The cooling conduit210continues through a series of pins270or other types of turbulators through the airfoil110. Likewise, a number of cooling tubes280leading to a number of cooling holes290may extend towards the trailing edge150so as to provide film cooling to the airfoil110.FIG. 5shows the contoured turn250of the cooling conduit210about the intersection240. Likewise,FIG. 6shows the expanded cooling section220about the intersection240. Other components and other configurations also may be used herein.

The use of the contoured turn250in the cooling conduit210about the intersection240between the airfoil110and the platform120reduces the stiffness at the intersection240via the reduced wall thickness255. The reduced stiffness thus reduces stress therein due to temperature differences between the airfoil110and the platform120. The reduced wall thickness255about the contoured turn250also allows for the larger edge radius260. The larger edge radius260also reduces the peak stresses therein. Reducing stress at the intersection240should provide increased overall lifetime with reduced maintenance and maintenance costs. Moreover, the reduced wall thickness255and increased edge radius260may make the overall trailing edge core cavity200stronger so as to prevent core breakage during manufacture and thus decreasing overall casting costs. Further, excessive amounts of the cooling medium170may not be required herein. The overall impact of thermal expansion to the turbine bucket100thus may be reduced.