Patent Publication Number: US-2012045337-A1

Title: Turbine bucket assembly and methods for assembling same

Description:
BACKGROUND OF THE INVENTION 
     The subject matter described herein relates generally to gas turbine engines and, more particularly, to a bucket assembly for use with a turbine engine. 
     At least some known rotor assemblies used with turbine engines include at least one row of circumferentially-spaced rotor blades. Each rotor blade includes an airfoil that includes a pressure side and a suction side that are connected together along leading and trailing edges. Each airfoil extends radially outward from a rotor blade platform. Each rotor blade also includes a dovetail that extends radially inward from a shank defined between the platform and the dovetail. The dovetail is used to mount the rotor blade to a rotor disk or spool. Known blades are hollow and include an internal cooling cavity that is defined at least partially by the airfoil, platform, shank, and dovetail and that is used to channel a flow of cooling fluid. Leakage of cooling fluid may occur between adjacent rotor blades. Depending on the amount of leakage, turbine performance and output may be adversely impacted. 
     Furthermore, the airfoil portions of at least some known rotor blades are generally exposed to higher temperatures than the dovetail portions. Higher temperatures may cause temperature mismatches to develop at the interface between the airfoil and the platform, and/or between the shank and the platform. These temperature mismatches may cause compressive thermal stresses to be induced to the rotor blade platform. Over time, continued operation with high compressive thermal stresses may cause platform oxidation, platform cracking, and/or platform creep deflection, any or all of which may shorten the useful life of the rotor assembly. 
     BRIEF SUMMARY OF THE INVENTION 
     In one aspect, a method for assembling a rotor assembly for use with a turbine engine is provided. The method includes providing at least two rotor blades that each include a shank extending between a dovetail and a platform. The shank includes at least one cover plate that extends inwardly from the platform towards the dovetail. An airfoil extends outwardly from the platform. A first rotor blade is coupled to a rotor disk. A second rotor blade is coupled to the rotor disk, such that a cavity is defined between the first and second rotor blades, and such that a seal path is defined between a first rotor blade cover plate and a second rotor blade cover plate. 
     In a further aspect, a rotor blade for a turbine engine is provided. The rotor blade includes a platform that includes a radially outer surface and a radially inner surface. An airfoil extends radially outwardly from the platform. A dovetail is adapted to be coupled to a rotor wheel. A shank extends between the platform and the dovetail. The shank includes at least one cover plate that extends inwardly from the platform towards the dovetail. At least one sealing assembly is coupled to the cover plate. The sealing assembly extends from the dovetail to the platform. The sealing assembly forms a seal path between the rotor blade and a circumferentially adjacent rotor blade. 
     In another aspect, a gas turbine engine is provided. The gas turbine engine includes a compressor and a combustor coupled downstream from the compressor to receive at least some of the air discharged by the compressor. A rotor shaft is coupled to the compressor. A plurality of circumferentially-spaced rotor blades are coupled to the rotor shaft. Each of the plurality of rotor blades includes a platform. An airfoil extends radially outwardly from the platform. A dovetail is coupled to the rotor shaft. A shank extends between the platform and the dovetail. The shank includes at least one cover plate that extends inwardly from the platform towards the dovetail. At least one sealing assembly is coupled to the cover plate such that a seal path is defined between adjacent rotor blades. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is schematic illustration of an exemplary known turbine engine system. 
         FIG. 2  is an enlarged perspective view of an exemplary rotor assembly that may be used with the turbine engine system shown in  FIG. 1 . 
         FIG. 3  is an enlarged sectional view of a portion of the rotor assembly shown in  FIG. 2   
         FIG. 4  is a cross-sectional view of the rotor assembly shown in  FIG. 2 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The exemplary methods and systems described herein overcome disadvantages of known rotor blade assemblies by providing a rotor blade that facilitates reducing leakage of cooling fluid from the rotor blade. More specifically, the embodiments described herein include a labyrinth seal path that is positioned between adjoining rotor blades to facilitate increasing a back pressure between adjacent rotor blades and to facilitate reducing leakage of cooling fluid through the rotor blades. 
     As used herein, the term “rotor blade” is used interchangeably with the term “bucket” and thus can include any combination of a bucket including a platform and dovetail and/or a bucket integrally formed with the rotor disk, either of which may include at least one airfoil segment. 
