U.S. patent application number 14/136403 was filed with the patent office on 2015-06-25 for snubber configurations for turbine rotor blades.
This patent application is currently assigned to General Electric Company. The applicant listed for this patent is General Electric Company. Invention is credited to Mark Andrew Jones, Joseph Anthony Weber.
Application Number | 20150176413 14/136403 |
Document ID | / |
Family ID | 53275478 |
Filed Date | 2015-06-25 |
United States Patent
Application |
20150176413 |
Kind Code |
A1 |
Weber; Joseph Anthony ; et
al. |
June 25, 2015 |
SNUBBER CONFIGURATIONS FOR TURBINE ROTOR BLADES
Abstract
A rotor blade for use in a turbine of a combustion turbine
engine. The rotor blade may include an airfoil having a concave
pressure sidewall and a convex suction sidewall extending axially
between corresponding leading and trailing edges and radially
between the root and an outboard tip. The rotor blade may further
include dual snubber shrouds positioned on the airfoil. Each of the
dual snubber shrouds may be configured to engage a corresponding
snubber shroud on at least one neighboring rotor blade upon
installation.
Inventors: |
Weber; Joseph Anthony;
(Simpsonville, SC) ; Jones; Mark Andrew; (Greer,
SC) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company |
Schenectady |
NY |
US |
|
|
Assignee: |
General Electric Company
Schenectady
NY
|
Family ID: |
53275478 |
Appl. No.: |
14/136403 |
Filed: |
December 20, 2013 |
Current U.S.
Class: |
416/212A |
Current CPC
Class: |
Y02T 50/60 20130101;
Y02T 50/673 20130101; F01D 5/225 20130101; Y02T 50/671
20130101 |
International
Class: |
F01D 5/22 20060101
F01D005/22 |
Claims
1. A rotor blade for use in a turbine of a combustion turbine
engine, the rotor blade comprising an airfoil that extends from a
connection with a root, the airfoil including a concave pressure
sidewall and a convex suction sidewall extending axially between
corresponding leading and trailing edges and radially between the
root and an outboard tip, the rotor blade further comprising: dual
snubber shrouds positioned on the airfoil; wherein each of the dual
snubber shrouds is configured to engage a corresponding dual
snubber shroud on at least one neighboring rotor blade upon
installation.
2. The rotor blade according to claim 1, wherein both of the dual
snubber shrouds comprises a position at or outboard of a radial
mid-region of the airfoil; and wherein each of the dual snubber
shrouds comprise an axial thickness that is less than half of an
axial thickness that the airfoil has at a radial height of each of
the snubber shrouds.
3. The rotor blade according to claim 2, wherein the dual snubber
shrouds include an inboard snubber shroud and an outboard snubber
shroud, the inboard snubber shroud comprising an inboard position
relative to the outboard snubber shroud; wherein the inboard
snubber shroud comprises a circumferentially extending projection
from at least one of the pressure sidewall and the suction sidewall
of the airfoil; wherein the outboard snubber shroud comprises a
circumferentially extending projection from at least one of the
pressure sidewall and the suction sidewall of the airfoil; and
wherein the rotor blade comprises a turbine rotor blade.
4. The rotor blade according to claim 3, wherein: the inboard
snubber shroud comprises circumferentially extending projections
from each of the pressure sidewall and the suction sidewall of the
airfoil; and the outboard snubber shroud comprises
circumferentially extending projections from each of the pressure
sidewall and the suction sidewall of the airfoil.
5. The rotor blade according to claim 4, wherein: each of the
inboard snubber shroud and the outboard snubber shroud comprises
pressure side contact faces at distal ends of each of the
circumferential projections extending from the pressure sidewall of
the airfoil; and each of the inboard snubber shroud and the
outboard snubber shroud comprises suction side contact faces at
distal ends of each of the circumferential projections extending
from the suction sidewall of the airfoil.
6. The rotor blade according to claim 5, wherein the inboard
snubber shroud and the outboard snubber shroud are configured such
that the pressure side contact faces correspond to the suction side
contact faces so to form an interface therebetween when the rotor
blade comprises an installed position between neighboring rotor
blades having a same design as the rotor blade.
