U.S. patent application number 14/453914 was filed with the patent office on 2016-02-11 for turbine blade mid-span shroud assembly.
The applicant listed for this patent is General Electric Company. Invention is credited to Kevin Leon Bruce, John Wesley Harris, JR., Brian Denver Potter, David Randolph Spracher, William Scott Zemitis.
Application Number | 20160040537 14/453914 |
Document ID | / |
Family ID | 55134962 |
Filed Date | 2016-02-11 |
United States Patent
Application |
20160040537 |
Kind Code |
A1 |
Spracher; David Randolph ;
et al. |
February 11, 2016 |
TURBINE BLADE MID-SPAN SHROUD ASSEMBLY
Abstract
A mid-span shroud assembly for a turbine blade includes a
pressure side shroud body defining a spar pocket and a fastener
hole and a suction side shroud body defining a spar pocket and a
fastener hole. The mid-span shroud assembly further includes a spar
having a first end portion which extends within the spar pocket of
the pressure side shroud body and a second end portion which
extends within the spar pocket of the suction side shroud body. The
spar is formed to extend through a bore hole of the turbine blade.
A fastener is formed to extend through the fastener hole of the
pressure side shroud body, a fastener orifice of the turbine blade
and the fastener hole of the suction side shroud body to provide a
clamping force to hold the pressure side and suction side shroud
bodies against the airfoil.
Inventors: |
Spracher; David Randolph;
(Simpsonville, SC) ; Bruce; Kevin Leon;
(Greenville, SC) ; Harris, JR.; John Wesley;
(Taylors, SC) ; Potter; Brian Denver; (Greer,
SC) ; Zemitis; William Scott; (Simpsonville,
SC) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company |
Schenectady |
NY |
US |
|
|
Family ID: |
55134962 |
Appl. No.: |
14/453914 |
Filed: |
August 7, 2014 |
Current U.S.
Class: |
416/196R |
Current CPC
Class: |
F01D 5/225 20130101;
Y02T 50/60 20130101; Y02T 50/673 20130101; F01D 5/147 20130101 |
International
Class: |
F01D 5/22 20060101
F01D005/22 |
Claims
1. A mid-span shroud assembly for a turbine blade, the mid-span
shroud assembly comprising: a pressure side shroud body, the
pressure side shroud body defining a spar pocket and a fastener
hole; a suction side shroud body, the suction side shroud body
defining a spar pocket and a fastener hole; a spar having a first
end portion extending within the spar pocket of the pressure side
shroud body and a second end portion extending within the spar
pocket of the suction side shroud body, wherein the spar is formed
to extend through a bore hole of the turbine blade; and a fastener
formed to extend through the fastener hole of the pressure side
shroud body, a fastener orifice of the turbine blade and the
fastener hole of the suction side shroud body, wherein the fastener
provides a clamping force to hold the pressure side shroud body
against a pressure side wall of the turbine blade and the suction
side shroud body against a suction side wall of the turbine
blade.
2. The mid-span shroud assembly as in claim 1, further comprising
an interlock feature defined within the spar pocket of the pressure
side shroud body.
3. The mid-span shroud assembly as in claim 1, further comprising
an interlock feature defined within the spar pocket of the suction
side shroud body.
4. The mid-span shroud assembly as in claim 1, wherein at least one
of the first and second end portions of the spar includes
interlocking features which extend radially outwardly with respect
to an axial centerline of the spar.
5. The mid-span shroud assembly as in claim 1, wherein the first
end portion of the spar is non-rigidly situated within the spar
pocket of the pressure side shroud body.
6. The mid-span shroud assembly as in claim 1, wherein the second
end portion of the spar is non-rigidly situated within the spar
pocket of the suction side shroud body.
7. The mid-span shroud assembly as in claim 1, wherein the pressure
side shroud body includes a mating side portion formed to contour
with the pressure side wall of the turbine blade.
8. The mid-span shroud assembly as in claim 1, wherein the suction
side shroud includes a mating side portion formed to contour with
the suction side wall of the turbine blade.
9. The mid-span shroud assembly as in claim 1, wherein at least a
portion of the spar is non-cylindrical.