       FIG. 1  is a schematic view of an exemplary gas turbine engine  10 . In the exemplary embodiment, gas turbine engine  10  includes an intake section  12 , a compressor section  14  coupled downstream from intake section  12 , a combustor section  16  coupled downstream from compressor section  14 , a turbine section  18  coupled downstream from combustor section  16 , and an exhaust section  20 . Turbine section  18  is includes a rotor assembly  22  that is coupled to compressor section  14  via a drive shaft  32 . Combustor section  16  includes a plurality of combustors  24 . Combustor section  16  is coupled to compressor section  14  such that each combustor  24  is in flow communication with compressor section  14  and such that fuel nozzle assembly  26  is coupled to each combustor  24 . Turbine section  18  is rotatably coupled to compressor section  14  and to a load  28  such as, but not limited to, an electrical generator and a mechanical drive application. In the exemplary embodiment, compressor section  14  and turbine section  18  each include at least one turbine blade or bucket  30  coupled to rotor assembly  22  that include airfoil portions (not shown in  FIG. 1 ). 
     During operation, intake section  12  channels air towards compressor section  14 . Compressor section  14  compresses the inlet air to a higher pressure and temperature and discharges the compressed air towards combustor section  16 . The compressed air is mixed with fuel and ignited to generate combustion gases that flow to turbine section  18 . Turbine section  18  drives compressor section  14  and/or load  28 . Specifically, at least a portion of compressed air supplied to fuel nozzle assembly  26 . Fuel is channeled to fuel nozzle assembly  26  wherein it is mixed with the air and ignited in combustor section  16 . Combustion gases are generated and channeled to turbine section  18  wherein gas stream thermal energy is converted to mechanical rotational energy. Exhaust gases exit turbine section  18  and flow through exhaust section  20  to ambient atmosphere. 
       FIG. 2  is an enlarged perspective view of an exemplary rotor assembly  22  that may be used with gas turbine engine  10  (shown in  FIG. 1 ).  FIG. 3  is an enlarged sectional view of a portion of rotor assembly  22 , and  FIG. 4  is a cross-sectional view of rotor assembly  22  taken along sectional line  4 - 4  in  FIG. 3 . In the exemplary embodiment, rotor assembly  22  includes at least one rotor blade  100  coupled to a rotor disk  102 . Moreover, in the exemplary embodiment, rotor assembly  22  includes a first rotor blade  104 , a second rotor blade  106 , and at least a third rotor blade  107 . In the exemplary embodiment, each rotor blade  100  is coupled to a rotor disk  102  that is rotatably coupled to a rotor shaft, such as drive shaft  32  (shown in  FIG. 1 ). In an alternative embodiment, rotor blades  100  are mounted within a rotor spool (not shown). More specifically, when rotor blades  100  are coupled to rotor disk  102 , a gap  108  is defined between adjacent circumferentially-spaced rotor blades  100 . In the exemplary embodiment, each rotor blade  100  extends radially outward from rotor disk  102  and includes an airfoil  110 , a platform  112 , a shank  114 , and a dovetail  116 . Each airfoil  110  includes a first sidewall  118  and a second sidewall  120  that is coupled to first sidewall  118  to form airfoil  110 . 
     In the exemplary embodiment, first sidewall  118  is convex and defines a suction side  119  of airfoil  110 , and second sidewall  120  is concave and defines a pressure side  121  of airfoil  110 . First sidewall  118  is coupled to second sidewall  120  along a leading edge  122  and along an axially-spaced trailing edge  124  of airfoil  110 . More specifically, airfoil trailing edge  124  is spaced chord-wise and downstream from airfoil leading edge  122 . First sidewall  118  and second sidewall  120  each extend longitudinally or radially outwardly in span from a blade root  126  positioned adjacent to platform  112 , to an airfoil tip  128 . In the exemplary embodiment, an internal cooling chamber  130  is defined within airfoil  110  between first sidewall  118  and second sidewall  120 , and extends through platform  112 , through shank  114 , and into dovetail  116 . 