7. The rotor blade according to claim 5, wherein the inboard
snubber shroud and the outboard snubber shroud are configured such
that the pressure side contact faces engage suction side contact
faces of a first neighboring rotor blade if: the first neighboring
rotor blade comprises a same design as the rotor blade; and the
first neighboring rotor blade and the rotor blade comprise a
predetermined installed position relative to each other; wherein
the inboard snubber shroud and the outboard snubber shroud are
configured such that the suction side contact faces engage pressure
side contact faces of a second neighboring rotor blade if: the
second neighboring rotor blade comprises the same design as the
rotor blade; and the second neighboring rotor blade and the rotor
blade comprise a predetermined installed position relative to each
other.
8. The rotor blade according to claim 4, wherein the outboard
snubber shroud comprises a position near the outboard tip of the
airfoil and the inboard snubber shroud comprises a position near
the radial mid-region of the airfoil.
9. The rotor blade according to claim 4, wherein the outboard
snubber shroud comprises a shroud positioned just inside of the
outboard tip of the airfoil and the inboard snubber shroud
comprises a shroud positioned near a radial mid-point of the
airfoil.
10. The rotor blade according to claim 4, wherein the inboard
snubber shroud comprises one disposed within a first range of
radial heights defined on the airfoil, wherein the first range
includes an inboard boundary at 25% of a radial height of the
airfoil and an outboard boundary at 75% of the radial height of the
airfoil; and wherein the outboard snubber shroud comprises one
disposed within a second range of radial heights defined on the
airfoil, wherein the second range includes an inboard boundary at
60% of the radial height of the airfoil.
11. The rotor blade according to claim 10, wherein the inboard
boundary of the first range comprises 40% of a radial height of the
airfoil and the outboard boundary of the first range comprises 60%
of the radial height of the airfoil; and wherein the inboard
boundary of the second range comprises 75% of the radial height of
the airfoil and an outboard boundary of the second range comprises
95% of the radial height of the airfoil.
12. The rotor blade according to claim 10, wherein the inboard
boundary of the first range comprises 40% of a radial height of the
airfoil and the outboard boundary of the first range comprises 60%
of the radial height of the airfoil; and wherein the inboard
boundary of the second range comprises 90% of the radial height of
the airfoil.
13. The rotor blade according to claim 12, wherein an outboard tip
of the airfoil comprises a winglet.
14. The rotor blade according to claim 13, wherein the outboard
snubber shroud is radial spaced apart from the winglet.
15. The rotor blade according to claim 14, wherein the winglet
comprises an outboard flaring disposed within a narrow radial
section of the airfoil, the narrow radial section having a side
contiguous with the outboard tip of the airfoil.
16. The rotor blade according to claim 15, wherein the outboard
shroud extends from the winglet.
17. The rotor blade according to claim 15, wherein the winglet
includes an inboard edge and an outboard edge, and wherein the
outboard flaring includes the winglet having a smaller
cross-sectional area at the inboard edge and a larger
cross-sectional airfoil area at an outboard edge of winglet;
wherein the outboard edge of the winglet comprises the outboard tip
of the airfoil and the inboard edge of the winglet is spaced a
fixed distance from the outboard edge of the winglet about a
circumference of the airfoil.
18. The rotor blade according to claim 17, wherein the outboard
edge of the winglet comprises a sharp edge defined about the
circumference to the airfoil; and wherein the outboard flaring of
the winglet comprises a concave surface profile.
19. A gas turbine engine having a turbine that comprises a row of
circumferentially spaced rotor blades, wherein each of the rotor
blades includes an airfoil extending from a root that connects to a
rotor disc, the airfoil including a pressure sidewall and a suction
sidewall extending axially between a leading edge and a trailing
edges and radially between a platform that forms an outboard
boundary of the root and an outboard tip of the airfoil, the
airfoil of each of the rotor blades further comprising: dual
snubber shrouds, an inboard snubber shroud and an outboard snubber
shroud; wherein each of the snubber shrouds comprises a position
between the platform and the outboard tip of the airfoil and a
configuration that connects each of the rotor blades to neighboring
rotor blades positioned to each side; and wherein the outboard
snubber shroud comprises a shroud positioned near the outboard tip
of the airfoil and the inboard snubber shroud comprises a shroud
positioned near a radial mid-region of the airfoil.
20. The gas turbine engine according to claim 19, wherein the
inboard snubber shroud comprises circumferentially extending
projections from each of the pressure sidewall and the suction
sidewall of the airfoil, and the outboard snubber shroud comprises
circumferentially extending projections from each of the pressure
sidewall and the suction sidewall of the airfoil; and wherein each
of the inboard snubber shroud and the outboard snubber shroud
comprises pressure side contact faces at distal ends of each of the
circumferential projections extending from the pressure sidewall of
the airfoil, and each of the inboard snubber shroud and the
outboard snubber shroud comprises suction side contact faces at
distal ends of each of the circumferential projections extending
from the suction sidewall of the airfoil.