10. A turbine blade, comprising: an airfoil having a pressure side
wall, a suction side wall, a bore hole and a fastener orifice, the
bore hole and the fastener orifice extending through the pressure
and suction side walls; and a mid-span shroud assembly, the
mid-span shroud assembly comprising: a pressure side shroud body
having a mating side portion formed to contour to the pressure side
wall, the pressure side shroud body defining a spar pocket and a
fastener hole; a suction side shroud body having a mating side
portion formed to contour to the suction side wall, the suction
side shroud body defining a spar pocket and a fastener hole; a spar
which extends through the bore hole, the spar including a first end
portion and a second end portion, the first end portion situated
within the spar pocket of the pressure side shroud body and the
second end portion situated within the spar pocket of the suction
side shroud body; and a fastener which extends through the fastener
hole of the pressure side shroud body, the fastener orifice and the
fastener hole of the suction side shroud body, wherein the fastener
provides a clamping force to hold the pressure side shroud body and
the suction side shroud body against the corresponding pressure
side wall and suction side wall.
11. The turbine blade as in claim 10, further comprising an
interlock feature defined within the spar pocket of the pressure
side shroud body.
12. The turbine blade as in claim 10, further comprising an
interlock feature defined within the spar pocket of the suction
side shroud body.
13. The turbine blade as in claim 10, wherein at least one of the
first and second end portions of the spar includes interlocking
features which extend radially outwardly with respect to an axial
centerline of the spar.
14. The turbine blade as in claim 10, wherein the first end portion
of the spar is non-rigidly situated within the spar pocket of the
pressure side shroud body.
15. The turbine blade as in claim 10, wherein the second end
portion of the spar is non-rigidly situated within the spar pocket
of the suction side shroud body.
16. The turbine blade as in claim 10, wherein the spar includes an
intermediate portion which extends between the first and second end
portions and extends through the bore hole, wherein the
intermediate portion of the spar and the bore hole are
non-cylindrical.
17. A gas turbine, comprising: a compressor section; a combustor
section; and a turbine section, the turbine section including a
plurality of turbine blades coupled to a rotor shaft, each turbine
blade including an airfoil having a pressure side wall, a suction
side wall, a bore hole and a fastener orifice, the bore hole and
the fastener orifice extending through the pressure and suction
side walls, each turbine blade including a mid-span shroud assembly
coupled to the turbine blade, the mid-span shroud assembly
comprising: a pressure side shroud body having a mating side
portion formed to contour to the pressure side wall, the pressure
side shroud body defining a spar pocket and a fastener hole; a
suction side shroud body having a mating side portion formed to
contour to the suction side wall, the suction side shroud body
defining a spar pocket and a fastener hole; a spar which extends
through the bore hole, the spar including a first end portion and a
second end portion, the first end portion extending within the spar
pocket of the pressure side shroud body and the second end portion
extending within the spar pocket of the suction side shroud body;
and a fastener which extends through the fastener hole of the
pressure side shroud body, the fastener orifice and the fastener
hole of the suction side shroud body, wherein the fastener provides
a clamping force to hold the pressure side shroud body and the
suction side shroud body against the corresponding pressure side
wall and suction side wall.
18. The gas turbine as in claim 17, wherein the pressure side
shroud body includes an interlock feature defined within the
pressure side shroud body spar pocket and the suction side shroud
body includes an interlock feature defined within the suction side
shroud body spar pocket.
19. The gas turbine as in claim 17, wherein at least one of the
first and second end portions of the spar includes interlocking
members which extend radially outwardly with respect to an axial
centerline of the spar.
20. The gas turbine as in claim 17, wherein the first end portion
of the spar is non-rigidly situated within the spar pocket of the
pressure side shroud body, and wherein the second end portion of
the spar is non-rigidly situated within the spar pocket of the
suction side shroud body.
Description
FIELD OF THE INVENTION
[0001] The present invention generally relates to a turbine blade.
More particularly, this invention involves a turbine blade mid-span
shroud assembly.
BACKGROUND OF THE INVENTION
[0002] A rotating turbine blade, also known as a turbine bucket or
turbine rotor blade, converts energy from a flowing fluid such as
hot combustion gas or steam into mechanical energy by causing a
shaft of a turbomachine to rotate. As the turbomachine transitions
through various operating modes, the turbine blades are subjected
to both mechanical and thermal stresses.
[0003] Mechanical fatigue may be caused by fluctuating forces in
combination with steady state forces. More specifically, the
turbine blades may experience fluctuating forces when they rotate
through non-uniform fluid flow downstream from stationary vanes,
also known as nozzles, positioned between adjacent rows of turbine
blades. A basic design consideration for turbomachines is to avoid
or to minimize resonance with natural frequencies of the turbine
blades and the dynamic stresses produced by forced response and/or
aero-elastic instability.
