U.S. patent application number 13/725617 was filed with the patent office on 2014-09-11 for turbine rotor blades having mid-span shrouds.
This patent application is currently assigned to General Electric Company. The applicant listed for this patent is General Electric Company. Invention is credited to Bradley Taylor Boyer.
Application Number | 20140255207 13/725617 |
Document ID | / |
Family ID | 49956438 |
Filed Date | 2014-09-11 |
United States Patent
Application |
20140255207 |
Kind Code |
A1 |
Boyer; Bradley Taylor |
September 11, 2014 |
TURBINE ROTOR BLADES HAVING MID-SPAN SHROUDS
Abstract
A rotor blade for use in a turbine of a combustion turbine
engine is described. The rotor blade may include an airfoil that
extends from a connection with a root. The rotor blade may further
include a mid-span shroud configured to engage a corresponding
mid-span shroud on at least one neighboring rotor blades during
operation. Outboard of the mid-span shroud, the airfoil may include
an outboard region that is substantially hollow, and inboard of the
mid-span shroud, the airfoil may include an inboard region that is
substantially solid.
Inventors: |
Boyer; Bradley Taylor;
(Greenville, SC) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company; |
|
|
US |
|
|
Assignee: |
General Electric Company
Schenectady
NY
|
Family ID: |
49956438 |
Appl. No.: |
13/725617 |
Filed: |
December 21, 2012 |
Current U.S.
Class: |
416/97R ;
416/196R |
Current CPC
Class: |
Y02T 50/60 20130101;
Y02T 50/671 20130101; F05D 2260/96 20130101; F01D 5/225 20130101;
F01D 5/186 20130101; Y02T 50/676 20130101; F01D 5/18 20130101 |
Class at
Publication: |
416/97.R ;
416/196.R |
International
Class: |
F01D 5/22 20060101
F01D005/22; F01D 5/18 20060101 F01D005/18 |
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: a
mid-span shroud configured to engage a corresponding mid-span
shroud on at least one neighboring rotor blade during operation;
wherein: outboard of the mid-span shroud, the airfoil includes an
outboard region that is substantially hollow; and inboard of the
mid-span shroud, the airfoil includes an inboard region that is
substantially solid.
2. The rotor blade of claim 1, wherein the mid-span shroud includes
a circumferentially extending projection from at least one of the
pressure sidewall and the suction sidewall of the airfoil.
3. The rotor blade of claim 1, wherein the mid-span shroud
comprises a circumferential projection from each of the pressure
sidewall and the suction sidewall of the airfoil; wherein the
mid-span shroud comprises a pressure side contact face at a distal
end of the circumferential projection from the pressure sidewall
and a suction side contact face at a distal end of the
circumferential projection from the pressure sidewall; wherein,
upon installation, the pressure side contact face is configured to
form a first shroud-to-shroud interface with the suction side
contact face of a first neighboring rotor blade of a same design as
the rotor blade; and wherein, upon installation, the suction side
contact face is configured to form a second shroud-to-shroud
interface with the pressure side contact face of a second
neighboring rotor blade of the same design as the rotor blade.
4. The rotor blade of claim 3, wherein the mid-span shroud
comprises a shroud that is positioned inboard of an outboard tip of
the airfoil and outboard of a platform of the rotor blade.
5. The rotor blade of claim 3, wherein the mid-span shroud
comprises a shroud that is disposed within a range of positions on
the airfoil; and wherein the range of positions is defined between
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.
6. The rotor blade of claim 3, wherein the mid-span shroud
comprises a shroud that is disposed within a range of positions on
the airfoil; and wherein the range of positions is defined between
an inboard boundary at 33% of a radial height of the airfoil and an
outboard boundary at 66% of the radial height of the airfoil.