Description
BACKGROUND OF THE INVENTION
[0001] The present application relates generally to apparatus,
methods and/or systems concerning the design and manufacture of
turbine rotor blades. More specifically, but not by way of
limitation, the present application relates to apparatus and
assemblies pertaining to turbine rotor blades having multiple
snubber shrouds.
[0002] In a combustion turbine engine, it is well known that air
pressurized in a compressor is used to combust a fuel in a
combustor to generate a flow of hot combustion gases, whereupon
such gases flow downstream through one or more turbines so that
energy can be extracted therefrom. In accordance with such a
turbine, generally, rows of circumferentially spaced rotor blades
extend radially outwardly from a supporting rotor disc. Each rotor
blade typically includes a dovetail that permits assembly and
disassembly of the blade in a corresponding dovetail slot in the
rotor disc, as well as an airfoil that extends radially outwardly
from the dovetail and interacts with the flow of the working fluid
through the engine. The airfoil has a concave pressure side and
convex suction side extending axially between corresponding leading
and trailing edges, and radially between a root and a tip. It will
be understood that the blade tip is spaced closely to a radially
outer stationary surface for minimizing leakage therebetween of the
combustion gases flowing downstream between the turbine blades.
[0003] Shrouds at the tip of the airfoil or "tip shrouds" often are
implemented on aft stages or rotor blades to provide a point of
contact at the tip, manage bucket frequencies, enable a damping
source (i.e., by connecting the tips of neighboring rotor blades),
and to reduce the over-tip leakage of the working fluid. Given the
length of the rotor blades in the aft stages, the damping function
of the tip shrouds provides a significant benefit to durability.
However, taking full advantage of the benefits is difficult
considering the weight that the tip shroud adds to the assembly and
the other design criteria, which include enduring thousands of
hours of operation exposed to high temperatures and extreme
mechanical loads. Thus, while large tip shrouds are desirable
because of the effective manner in which they seal the gas path and
robust connection they may form between neighboring rotor blades,
one of ordinary skill in the art will appreciate that larger tip
shrouds are troublesome because of the increased pull loads on the
rotor disc, particularly at the base of the airfoil because it must
support the entire load of blade.
[0004] Another consideration is that the output and efficiency of
gas turbine engines improve as the size of the engine and, and more
specifically, the amount of air able to pass through it increase.
The size of the engine, however, is limited by the operable length
of the turbine blades, with longer turbine rotor blades enabling
enlargement of the flow path through engine. Longer rotor blades,
though, incur increased mechanical loads, which place further
demands on the blades and the rotor disc that holds them. Longer
rotor blades also decrease the natural vibrational frequencies of
the blades during operation, which increases the vibratory response
of the rotor blades. This additional vibratory load place even
greater demands on rotor blade design, which may further shorten
the life of the component and, in some cases, may cause vibratory
loads that damage other functions of the turbine engine. One way to
address the vibratory load of longer rotor blades is through the
use of shrouds that connect adjacent rotor blades to each other. As
mentioned, though, the added weight of the shroud may negate much
of the benefit.
[0005] One way to address this is to position the shroud lower on
the airfoil of the rotor blade. That is, instead of adding the
shroud to the tip of the rotor blade, the shroud is positioned near
the middle radial portion of the airfoil. As used herein, such a
shroud will be referred to as a "snubber shroud." At this lower (or
more inboard) radius, the mass of the shroud causes a reduced level
of stress to the rotor blade. However, this type of shroud leaves a
portion of the airfoil of the rotor blade unrestrained (i.e., that
portion of the airfoil that extends outboard of the snubber
shroud). This cantilevered portion of the airfoil typically results
in lower frequency vibration and increased vibratory loads, which
may be damaging to the engine. Accordingly, a novel rotor blade
design that reduced or limited these loads would have value in the
market for such products.
BRIEF DESCRIPTION OF THE INVENTION
[0006] The present application thus describes a rotor blade for use
in a turbine of a combustion turbine engine. The rotor blade may
include an airfoil that extends from a connection with a root. The
airfoil may include a concave pressure sidewall and a convex
suction sidewall extending axially between corresponding leading
and trailing edges and radially between the root and an outboard
tip. The rotor blade may further include dual snubber shrouds
positioned on the airfoil. Each of the dual snubber shrouds may be
configured to engage a corresponding snubber shroud on at least one
neighboring rotor blade upon installation.