[0004] For example, each turbine blade on a rotating turbine disc
experiences a dynamic force when rotated through the non-uniform
flow from stationary vanes. As the turbine blades rotate through
areas of non-uniform flow, they may exhibit a dynamic response,
such as, for example, stress, displacements, etc. Additionally, a
turbine bladed disc may be induced into a state of vibration
wherein the energy build up is a maximum. This is exemplified by
areas of the blade or disc where the stress or displacement is at a
maximum level, and the resistance to the exciting force of the
blade or disc is at a minimum. Such a condition is known as a state
of resonance.
[0005] When analysis or empirical testing indicates that a turbine
blade and/or rotor disk may encounter a resonance condition during
operation of the turbomachine, steps may be taken to facilitate
minimizing the probability of encountering resonance. For example,
shroud sets may be formed along the span of each of the turbine
blades. Each shroud set generally includes a pair of
circumferentially extending shrouds, one shroud projecting from a
suction side surface of a turbine blade and one shroud projecting
from a pressure side surface of the same turbine blade. Because the
shrouds are located intermediate to a blade root portion and a
blade tip portion of each turbine blade, they are often referred to
as mid-span shrouds. However, mid-span shrouds can be located
anywhere along the turbine blade span, not just at the physical
mid-point of the span.
[0006] Mid-span shrouds are generally effective for avoiding or
minimizing resonance with natural frequencies of the turbine blades
and/or the dynamic stresses produced by fluctuating forces or
"flutter". However, mid-span shrouds are typically cast as part of
the turbine blade and may require additional machining or other
finishing processes to produce a finished turbine blade. This may
only be cost-effective during a design phase of the turbine blade.
In addition, a cast in mid-span shroud may not be retrofitted to
pre-existing turbine blade designs.
[0007] Another method for providing mid-span shrouds to the turbine
blade includes press fitting a support member through a bore hole
defined in the turbine blade and connecting each shroud to the
support member. However, this method may result in undesirable
stresses on the turbine blade and/or may result in the support
member becoming loose within the bore hole due to differences in
thermal expansion between the turbine blade and the press-fit
support member during operation of the turbomachine. Therefore, a
non-cast or non-integral mid-span shroud assembly which connects to
a new or pre-existing turbine blade to alter frequency and mode
shape in order to mitigate flutter and/or modify bucket vibratory
characteristics would be useful.
BRIEF DESCRIPTION OF THE INVENTION
[0008] Aspects and advantages of the invention are set forth below
in the following description, or may be obvious from the
description, or may be learned through practice of the
invention.
[0009] One embodiment of the present invention is a mid-span shroud
assembly for a turbine blade airfoil. The mid-span shroud assembly
includes a pressure side shroud body which defines a spar pocket
and a fastener hole, and a suction side shroud body which defines a
spar pocket and a fastener hole. The mid-span shroud assembly
further includes a spar having a first end portion which extends
within the spar pocket of the pressure side shroud body and a
second end portion which extends within the spar pocket of the
suction side shroud body. The spar is formed to extend through a
bore hole of the turbine blade. A fastener is formed to extend
through the fastener hole of the pressure side shroud body, a
fastener orifice of the turbine blade and the fastener hole of the
suction side shroud body to provide a clamping force which holds
the pressure side shroud body to the pressure side wall of the
airfoil and the suction side shroud body against the suction side
wall of the airfoil.
[0010] Another embodiment of the present invention is a turbine
blade. The turbine blade includes an airfoil having a pressure side
wall, a suction side wall, a bore hole and a fastener orifice. The
bore hole and the fastener orifice each extend through the pressure
and suction side walls. The turbine blade further comprises a
mid-span shroud assembly. The mid-span shroud assembly includes a
pressure side shroud body having a mating side portion which is
formed to contour to the pressure side wall. The pressure side
shroud body defines a spar pocket and a fastener hole. The mid-span
shroud assembly also includes a suction side shroud body having a
mating side portion which is formed to contour to the suction side
wall of the airfoil. The suction side shroud body defines a spar
pocket and a fastener hole. A spar extends through the bore hole.
The spar includes a first end portion and a second end portion. The
first end portion is situated or extends within the spar pocket of
the pressure side shroud body and the second end portion is
situated or extends within the spar pocket of the suction side
shroud body. A fastener extends through the fastener hole of the
pressure side shroud body, the fastener orifice and the fastener
hole of the suction side shroud body. The fastener provides a
clamping force to hold the pressure side shroud body and the
suction side shroud body against the corresponding pressure side
wall and suction side wall.
[0011] Another embodiment of the present invention is a gas
turbine. The gas turbine includes a compressor, a combustion
section and a turbine section. The turbine section includes a
plurality of turbine blades which are coupled to a rotor shaft.