     Platform  112  extends between airfoil  110  and shank  114  such that each airfoil  110  extends radially outwardly from platform  112 . Shank  114  extends radially inwardly from platform  112  to dovetail  116 . Dovetail  116  extends radially inwardly from shank  114  to enable rotor blades  100  to be coupled to rotor disk  102 . Platform  112  includes an upstream side or skirt  132 , and a downstream side or skirt  134  that are connected together with a pressure-side edge  136  and an opposite suction-side edge  138 . When rotor blades  100  are coupled to rotor disk  102 , a gap  140  is defined between circumferentially adjacent rotor blade platforms  112 , and more specifically between pressure-side edge  136  and an adjacent suction-side edge  138 . 
     In the exemplary embodiment, shank  114  includes a first sidewall  142 , a second sidewall  144 , an upstream sidewall or forward cover plate  146 , and an opposite downstream sidewall or aft cover plate  148 . Moreover, in the exemplary embodiment, first sidewall  142  is substantially concave and is coupled between forward cover plate  146  and aft cover plate  148  such that forward cover plate  146  is opposite aft cover plate  148 . Second sidewall  144  is substantially convex and is coupled between forward cover plate  146  and aft cover plate  148 . In one embodiment, first sidewall  142  is coupled to second sidewall  144  such that a cavity  150  is defined at least partially between first sidewall  142  and second sidewall  144 . In an alternative embodiment, first sidewall  142  is coupled to second sidewall  144  such that a unitary member extending between forward cover plate  146  and aft cover plate  148  is formed. In another alternative embodiment, shank  114  is formed as a unitary member. In the exemplary embodiment, first sidewall  142  and second sidewall  144  are each recessed with respect to forward cover plate  146  and aft cover plate  148 , respectively, such that when rotor blades  100  are coupled to rotor disk  102 , a shank cavity  152  is defined between first sidewall  142  and an adjacent second sidewall  144 . 
     In the exemplary embodiment, a forward angel wing  154  extends outwardly from forward cover plate  146 . An aft angel wing  156  extends outwardly from aft cover plate  148 . Forward angel wing  154  and aft angel wing  156  each facilitate sealing forward and aft angel wing buffer cavities (not shown) defined within rotor assembly  22 . In addition, a forward lower angel wing  158  extends outwardly from forward cover plate  146 , and is configured to facilitate sealing between rotor blade  100  and rotor disk  102 . More specifically, forward lower angel wing  158  extends outwardly from forward cover plate  146  between dovetail  116  and forward angel wing  154 . 
     In the exemplary embodiment, aft cover plate  148  includes a leading edge portion  164  and a circumferentially-spaced trailing edge portion  166 . A first sealing assembly  168  is coupled to leading edge portion  164 , and a second sealing assembly  170  is coupled to trailing edge portion  166 . In the exemplary embodiment, first sealing assembly  168  cooperates with an adjacent second sealing assembly  170  when rotor blades  100  are coupled to rotor disk  102 . First sealing assembly  168  and second sealing assembly  170  each extend between dovetail  116  and platform  112 , and each facilitates sealing shank cavity  152 . In the exemplary embodiment, first sealing assembly  168  and second sealing assembly  170  cooperate to form a seal path  172  between a first aft cover plate  148  and an adjacent second aft cover plate  148 . Seal path  172  facilitates reducing a volume of air channeled between circumferentially adjacent rotor blade shanks  114 . More specifically, seal path  172  facilitates reducing the volume of air that must be channeled from forward cover plate  146  to aft cover plate  148  through shank cavity  152  to facilitate preventing a flow of hot gases from entering shank cavity  152 . 
     In the exemplary embodiment, aft cover plate  148  extends a radial height r 1  from dovetail  116  to a platform inner surface  174 . First sealing assembly  168  and second sealing assembly  170  each extend a radial height r 2  from dovetail  116  to platform inner surface  174 . Radial height r 2  is approximately the same height as radial height r 1  of aft cover plate  148 . In one embodiment, first sealing assembly  168  and/or second sealing assembly  170  extends the full radial height r 1  of aft cover plate  148 . 
     In one embodiment, first sealing assembly  168  includes a sealing extension  176  that extends outwardly from leading edge portion  164  towards an adjacent rotor blade trailing edge portion  166 . Second sealing assembly  170  includes a recessed sealing groove  178  that is defined within trailing edge portion  166 . Recessed sealing groove  178  is sized to receive an adjacent sealing extension  176  such that recessed sealing groove  178  and sealing extension  176  cooperate to form seal path  172 . In an alternative embodiment, first sealing assembly  168  includes recessed sealing groove  178  and second sealing assembly  170  includes sealing extension  176 . 