[0007] The present application further describes a gas turbine
engine having a turbine that includes a row of circumferentially
spaced rotor blades. Each of the rotor blades may include an
airfoil extending from a root that connects to a rotor disc. The
airfoil may include a pressure sidewall and a suction sidewall
extending axially between a leading edge and a trailing edges and
radially between a platform that forms an outboard boundary of the
root and an outboard tip of the airfoil. The airfoil of each of the
rotor blades may further include dual snubber shrouds, an inboard
snubber shroud and an outboard snubber shroud. Each of the snubber
shrouds may have a position between the platform and the outboard
tip of the airfoil and a configuration that connects each of the
rotor blades to neighboring rotor blades positioned to each side.
The outboard snubber shroud may be positioned near the outboard tip
of the airfoil, and the inboard snubber shroud may be positioned
near a radial mid-region of the airfoil.
[0008] These and other features of the present application will
become apparent upon review of the following detailed description
of the preferred embodiments when taken in conjunction with the
drawings and the appended claims.
BRIEF DESCRIPTION OF THE DRAWINGS
[0009] These and other features of this invention will be more
completely understood and appreciated by careful study of the
following more detailed description of exemplary embodiments of the
invention taken in conjunction with the accompanying drawings, in
which:
[0010] FIG. 1 is a schematic representation of an exemplary
combustion turbine engine in which embodiments of the present
application may be used;
[0011] FIG. 2 is a sectional view of the compressor in the
combustion turbine engine of FIG. 1;
[0012] FIG. 3 is a sectional view of the turbine in the combustion
turbine engine of FIG. 1;
[0013] FIG. 4 is a perspective view of an exemplary turbine rotor
blade having a tip shroud of conventional design;
[0014] FIG. 5 is a perspective view of an exemplary turbine rotor
blade having a conventional mid-span snubber;
[0015] FIG. 6 is a perspective view of installed turbine rotor
blades connected via a conventional mid-span snubber;
[0016] FIG. 7 is a top view of installed turbine rotor blades
connected via a conventional mid-span snubber;
[0017] FIG. 8 is a side view of an exemplary turbine rotor blade
and stationary shroud assembly in which the rotor blade includes a
conventional tip shroud;
[0018] FIG. 9 is a side view of an exemplary turbine rotor blade
and stationary shroud assembly in which the rotor blade includes
dual snubbers according to an exemplary embodiment of the present
invention;
[0019] FIG. 10 is a perspective view of the outboard portion of the
airfoil of FIG. 9;
[0020] FIG. 11 is a side view of the outboard portion of an airfoil
having an outboard snubber and a tip winglet according to an
exemplary embodiment of the present invention;
[0021] FIG. 12 is a side view of the outboard portion of an airfoil
having an outboard snubber and a tip winglet according to an
alternative embodiment of the present invention;
[0022] FIG. 13 is a perspective view of the airfoil tip of the
turbine rotor blade of FIG. 12.
DETAILED DESCRIPTION OF THE INVENTION
[0023] While the following examples of the present invention may be
described in reference to particular types of turbine engine, those
of ordinary skill in the art will appreciate that the present
invention may not be limited to such use and applicable to other
types of turbine engines, unless specifically limited therefrom.
Further, it will be appreciated that in describing the present
invention, certain terminology may be used to refer to certain
machine components within the gas turbine engine. Whenever
possible, common industry terminology will be used and employed in
a manner consistent with its accepted meaning. However, such
terminology should not be narrowly construed, as those of ordinary
skill in the art will appreciate that often a particular machine
component may be referred to using differing terminology.
Additionally, what may be described herein as being single
component may be referenced in another context as consisting of
multiple components, or, what may be described herein as including
multiple components may be referred to elsewhere as a single one.
As such, in understanding the scope of the present invention,
attention should not only be paid to the particular terminology,
but also the accompanying description, context, as well as the
structure, configuration, function, and/or usage of the component,
particularly as may be provided in the appended claims.