Each turbine blade includes an airfoil having a pressure side wall,
a suction side wall, a bore hole and a fastener orifice where the
bore hole and the fastener orifice each extend through the pressure
and suction side walls. Each turbine blade includes a mid-span
shroud assembly which is coupled to the turbine blade. The mid-span
shroud assembly comprises a pressure side shroud body having a
mating side portion which is formed to contour to the pressure side
wall. The pressure side shroud body defines a spar pocket and a
fastener hole. The mid-span shroud assembly also includes a suction
side shroud body having a mating side portion which is formed to
contour to the suction side wall. The suction side shroud body
defines a spar pocket and a fastener hole. The mid-span shroud
assembly further includes a spar which extends through the bore
hole and includes a first end portion and a second end portion. The
first end portion extends within the spar pocket of the pressure
side shroud body and the second end portion extends within the spar
pocket of the suction side shroud body. A fastener extends through
the fastener hole of the pressure side shroud body, the fastener
orifice and the fastener hole of the suction side shroud body. The
fastener provides a clamping force which holds the pressure side
shroud body and the suction side shroud body against the
corresponding pressure side wall and suction side wall.
[0012] Those of ordinary skill in the art will better appreciate
the features and aspects of such embodiments, and others, upon
review of the specification.
BRIEF DESCRIPTION OF THE DRAWINGS
[0013] A full and enabling disclosure of the present invention,
including the best mode thereof to one skilled in the art, is set
forth more particularly in the remainder of the specification,
including reference to the accompanying figures, in which:
[0014] FIG. 1 illustrates a functional diagram of an exemplary gas
turbine as may incorporate at least one embodiment of the present
invention;
[0015] FIG. 2 is a perspective view of an exemplary turbine blade
according to at least one embodiment of the present invention;
[0016] FIG. 3 is an exploded perspective view of an exemplary
turbine blade according to at least one embodiment of the present
invention;
[0017] FIG. 4 is a cross sectional top view of a portion of an
exemplary turbine blade according to one embodiment of the present
invention; and
[0018] FIG. 5 is a cross sectional top view of a portion of an
exemplary turbine blade according to one embodiment of the present
invention.
DETAILED DESCRIPTION OF THE INVENTION
[0019] Reference will now be made in detail to present embodiments
of the invention, one or more examples of which are illustrated in
the accompanying drawings. The detailed description uses numerical
and letter designations to refer to features in the drawings. Like
or similar designations in the drawings and description have been
used to refer to like or similar parts of the invention. As used
herein, the terms "first", "second", and "third" may be used
interchangeably to distinguish one component from another and are
not intended to signify location or importance of the individual
components. The terms "upstream" and "downstream" refer to the
relative direction with respect to fluid flow in a fluid pathway.
For example, "upstream" refers to the direction from which the
fluid flows, and "downstream" refers to the direction to which the
fluid flows. The term "radially" refers to the relative direction
that is substantially perpendicular to an axial centerline of a
particular component, and the term "axially" refers to the relative
direction that is substantially parallel and/or coaxially aligned
to an axial centerline of a particular component.
[0020] Each example is provided by way of explanation of the
invention, not limitation of the invention. In fact, it will be
apparent to those skilled in the art that modifications and
variations can be made in the present invention without departing
from the scope or spirit thereof. For instance, features
illustrated or described as part of one embodiment may be used on
another embodiment to yield a still further embodiment. Thus, it is
intended that the present invention covers such modifications and
variations as come within the scope of the appended claims and
their equivalents. Although an industrial or land based gas turbine
is shown and described herein, the present invention as shown and
described herein is not limited to a land based and/or industrial
gas turbine unless otherwise specified in the claims. For example,
the invention as described herein may be used in any type of
turbomachine including but not limited to a steam turbine, an
aircraft gas turbine or marine gas turbine.
[0021] Referring now to the drawings, FIG. 1 illustrates a
schematic diagram of an exemplary gas turbine 10 turbomachine as
may incorporate various embodiments of the present invention. As
illustrated, the gas turbine 10 generally includes an inlet section
12, a compressor section 14 disposed downstream of the inlet
section 12, a plurality of combustors (not shown) within a
combustor section 16 which is disposed downstream of the compressor
section 14, a turbine section 18 disposed downstream of the
combustor section 16 and an exhaust section 20 disposed downstream
of the turbine section 18. Additionally, the gas turbine 10 may
include one or more shafts 22 coupled between the compressor
section 14 and the turbine section 18.