7. The rotor blade of claim 5, wherein the outboard region of
airfoil comprises a portion of the airfoil that is outboard of the
mid-span shroud, and wherein the inboard region of the airfoil
comprises a portion of the airfoil that is inboard of the mid-span
shroud; wherein a hollow region comprises a region within the
airfoil that is hollow; and wherein a solid region comprises a
region within the airfoil that is solid material.
8. The rotor blade of claim 7, wherein the hollow region of the
outboard region of the airfoil comprises a hollow chamber.
9. The rotor blade of claim 8, wherein the hollow chamber comprises
a connector extending therethrough that structurally connects the
pressure sidewall to the suction sidewall of the airfoil.
10. The rotor blade of claim 9, wherein the connector comprises a
plurality of ribs, a plurality of which extends across the hollow
chamber from the suction sidewall to the pressure sidewall of the
airfoil.
11. The rotor blade of claim 9, wherein the connector comprises a
plurality of pins, a plurality of which extends across the hollow
chamber from the suction sidewall to the pressure sidewall of the
airfoil.
12. The rotor blade of claim 9, wherein the connector comprises an
internal wall that divides the hollow chamber into a plurality of
hollow chambers.
13. The rotor blade of claim 7, wherein the outboard region of the
airfoil comprises a hollowness percentage that defines what portion
of a volume of the outboard region is the hollow region; and
wherein the inboard region of the airfoil comprises a solidness
percentage that defines what portion of a volume of the inboard
region is the solid region.
14. The rotor blade of claim 13, wherein the outboard region being
substantially hollow is defined as the outboard region having a
hollowness percentage of at least 70%; and wherein the inboard
region being substantially solid is defined as the inboard region
having the solidness percentage of at least 90%.
15. The rotor blade of claim 13, wherein the outboard region being
substantially hollow is defined as the outboard region having a
hollowness percentage of at least 80%; and wherein the inboard
region being substantially solid is defined as the inboard region
having the solidness percentage of at least 95%.
16. The rotor blade of claim 13, wherein the outboard region being
substantially hollow is defined as the outboard region having a
hollowness percentage of at least 90%; and wherein the inboard
region being substantially solid is defined as the inboard region
having the solidness percentage of 100%.
17. The rotor blade of claim 7, wherein the solid region of the
outboard region is limited to: a) a thin outer wall that along an
inner surface defines a hollow chamber in the airfoil and along an
outer surface defines the suction sidewall and pressure sidewall of
the airfoil; and b) pins that span the hollow chamber structurally
connecting the pressure sidewall to the suction sidewall; and
wherein the hollow region of the inboard region is limited to
narrow interior cooling passages configured to transport coolant
across the inboard region from a coolant source formed through the
root of the rotor blade to the hollow chamber of the outboard
region.
18. The rotor blade of claim 7, wherein the solid region of the
outboard region is limited to a thin outer wall that along an inner
surface defines a hollow chamber in the airfoil and along an outer
surface defines the suction sidewall and pressure sidewall of the
airfoil; and wherein the inboard region comprises no hollow
region.
19. The rotor blade of claim 18, wherein the outboard tip of the
airfoil comprises a tip plate that encloses the hollow chamber of
the airfoil.
20. The rotor blade of claim 19, wherein the tip plate includes one
or more film cooling apertures that, during operation, are
configured to meter a release of the coolant flowing through the
hollow chamber of the airfoil.
21. The rotor blade of claim 18, wherein the outboard tip of the
airfoil comprises an open face opening to the hollow chamber of the
airfoil.
22. A combustion turbine engine that includes: a 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: a mid-span shroud
configured to engage a corresponding mid-span shroud on at least
one neighboring rotor blade during operation; wherein, outboard of
the mid-span shroud, the airfoil includes an outboard region that
is substantially hollow, and, inboard of the mid-span shroud, the
airfoil includes an inboard region that is substantially solid;
wherein a hollow region comprises a region within the airfoil that
is hollow, and wherein a solid region comprises a region within the
airfoil that is solid material; wherein the outboard region of the
airfoil comprises a hollowness percentage that defines what portion
of a volume of the outboard region is the hollow region; wherein
the inboard region of the airfoil comprises a solidness percentage
that defines what portion of a volume of the inboard region is the
solid region; and wherein the outboard region being substantially
hollow is defined as the outboard region having a hollowness
percentage of at least 70% and wherein the inboard region being
substantially solid is defined as the inboard region having the
solidness percentage of at least 90%.