[0024] Several descriptive terms may be used regularly herein, and
it may be helpful to define these terms at the onset of this
section. Accordingly, these terms and their definitions, unless
stated otherwise, are as follows. As used herein, "downstream" and
"upstream" are terms that indicate direction relative to the flow
of a fluid, such as, for example, the working fluid through the
compressor, combustor and turbine sections of the gas turbine, or
the flow coolant through one of the component systems of the
engine. The term "downstream" corresponds to the direction of fluid
flow, while the term "upstream" refers to the direction opposite or
against the direction of fluid flow. The terms "forward" and "aft",
without any further specificity, refer to directions relative to
the orientation of the gas turbine, with "forward" referring to the
forward or compressor end of the engine, and "aft" referring to the
aft or turbine end of the engine. Additionally, given a gas turbine
engine's configuration about a central axis as well as this same
type of configuration in some component systems, terms describing
position relative to an axis likely will be used. In this regard,
it will be appreciated that the term "radial" refers to movement or
position perpendicular to an axis. Related to this, it may be
required to describe relative distance from the central axis. In
this case, for example, if a first component resides closer to the
center axis than a second component, it will be stated herein that
the first component is "radially inward" or "inboard" of the second
component. If, on the other hand, the first component resides
further from the axis than the second component, it may be stated
herein that the first component is "radially outward" or "outboard"
of the second component. Additionally, it will be appreciated that
the term "axial" refers to movement or position parallel to an
axis. And, finally, the term "circumferential" refers to movement
or position around an axis.
[0025] By way of background, referring now to the figures, FIGS. 1
through 3 illustrate an exemplary combustion turbine engine in
which embodiments of the present application may be used. It will
be understood by those skill in the art that the present invention
is not limited to this type of usage. As stated, the present
invention may be used in combustion turbine engines, such as the
engines used in power generation and airplanes, steam turbine
engines, and other type of rotary engines. FIG. 1 is a schematic
representation of a combustion turbine engine 10. In general,
combustion turbine engines operate by extracting energy from a
pressurized flow of hot gas produced by the combustion of a fuel in
a stream of compressed air. As illustrated in FIG. 1, combustion
turbine engine 10 may be configured with an axial compressor 11
that is mechanically coupled by a common shaft or rotor to a
downstream turbine section or turbine 13, and a combustor 12
positioned between the compressor 11 and the turbine 12.
[0026] FIG. 2 illustrates a view of an exemplary multi-staged axial
compressor 11 that may be used in the combustion turbine engine of
FIG. 1. As shown, the compressor 11 may include a plurality of
stages. Each stage may include a row of compressor rotor blades 14
followed by a row of compressor stator blades 15. Thus, a first
stage may include a row of compressor rotor blades 14, which rotate
about a central shaft, followed by a row of compressor stator
blades 15, which remain stationary during operation. The compressor
stator blades 15 generally are circumferentially spaced one from
the other and fixed about the axis of rotation. The compressor
rotor blades 14 are circumferentially spaced and attached to the
shaft; when the shaft rotates during operation, the compressor
rotor blades 14 rotate about it. As one of ordinary skill in the
art will appreciate, the compressor rotor blades 14 are configured
such that, when spun about the shaft, they impart kinetic energy to
the air or fluid flowing through the compressor 11. The compressor
11 may have other stages beyond the stages that are illustrated in
FIG. 2. Additional stages may include a plurality of
circumferential spaced compressor rotor blades 14 followed by a
plurality of circumferentially spaced compressor stator blades
15.
[0027] FIG. 3 illustrates a partial view of an exemplary turbine
section or turbine 13 that may be used in the combustion turbine
engine of FIG. 1. The turbine 13 also may include a plurality of
stages. Three exemplary stages are illustrated, but more or less
stages may present in the turbine 13. A first stage includes a
plurality of turbine buckets or turbine rotor blades 16, which
rotate about the shaft during operation, and a plurality of nozzles
or turbine stator blades 17, which remain stationary during
operation. The turbine stator blades 17 generally are
circumferentially spaced one from the other and fixed about the
axis of rotation. The turbine rotor blades 16 may be mounted on a
turbine wheel or disc (not shown) for rotation about the shaft (not
shown). A second stage of the turbine 13 also is illustrated. The
second stage similarly includes a plurality of circumferentially
spaced turbine stator blades 17 followed by a plurality of
circumferentially spaced turbine rotor blades 16, which are also
mounted on a turbine wheel for rotation. A third stage also is
illustrated, and similarly includes a plurality of turbine stator
blades 17 and rotor blades 16. It will be appreciated that the
turbine stator blades 17 and turbine rotor blades 16 lie in the hot
gas path of the turbine 13. The direction of flow of the hot gases
through the hot gas path is indicated by the arrow. As one of
ordinary skill in the art will appreciate, the turbine 13 may have
other stages beyond the stages that are illustrated in FIG. 3. Each
additional stage may include a row of turbine stator blades 17
followed by a row of turbine rotor blades 16.
