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.
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.
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.
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.
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 r1 from dovetail 116 to a platform inner surface 174. First sealing assembly 168 and second sealing assembly 170 each extend a radial height r2 from dovetail 116 to platform inner surface 174. Radial height r2 is approximately the same height as radial height r1 of aft cover plate 148. In one embodiment, first sealing assembly 168 and/or second sealing assembly 170 extends the full radial height r1 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
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.