Description
BACKGROUND OF THE INVENTION
[0001] The present application relates generally to apparatus,
methods and/or systems concerning the design and operation of
turbine rotor blades. More specifically, but not by way of
limitation, the present application relates to apparatus and
systems pertaining to turbine rotor blades and configurations of
turbine rotor blades having mid-span 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 turbine rotor
blades extend radially outwardly from a supporting rotor disc. Each
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 generally concave pressure
side and generally 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 shroud 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 damping and
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 performance benefit. However,
taking full advantage of the damping function 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 they
seal the gas path more effectively and may be designed to provide
more significant connection between neighboring rotor blades, which
may improves damping, one of ordinary skill in the art will
appreciate that larger tip shrouds are troublesome because of the
increased pull load on the rotor 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
further 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 "mid-span 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,
which is the portion of the airfoil that extends outboard of the
mid-span shroud. This cantilevered portion of the airfoil typically
results in lower frequency vibration and increased vibratory loads,
which may be damaging to the engine. 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 a root and an outboard tip.
The rotor blade may further include a mid-span shroud configured to
engage a corresponding mid-span on at least one neighboring rotor
blades during operation. Outboard of the mid-span shroud, the
airfoil includes an outboard region that is substantially hollow.
Inboard of the mid-span shroud, the airfoil includes an inboard
region that is substantially solid.
[0007] 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
[0008] 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:
[0009] FIG. 1 is a schematic representation of an exemplary
combustion turbine engine in which embodiments of the present
application may be used;
[0010] FIG. 2 is a sectional view of the compressor in the
combustion turbine engine of FIG. 1;
[0011] FIG. 3 is a sectional view of the turbine in the combustion
turbine engine of FIG. 1;
[0012] FIG. 4 is a perspective view of an exemplary turbine rotor
blade having a tip shroud of conventional design;
[0013] FIG. 5 is a perspective view of an exemplary turbine rotor
blade having a mid-span shroud that may be used with embodiments of
the present invention;
[0014] FIG. 6 is a perspective view of turbine rotor blades having
mid-span shrouds as in FIG. 5 in an installed condition;
[0015] FIG. 7 is a top view of turbine rotor blades having mid-span
shrouds as in FIG. 5 in an installed condition;
[0016] FIG. 8 is a side view of a turbine rotor blade having a
mid-span shroud and internal configuration according to an
embodiment of the present invention;
[0017] FIG. 9 is a side view of a turbine rotor blade having a
mid-span shroud and internal configuration according to an
alternative embodiment of the present invention;
[0018] FIG. 10 is a top cross-sectional view of the outboard region
of an exemplary turbine rotor blade in accordance with an
embodiment of the present invention;
[0019] FIG. 11 is a top cross-sectional view of the outboard region
of an exemplary turbine rotor blade in accordance with an
alternative embodiment of the present invention;
[0020] FIG. 12 is a top cross-sectional view of the outboard region
of an exemplary turbine rotor blade in accordance with an
alternative embodiment of the present invention; and
[0021] FIG. 13 is a top cross-sectional view of the inboard region
of an exemplary turbine rotor blade in accordance with an
embodiment of the present invention.