[0028] In use, the rotation of compressor rotor blades 14 within
the axial compressor 11 may compress a flow of air. In the
combustor 12, energy may be released when the compressed air is
mixed with a fuel and ignited. The resulting flow of hot gases from
the combustor 12, which may be referred to as the working fluid, is
then directed over the turbine rotor blades 16, the flow of working
fluid inducing the rotation of the turbine rotor blades 16 about
the shaft. Thereby, the energy of the flow of working fluid is
transformed into the mechanical energy of the rotating blades and,
because of the connection between the rotor blades and the shaft,
the rotating shaft. The mechanical energy of the shaft may then be
used to drive the rotation of the compressor rotor blades 14, such
that the necessary supply of compressed air is produced, and also,
for example, a generator to produce electricity.
[0029] FIG. 4 is a perspective view of an exemplary turbine rotor
blade 16 that has a tip shroud 37 of conventional design. The
turbine rotor blade 16 generally includes a root 21, which may
include means by which the rotor blade 16 attaches to a rotor disc
41 (as shown in FIG. 6), such as an axial dovetail configured for
mounting in a corresponding dovetail slot in the perimeter of the
rotor disc 41. The root 21 may include a shank that extends between
the dovetail and a platform 24, with the platform 24 being disposed
at the junction of the airfoil 25 and the root 21. The platform 24
defines a portion of the inboard boundary of the flowpath through
the turbine engine 10. The airfoil 25 is the active component of
the rotor blade 16 that intercepts the flow of the working fluid
and induces the rotor disc 41 to rotate. As illustrated, at the
outboard tip of the rotor blade 16, the tip shroud 37 may be
positioned. The tip shroud 37 essentially is an axially and
circumferentially extending planar component that is perched atop
the airfoil 25 and supported by it. As shown, positioned along the
top of the tip shroud 37 may be one or more seal rails 38.
Generally, seal rails 38 project radially outward from the outboard
surface of the tip shroud 37 and extend circumferentially between
opposite ends of the tip shroud 37 in the general direction of
rotation. Seal rails 38 are formed to deter the flow of working
fluid through the gap between the tip shroud 37 and the inner
surface of the surrounding stationary components of the turbine 13.
As discussed in more detail below, tip shrouds 37 may be formed
with contact faces 55 such that the shrouds on adjacent rotor
blades contact or engage each other, which typically damps
vibration in the assembly and prolong the life of the rotor blade
16. (Note that while a preferred embodiment of the present
application is directed toward turbine rotor blades 16, it will be
understood that aspects of the present invention may be applicable
to compressor rotor blades 14 and that, unless otherwise stated,
the present invention should be understood to apply to each type of
rotor blade 14, 16.)
[0030] FIG. 5 provides a perspective view of an exemplary turbine
rotor blade 16 which has a snubber shroud 51 consistent with one
that might be used with rotor blades 16 having an internal
structural configurations in accordance with the present invention,
as discussed in detail below. As is known in the art, a snubber or
snubber shroud 51 such as the one shown may be used to connect
adjacent rotor blades 16. The linking of adjacent rotor blades 16
may occur between a shroud-to-shroud interface 54 (which is shown
in FIG. 7) at which a pressure side contact face 55 and a suction
side contact face 56 contact each other. This linking of rotor
blades 16 in this manner tends to increase the natural frequency of
the assembly and damp operational vibrations, which means rotor
blades 16 are subject to less mechanical stress during operation
and degrade more slowing. Shrouds 51, however, add weight to the
assembly, which tends to negate some of these benefits,
particularly when the shroud is located at the outboard tip 41 of
the rotor blade 16. As mentioned above, one way to lessen the
impact of the added weight of the shroud is to position the shroud
lower on the airfoil 25, as shown in FIG. 5. At this lower (or more
inboard) radius, the mass of the shroud 51 causes a reduction in
the applied stress to the rotor blade. However, a snubber shroud
leaves a portion of the airfoil 25 unrestrained, i.e., the portion
of the airfoil 25 that extends outboard of the snubber shroud 51,
and this cantilevered portion of the airfoil 25 results in a lower
natural frequency and increased vibratory response during
operation, which, as stated, increases may damage the rotor blades
and the engine.