     In the exemplary embodiment, first rotor blade  104  includes first sealing assembly  168 , including sealing extension  176 , and second sealing assembly  170 , including recessed sealing groove  178 . In an alternative embodiment, first rotor blade  104  includes first sealing assembly  168 , including recessed sealing groove  178 , and second sealing assembly  170 , including a sealing extension  176 . In one embodiment, second rotor blade  106  includes first sealing assembly  168  and second sealing assembly  170  each including sealing extension  176 . In an alternative embodiment, second rotor blade  106  includes first sealing assembly  168  and second sealing assembly  170  each including recessed sealing groove  178 . 
     In the exemplary embodiment, recessed sealing groove  178  includes a radially outer surface  184  that extends between dovetail  116  and platform inner surface  174 . An abradable layer  186  is coupled to recessed sealing groove outer surface  184 . Alternatively, in one embodiment, abradable layer  186  includes an aluminum composite material. In the exemplary embodiment, sealing extension  176  includes a plurality of labyrinth teeth  188  that extend outwardly from an inner surface  190  of sealing extension  176 . Labyrinth teeth  188  are each positioned adjacent to an opposing recessed sealing groove outer surface  184  such that a labyrinth seal  191  is defined between sealing extension  176  and recessed sealing groove  178 . 
     In the exemplary embodiment, shank  114  includes a leading edge radial seal pin slot  192  that extends generally radially through shank  114  at least partially between platform  112  and dovetail  116 . More specifically, leading edge radial seal pin slot  192  is defined within shank forward cover plate  146  and is adjacent to shank convex sidewall  144 . Leading edge radial seal pin slot  192  is sized to receive a radial seal pin  194  to facilitate sealing between adjacent forward cover plates  146  when rotor blades  100  are coupled within rotor disk  102 . In one embodiment, radial seal pin  194  is not inserted into leading edge radial seal pin slot  192 . In an alternative embodiment, forward cover plate  146  includes a first sealing assembly  168  and a second sealing assembly  170 . 
     Referring to  FIG. 3 , in the exemplary embodiment, during operation of gas turbine engine assembly  10 , combustor section  16  generates and channels combustion gases, represented by arrows  196 , to rotor assembly  22 . Combustion gases  196  contact rotor blades  100  causing rotor assembly  22  to rotate about drive shaft  32 . At least a portion of combustion gases  196  pass through adjacent forward cover plates  146 , around radial seal pin  194 , and into shank cavity  152 . First sealing assembly  168  and second sealing assembly  170  each facilitate preventing combustion gases  196  from passing through adjacent aft cover plates  148  causing an increase in a fluid pressure within shank cavity  152  that facilitates reducing a volume of combustion gases  196  entering shank cavity  152 . 
     The above-described methods and apparatus facilitate reducing an operating temperature of a rotor assembly. More specifically, the labyrinth seal defined between adjacent rotor blades facilitate reducing leakage of cooling fluid between adjacent rotor blades. In addition, the embodiments described herein facilitate increasing a back pressure of cooling fluid within a shank cavity, which facilitates increasing a flow of cooling fluid to the rotor blades to reduce an operating temperature of the rotor assembly. As such, the cost of maintaining the gas turbine engine system is facilitated to be reduced. 
     Exemplary embodiments of methods and apparatus for a turbine bucket assembly are described above in detail. The methods and apparatus are not limited to the specific embodiments described herein, but rather, components of systems and/or steps of the method may be utilized independently and separately from other components and/or steps described herein. For example, the methods and apparatus may also be used in combination with other combustion systems and methods, and are not limited to practice with only the gas turbine engine assembly as described herein. Rather, the exemplary embodiment can be implemented and utilized in connection with many other combustion system applications. 
     Although specific features of various embodiments of the invention may be shown in some drawings and not in others, this is for convenience only. Moreover, references to “one embodiment” in the above description are not intended to be interpreted as excluding the existence of additional embodiments that also incorporate the recited features. In accordance with the principles of the invention, any feature of a drawing may be referenced and/or claimed in combination with any feature of any other drawing. 
     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 have 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 languages of the claims.