TURBINE BUCKET ASSEMBLY AND METHODS FOR ASSEMBLING SAME

Abstract

A method for assembling a rotor assembly for use with a turbine engine. 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.

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 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 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.

Claims

1. A method for assembling a rotor assembly for use with a turbine engine, said method comprising:

providing at least two rotor blades that each include a shank extending between a dovetail and a platform, wherein each shank includes at least one cover plate that extends inwardly from the platform towards the dovetail, and an airfoil that extends outwardly from the platform;

coupling a first rotor blade to a rotor disk;

coupling a second rotor blade 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.

2. A method in accordance with claim 1, further comprising:

coupling a first sealing assembly to the first rotor blade cover plate; and

coupling a second sealing assembly to the second rotor blade cover plate to form a labyrinth seal path.

3. A method in accordance with claim 2, further comprising:

coupling a sealing extension to the first rotor blade cover plate to form the first sealing assembly; and

defining a sealing groove within the second rotor blade cover plate.

4. A method in accordance with claim 3, further comprising coupling an abradable surface to an outer surface of the sealing groove.

5. A method in accordance with claim 3, further comprising coupling a plurality of labyrinth teeth to the sealing extension such that a tortuous path is defined between the first sealing assembly and the second sealing assembly.

6. A method in accordance with claim 2, wherein the first sealing assembly extends between the platform and the dovetail.

7. A rotor blade for a turbine engine, said rotor blade comprising:

a platform comprising a radially outer surface and a radially inner surface;

an airfoil extending radially outwardly from said platform;

a dovetail adapted to be coupled to a rotor wheel;

a shank extending between said platform and said dovetail, said shank comprising at least one cover plate extending inwardly from said platform towards said dovetail; and

at least one sealing assembly coupled to said cover plate, said sealing assembly extending from said dovetail to said platform, said sealing assembly forms a seal path between said rotor blade and a circumferentially adjacent rotor blade.

8. A rotor blade in accordance with claim 7, wherein said sealing assembly comprises a sealing extension coupled to said cover plate, said sealing extension extending outwardly from said cover plate towards an adjacent rotor blade.

9. A rotor blade in accordance with claim 8, wherein said sealing extension comprises a plurality of labyrinth teeth extending outwardly from said sealing extension, said labyrinth teeth configured to form a tortuous path between said sealing extension and an adjacent rotor blade.

10. A rotor blade in accordance with claim 7, wherein said sealing assembly comprises a recessed sealing groove defined within said cover plate.

11. A rotor blade in accordance with claim 10, wherein said sealing groove comprises an abradable surface extending from an outer surface of said sealing groove.

12. A rotor blade in accordance with claim 7, further comprising a first sealing assembly coupled to said cover plate and an opposite second sealing assembly coupled said cover plate.

13. A rotor blade in accordance with claim 12, wherein said first sealing assembly comprises a sealing extension, said second sealing assembly comprises a recessed groove.

14. A rotor blade in accordance with claim 12, wherein said first sealing assembly and said second sealing assembly each comprise a sealing extension.

15. A rotor blade in accordance with claim 12, wherein said first sealing assembly and said second sealing assembly each comprise a recessed groove.

16. A gas turbine engine comprising:

a compressor;

a combustor coupled downstream from said compressor to receive at least some of the air discharged by said compressor;

a rotor shaft coupled to said compressor; and

a plurality of circumferentially-spaced rotor blades coupled to said rotor shaft, each of said plurality of rotor blades comprising:

a platform;

an airfoil extending radially outwardly from said platform;

a dovetail coupled to said rotor shaft;

a shank extending between said platform and said dovetail, said shank comprising at least one cover plate extending inwardly from said platform towards said dovetail; and

at least one sealing assembly coupled to said cover plate such that a seal path is defined between adjacent rotor blades.

17. A gas turbine engine in accordance with claim 16, wherein each of said plurality of rotor blades further comprises a first sealing assembly coupled to said cover plate and an opposite second sealing assembly coupled said cover plate.

18. A gas turbine engine in accordance with claim 17, wherein said first sealing assembly comprises a sealing extension, said second sealing assembly comprises a recessed groove.

19. A gas turbine engine in accordance with claim 17, wherein said first sealing assembly and said second sealing assembly each comprise a sealing extension.

20. A gas turbine engine in accordance with claim 17, wherein said first sealing assembly and said second sealing assembly each comprise a recessed groove.