DETAILED DESCRIPTION OF THE INVENTION
[0022] As an initial matter, it will be appreciated that to discuss
the invention of the present application, it may be necessary to
select terminology to refer to and describe particular components
within a combustion turbine engine. Whenever possible, common
industry terminology will be used and employed in a manner
consistent with its accepted meaning. However, it is meant that any
such terminology be given a broad meaning and not narrowly
construed such that the meaning intended herein and the scope of
the appended claims is unreasonably restricted. Those of ordinary
skill in the art will appreciate that often a particular component
may be referred to using several different terms. In addition, what
may be described herein as being single part may include and 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 part. As such, in
understanding the scope of the present invention, attention should
not only be paid to the terminology and description provided
herein, but also to the structure, configuration, function, and/or
usage of the component, particularly as provided in the appended
claims.
[0023] In addition, several descriptive terms may be used regularly
herein, and it should prove 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 a
direction relative to the flow of a fluid, such as the working
fluid through the turbine engine or, for example, the flow of air
through the combustor or coolant through one of the turbine's
component systems. As such, the term "downstream" corresponds to
the direction of flow of the fluid, and the term "upstream" refers
to the direction opposite to the flow. The terms "forward" and
"aft", without any further specificity, refer to directions, with
"forward" referring to the forward or compressor end of the engine,
and "aft" referring to the aft or turbine end of the engine. The
term "radial" refers to movement or position perpendicular to an
axis. It is often required to describe parts that are at differing
radial positions with regard to a center axis. In cases such as
this, if a first component resides closer to the 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. The term "axial" refers to movement or position parallel
to an axis. Finally, the term "circumferential" refers to movement
or position around an axis. It will be appreciated that such terms
may be applied in relation to the center axis of the turbine, or,
when referring to components within a combustor, the center axis of
the combustor.
[0024] 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.
[0025] 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.
[0026] 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.
[0027] 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.
[0028] 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.
[0029] FIG. 5 provides a perspective view of an exemplary turbine
rotor blade 16 which has a mid-span 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
mid-span 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 mid-span shroud
leaves a portion of the airfoil 25 unrestrained, i.e., the portion
of the airfoil 25 that extends outboard of the mid-span 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.
[0030] FIG. 6 is a perspective view of rotor blades 16 having
mid-span 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 mid-span 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.
[0031] FIG. 8 is a side view of a rotor blade having a mid-span
shroud 51 and internal structure or configuration (which is shown
via the dotted lines) according to an embodiment of the present
invention. FIG. 9 is a similar view illustrating an alternative
embodiment for the internal structure. As used herein and as
depicted on FIGS. 8 and 9, the airfoil 25 may be described as
having an inboard region 58, which is defined as the portion of the
airfoil 25 that is radially inside or inboard of the mid-span
shroud 51, and an outboard region 59, which is defined as the
portion of the airfoil 25 that is radially outside or outboard of
the mid-span shroud 51. According to embodiments of the present
invention, the inboard region 58 is solid or substantially solid,
and the outboard region 59 is hollow or substantially hollow. With
this in mind and as used herein, a hollow region 61 is any region
or space within the airfoil 25 that is hollow, such as a chamber or
passaged formed therein, and a solid region 62 is any region or
space within the airfoil 25 that is composed of a solid material.
According to certain embodiments of the present invention, the
outboard region 59 may include a substantial amount of hollow
region 61 formed within it. As illustrated in FIGS. 8 and 9, the
hollow region 61 of the outboard region 59 include a hollow chamber
65 that takes up much of the volume between a pressure sidewall 26
and a suction sidewall 27. In certain embodiments, such as the one
illustrated in FIG. 8, the inboard region 58 may include solid
region 62 that takes up all of the volume between the pressure
sidewall 26 and the suction sidewall 27. As illustrated in FIG. 9,
the inboard region may have solid region 62 that takes up almost
within this portion of the airfoil 25, with the exception being a
few interior cooling passages 81 configured to carry coolant to the
hollow chamber 65 of the outboard region 59.
[0032] The mid-span shroud 51, according to the present invention,
may be defined broadly any shroud that is positioned inboard of an
outboard tip 41 of the airfoil 25 and outboard of a platform 24.
According to certain embodiments of the present invention, a
mid-span shroud 51 is one positioned near the approximate radial
center of the airfoil 25.
