U.S. patent application number 12/860493 was filed with the patent office on 2012-02-23 for turbine bucket assembly and methods for assembling same.
Invention is credited to Michael James Fedor, David Martin Johnson.
Application Number | 20120045337 12/860493 |
Document ID | / |
Family ID | 45557455 |
Filed Date | 2012-02-23 |
United States Patent
Application |
20120045337 |
Kind Code |
A1 |
Fedor; Michael James ; et
al. |
February 23, 2012 |
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.
Inventors: |
Fedor; Michael James;
(Simpsonville, SC) ; Johnson; David Martin;
(Simpsonville, SC) |
Family ID: |
45557455 |
Appl. No.: |
12/860493 |
Filed: |
August 20, 2010 |
Current U.S.
Class: |
416/193A ;
29/889.21; 416/219R |
Current CPC
Class: |
F01D 11/006 20130101;
Y10T 29/49321 20150115 |
Class at
Publication: |
416/193.A ;
29/889.21; 416/219.R |
International
Class: |
F01D 11/00 20060101
F01D011/00; F01D 5/30 20060101 F01D005/30; B21K 25/00 20060101
B21K025/00 |
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.
Description
BACKGROUND OF THE INVENTION
[0001] The subject matter described herein relates generally to gas
turbine engines and, more particularly, to a bucket assembly for
use with a turbine engine.
[0002] 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.
[0003] 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
[0004] 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.
[0005] 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.
[0006] 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
[0007] FIG. 1 is schematic illustration of an exemplary known
turbine engine system.
[0008] 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.
[0009] FIG. 3 is an enlarged sectional view of a portion of the
rotor assembly shown in FIG. 2
[0010] FIG. 4 is a cross-sectional view of the rotor assembly shown
in FIG. 2.
DETAILED DESCRIPTION OF THE INVENTION
[0011] 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.
[0012] 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.
[0013] 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).
[0014] 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.
[0015] 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.
[0016] 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.
[0017] 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.
[0018] 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.
[0019] 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.
[0020] 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.
[0021] In the exemplary embodiment, aft cover plate 148 extends a
radial height r.sub.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.sub.2 from dovetail 116 to platform
inner surface 174. Radial height r.sub.2 is approximately the same
height as radial height r.sub.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.sub.1 of aft cover
plate 148.
[0022] 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.
[0023] 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.
[0024] 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.
[0025] 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.
[0026] 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.
[0027] 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.
[0028] 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.
[0029] 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.
[0030] 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.
* * * * *