[0031] FIG. 6 is a perspective view of rotor blades 16 having
snubber shrouds 51 as they might be arranged in an installed
condition. FIG. 7 provides a top view of the same installed
assembly. As shown, the snubber shrouds 51 are configured to link
or engage the shrouds 51 of the rotor blades 16 that are adjacent
to them. This linking or engagement may occur at shroud-to-shroud
interfaces 54 between the pressure side contact face 55 and the
suction side contact face 56, as illustrated.
[0032] FIG. 8 is a side view of an exemplary turbine rotor blade 16
having a tip shroud 37 according to a conventional design. It will
be appreciated that the mass tip shrouds add to the outboard
portion of the airfoil significantly increases the mechanical
forces acting on a rotating blade during operation. In order to
sustain these additional stresses, conventional rotor blades are
designed with airfoil structures that are strengthened by
increasing the width of the airfoil base (i.e., the distance
indicated by reference numeral 61). Over time, though, the
increased pull that results from weight of the tip shroud results
in airfoil creep, with is a general stretching along its length
(i.e., the distance indicated by reference numeral 63). This
increased elongation over the life of the rotor blade must be taken
into account when designing the clearances between the outboard
tips of rotor blade and the stationary structure that surrounds it
(i.e., the gap indicated by reference numeral 64). It will be
appreciated that this necessarily results in wider gap clearances,
which, because wider clearances result in increased leakage levels,
negatively impacts the efficiency of the engine.
[0033] FIGS. 9 through 13 illustrate several aspects of the present
invention. As illustrated in FIG. 9, the present invention
describes an airfoil 25 having dual snubber shrouds: an outboard
snubber shroud 52 and an inboard snubber shroud 53. The benefits of
this arrangement are several, including an overall reduced tip mass
as some of that mass is relocated closer to the axis of rotation,
which reduces mechanical stress on the airfoil. This reduction in
stress allows a reduction in the width of the airfoil, as indicated
in FIG. 9. Additionally, the reduction in stress enables tighter
initial gap clearances between the rotor blade and stationary
structure, as the airfoil will experience less elongation during
operation. The reduced mechanical pull on the rotor blade also may
allow longer blade life, as well as smaller rotor dovetail widths,
which may reduce costs by further reducing the overall size of the
rotor blade. In addition, the present invention results in tighter
clearances and aerodynamic performance compared to conventional
scalloped tip shrouds, which may be further improved with the
addition of a winglet 70, as discussed below.
[0034] The dual snubber shrouds of the present invention, as
mentioned, may include an inboard snubber shroud 53 and an outboard
snubber shroud 52, with the inboard snubber shroud 53 being
positioned radially inward relative to the outboard snubber shroud
52. As shown, each of the snubber shrouds may have a narrower axial
profile than other conventional shrouds, which, in preferred
embodiments is less than half of the axial width of airfoil at the
radial height of the snubber shroud.
[0035] The inboard snubber shroud 53 may be configured as a
circumferentially extending projection that protrudes from one or
both of the pressure sidewall 26 and the suction sidewall 27 of the
airfoil 25. Similarly, the outboard snubber shroud 52 may be
configured as a circumferentially extending projection that
protrudes from one or both of the pressure sidewall 26 and the
suction sidewall 27 of the airfoil 25. As previously discussed,
each of the snubber shrouds may be configured to engage a snubber
shroud formed on one or both neighboring rotor blades upon
installation. It will be appreciated that the dual points of
contact that the present invention enables may be used to
advantageously limit the vibratory response of the rotor blades
during operation. Accordingly, each of the inboard snubber shroud
53 and the outboard snubber shroud 52 may include pressure side
contact faces 55 at distal ends of each circumferential projection
on the pressure side of the airfoil 25. Likewise, each of the
inboard snubber shroud 53 and the outboard snubber shroud 52 may
include suction side contact faces 56 at distal ends of each
circumferential projection on the suction side of the airfoil 25.
Given this configuration, it will be appreciated that the two
pressure side contact faces 55 of a particular airfoil 25 may
engage the two suction side contact faces 56 of the neighboring
airfoil 25 positioned to one side of it, and the two suction side
contact faces 56 may engage the two pressure side contact faces 55
of the neighboring airfoil 25 position to the other side of it.