[0033] A mid-span shroud 51 according to present invention also may
be defined as a shroud disposed within a range of radial positions
on the airfoil 25. According to certain embodiments of the present
invention, the range of positions of a mid-span shroud 51 is
defined between an inboard boundary of approximately 25% of the
radial height of the airfoil 25 and an outboard boundary of
approximately 75% of the radial height of the airfoil 25. According
to other embodiments of the present invention, as may be defined by
the appended claims, the range of positions of a mid-span shroud 51
is defined between an inboard boundary of approximately 33% of the
radial height of the airfoil 25 and an outboard boundary of
approximately 66% of the radial height of the airfoil 25.
[0034] As to other characteristics of the mid-span shroud 51 of the
present invention, the mid-span shroud 51 may described as a
circumferentially extending projection that protrudes from at least
one of the pressure sidewall 26 and the suction sidewall 27 of the
airfoil 25. As shown in FIGS. 8 and 9, the mid-span shroud 51 may
include a circumferential projection protruding from each of the
pressure sidewall 26 and the suction sidewall 27. The mid-span
shroud 51, as mentioned, may be configured to engage the mid-span
shrouds 51 of neighboring rotor blades 16. The mid-span shroud 51
of the present invention may include a pressure side contact face
55 disposed at a distal end of the circumferential projection from
the pressure sidewall 26 of the airfoil 25, and a suction side
contact face 56 at a distal end of the circumferential projection
from the suction sidewall 27. The pressure side contact face 55 may
be configured to correspond to the suction side contact face 56
such that, when the rotor blade 16 is installed between two
neighboring rotor blades having the same configuration or design,
the mid-span shroud 51 of the rotor blade 16 links or engages both
of the neighboring rotor blades 16 via contact between: a) the
pressure side contact face 55 of the rotor blade 16 and the suction
side contact face 56 of one of the neighboring rotor blades; and b)
the suction side contact face 56 of the rotor blade 16 and the
pressure side contact face 55 of the other neighboring rotor blade.
The adjacent rotor blades 16 may contact each other at a
shroud-to-shroud interface 54 that is defined between the contact
faces of each of the mid-span shrouds 51.
[0035] As mentioned, the hollow region 61 of the outboard region 59
may include a hollow chamber 65. As illustrated via the dotted
lines of FIGS. 8 and 9, the hollow chamber 65 may have a profile
similar in shape to the profile of the airfoil 25. The hollow
chamber 65 may be the void that is defined between an inner
surfaces of the pressure sidewall 26 and the suction sidewall 27.
As illustrated in FIG. 9, the hollow chamber 65 may include
structure or connectors 66 that structurally support the outboard
region 59 of the airfoil 25. The connectors 66 may connect the
pressure sidewall 26 to the suction sidewall 27 of the airfoil
25.
[0036] According to the present invention, FIGS. 10 through 12
illustrate several possible embodiments for the configuration of
the connectors 66 formed within the chamber 65 of the outboard
region 59. FIG. 10 is a top cross-sectional view of the outboard
region 59 of an exemplary rotor blade 16. As illustrated, the
hollow chamber 65 may include a plurality of ribs 62. The ribs 62
may be configured to extend across the hollow chamber 65 between
the suction sidewall 27 and the pressure sidewall 26. In an
alternative embodiment, as illustrated in FIG. 11, the connectors
66 may include a number of pins 63 that extend between the suction
sidewall 27 and the pressure sidewall 26. The pins 63 may be more
numerous and thinner than the ribs 62 of FIG. 10. In another
alternative, as illustrated in FIG. 12, the connectors 66 may
include one or more internal walls 73 that extend between the
suction sidewall 27 and the pressure sidewall 26. The internal
walls 73 may extend radially and be configured such that they
separate the hollow chamber 65 into a plurality of chambers, as
provided in the illustration.
[0037] FIG. 13 is a top cross-sectional view of the inboard region
58 of a rotor blade 16 in accordance with an embodiment of the
present invention. It will be appreciated that FIG. 13, for
example, might be a cross-sectional view of the inboard region of
FIG. 9 in which a pair of interior cooling passages 81 stretch
between a coolant source formed through the root 21 of the rotor
blade 16 and the hollow chamber 61. As illustrated, but for the
narrow cooling passages 81, the inboard region 58 is substantially
solid in construction.
[0038] The outboard region 59 of the airfoil 25 may be described as
having a "hollowness percentage" that defines the percentage or
portion of the volume of the outboard region 59 that is comprised
of hollow region 61. In a similar manner, the inboard region 58 of
the airfoil 25 may be described as having a "solidness percentage"
that defines the portion of the volume of the inboard region 58 is
comprised of solid region 62. In other embodiments of the present
invention, the hollowness percentage of the outboard region 59 is
at least 70%, and the solidness percentage of the inboard region 58
is at least 90%. In other embodiments of the present invention, the
hollowness percentage of the outboard region 59 is at least 80%,
and the solidness percentage of the inboard region 58 is at least
95%. In still other embodiments of the present invention, the
hollowness percentage of the outboard region 59 is at least 90%,
and the solidness percentage of the inboard region 58 is 100%.
[0039] In certain embodiments of the present invention, the solid
region 62 of the outboard region 59 is limited to: a) a thin outer
wall that along an inner surface defines the hollow chamber 65 in
the airfoil 25 and along an outer surface defines the suction
sidewall 27 and pressure sidewall 26 of the airfoil 25; and b)
connectors 66 that span the hollow chamber 65 structurally
connecting the pressure sidewall 26 to the suction sidewall 25. In
such cases, the hollow region 61 of the inboard region 58 may be
limited to a few interior cooling passages 81 configured to
transport coolant across the inboard region 58 from a coolant
source formed through the root 21 to the hollow chamber 65 of the
outboard region 59. In other embodiments, the solid region 62 of
the outboard region 59 is limited to a thin outer wall that along
an inner surface defines a hollow chamber 65 in the airfoil 25 and
along an outer surface defines the suction sidewall 27 and pressure
sidewall 26 of the airfoil 25. In such cases, the inboard region 58
may have no hollow region 61.
[0040] In certain embodiments, the outboard tip 41 of the airfoil
25 has a tip plate 76 that encloses the hollow chamber 65 of the
airfoil 25, as illustrated in FIG. 9. The tip plate 76 may include
film cooling apertures 82 that are configured to meter the release
of a pressurized coolant within the hollow chamber 65 of the
airfoil 25 during operation. In other embodiments, the outboard tip
41 of the airfoil 25 may include an open face 77, an example of
which is shown in FIG. 8, that opens to the hollow chamber 65 of
the airfoil 25.
[0041] It will appreciated that, pursuant to the several
embodiments discussed above, the present invention provides a
manner by which the vibratory response of turbine rotor blades 16
may be reduced so to limit the damaging mechanical loads, which may
be used, in particular, to enable the lengthening of rotor blades
so that greater engine efficiencies are achieved. That is, the
present invention teaches a method by which turbine rotor blades
may be snubbed via mid-span shrouds 51 and configured internally to
limit the vibratory response of the cantilevered portion that
extends beyond the mid-span shroud 51. The method includes
increasing the stiffness and decreasing the mass of the portion of
the airfoil 25 outboard of the mid-span shroud 51 by hollowing out
a significant portion of the region 25 and, in some embodiments,
providing connecting structure through the hollowed region, while
the region of the airfoil 25 that is inboard of the mid-span shroud
51 remains solid. In this manner, natural frequencies of the
structure may be raised and harmful vibratory responses avoided,
thereby allowing for longer turbine blades, which, in turn, may be
used to enable larger turbine engines having greater output and
efficiency.
[0042] 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.
* * * * *