[0036] As indicated in FIG. 9, the outboard snubber shroud 52 may
be positioned near the outboard tip 41 of the airfoil 25, while the
inboard snubber shroud 53 may be positioned near the radial
mid-region of the airfoil 25. In an alternative embodiment, the
outboard snubber shroud 52 is positioned just inside of the
outboard tip 41 of the airfoil 25, and the inboard snubber shroud
53 is positioned at approximately the radial mid-point of the
airfoil 25. In another embodiment, the radial positioning of the
inboard and outboard snubber shrouds 52 is defined within a range
of radial heights defined relative to the airfoil 25. In one such
embodiment, the inboard snubber shroud 53 may be positioned within
a range of radial heights defined between an inboard boundary at
25% of a radial height of the airfoil 25 and an outboard boundary
at 75% of the radial height of the airfoil 25, and the outboard
snubber shroud 52 may be positioned outside of an inboard boundary
at 60% of the radial height of the airfoil 25. In an alternative
embodiment, the inboard snubber shroud 53 may be positioned within
a range of radial heights defined between an inboard boundary at
40% of a radial height of the airfoil 25 and an outboard boundary
at 60% of the radial height of the airfoil 25, and the outboard
snubber shroud 52 may be positioned within a range of radial
heights defined between an inboard boundary at 75% of the radial
height of the airfoil 25 and an outboard boundary that 95% of the
radial height of the airfoil 25. In another preferred embodiment,
the inboard snubber shroud 53 may be positioned within a range of
radial heights defined between an inboard boundary at 40% of a
radial height of the airfoil 25 and an outboard boundary at 60% of
the radial height of the airfoil 25, and the outboard snubber
shroud 52 may be positioned outside of an inboard boundary at 90%
of the radial height of the airfoil 25.
[0037] According to other embodiments of the present invention, as
illustrated in FIGS. 10 through 13, the outboard tip 41 of the
airfoil 25 may include a winglet 70. As used herein, a winglet 70
includes a flaring of the outboard tip 41 of the airfoil 25. As
indicated in FIGS. 11 and 12, the flaring occurs within a narrow
radial section that, on one side, is contiguous to the outboard tip
41. It will be appreciated that the winglet 70 may be described as
including an inboard edge 71 and an outboard edge 72. Described in
this manner, the outboard flaring of the winglet 70 may be oriented
such that the cross-sectional area at the inboard edge 71 of the
winglet 70 is smaller than that of the outboard edge 72 of winglet
70. As indicated, the inboard edge 71 of the winglet 70 may be
offset a constant distance from the outboard edge 72, which as
stated, may be the outboard tip 41 of the airfoil 25.
[0038] As illustrated in FIG. 11, the outboard snubber shroud 52
may be radial spaced apart from the winglet 70. In this case, the
outboard snubber shroud 52 is not connected to the winglet 70 and
each are separated by a radial gap. In an alternative preferred
embodiment, as illustrated in FIGS. 12 and 13, the outboard snubber
shroud 52 may be incorporated into the winglet 70, i.e., the
outboard snubber shroud 52 may extend from the winglet 70 and be
connected to the winglet 70 at its base. As also shown, preferred
embodiments of the winglet 70 include the outboard edge having a
sharp edge that extends about the circumference to the airfoil 25.
Further, as indicated most clearly in FIGS. 12 and 13, the outboard
flaring of the winglet 70 preferably includes a concave surface
profile 73. Other profiles, such as linear, are also possible.
[0039] It will be appreciated that, pursuant to the several
embodiments discussed above, the present invention provides a
manner by which the vibratory response of turbine rotor blades may
be reduced so to limit damaging vibrational mechanical loads while
also allowing for improved aerodynamic/leakage preventing
performance. That is, it will be understood that, according to the
present invention, natural frequencies of the rotor blade structure
may be raised and harmful vibratory responses avoided, thereby
enabling longer turbine blades, which, in turn, may be used to
enable larger turbine engines having greater output and efficiency.
Additionally, the reduced tip mass enabled by the present invention
and resulting mechanical pull may allow for a tighter clearance
between the rotor blade and the surrounding stationary
structure.
[0040] While the invention has been described in connection with
what is presently considered to be the most practical and preferred
embodiment, it is to be understood that the invention is not to be
limited to the disclosed embodiment, but on the contrary, is
intended to cover various modifications and equivalent arrangements
included within the spirit and scope of the appended claims.
* * * * *