[0022] The turbine section 18 may generally include a rotor shaft
24 having a plurality of rotor disks 26 (one of which is shown) and
a plurality of rotatable turbine blades 28 which extend radially
outwardly from and are interconnected to each rotor disk 26. Each
rotor disk 26 may, in turn, be coupled to a portion of the rotor
shaft 24 that extends through the turbine section 18. The turbine
section 18 further includes an outer casing 30 that
circumferentially surrounds the rotor shaft 24 and the turbine
blades 28, thereby at least partially defining a hot gas path 32
through the turbine section 18.
[0023] During operation, a working fluid such as air flows through
the inlet section 12 and into the compressor section 14 where the
air is progressively compressed, thus providing pressurized air to
the combustors of the combustion section 16. The pressurized air is
mixed with fuel and burned within each combustor to produce hot
gases of combustion 34. The hot gases of combustion 34 flow through
the hot gas path 32 from the combustor section 16 to the turbine
section 18, wherein energy (kinetic and/or thermal) is transferred
from the hot gases 34 to the turbine blades 28, thus causing the
rotor shaft 24 to rotate. The mechanical rotational energy may then
be used to various purposes such as to power the compressor section
14 and/or generate electricity. The hot gases of combustion 34
exiting the turbine section 18 may be exhausted from the gas
turbine 10 via the exhaust section 20.
[0024] FIG. 2 is a perspective view of an exemplary turbine blade
28 according to at least one embodiment of the present invention.
As shown in FIG. 2, the turbine blade 28 generally includes a
mounting portion 36, a platform portion 38 and an airfoil 40 that
extends substantially radially outwardly from the platform portion
38. The platform portion 38 generally serves as a radially inward
boundary for the hot gases of combustion 34 flowing through the hot
gas path 32 of the turbine section 18 (FIG. 1). As shown in FIG. 2,
the mounting portion 36 may extend substantially radially inwardly
from the platform portion 38 and may include a root structure, such
as a dovetail, formed to interconnect or secure the rotor blade 28
to the rotor disk 26 (FIG. 1). As illustrated in FIG. 2, the
airfoil 40 extends substantially radially outwardly from the
platform portion 38 in span from a root 42 of the airfoil 40 which
may be defined at an intersection between the airfoil 40 and the
platform 38, and a tip portion 44 of the airfoil 40. The tip
portion 44 is disposed radially opposite the root 42. As such, the
tip 44 may generally define the radially outermost portion of the
rotor blade 28.
[0025] FIG. 3 provides an exploded view of a portion of the turbine
blade 28 airfoil 40 according to one embodiment of the present
invention. As shown in FIGS. 2 and 3, the airfoil 40 further
includes a leading edge 46 which is oriented towards or into the
flow of hot gas 34, and a trialing edge 48 which is downstream from
the leading edge 46. As shown in FIG. 2, the leading edge 46 and
the trailing edge extend in span between the root 42 and tip
portion 44.
[0026] As shown in FIG. 3, the airfoil 40 includes a pair of
opposing side walls 50. In particular embodiments, the airfoil 40
includes a first or pressure side wall 52 and an opposing second or
suction side wall 54. The pressure side wall 52 and suction side
wall 54 extend in chord between the leading edge 46 and the
trialing edge 48 of the airfoil 40. As shown in FIG. 2, the
pressure side wall 52 and suction side wall 54 extend radially in
span between the root 42 and tip portion 44. As shown in FIG. 3,
the pressure side wall 52 generally comprises an aerodynamic,
substantially concave surface of the airfoil 40. In contrast, the
suction side wall 54 may generally define an aerodynamic,
substantially convex surface of the airfoil 40.
[0027] In particular embodiments, as shown in FIGS. 2 and 3 a
mid-span shroud assembly 100 is coupled to the airfoil 40. FIG. 3
shows the mid-span shroud assembly 100 exploded out from the
airfoil 40. The mid-span shroud assembly 100 may be located
anywhere along the airfoil 40 span and is not limited to a physical
mid-point of the span of the airfoil 40 unless otherwise provided
in the claims and/or the specification. The mid-span shroud
assembly 100 creates a contact between adjacent turbine blades 28
for a full 360 degrees around the rotor shaft 24 and/or rotor disk
26 at a desired percent of span and/or a desired percent of chord
of a given turbine blade 28. This contact alters the vibratory
characteristics (natural frequencies and mode shapes) of the
airfoil 40.
[0028] As shown in FIG. 3, the mid-span shroud assembly 100
generally includes a pair of shroud bodies 102. In one embodiment,
a first or pressure side shroud body 104 is associated with the
pressure side wall 52 of the airfoil 40 and a second or suction
side shroud body 106 is associated with the suction side wall 54 of
the airfoil 40.
[0029] As shown in FIG. 3, the pressure side shroud body 104
extends or projects outwardly from the pressure side wall 52. The
pressure side shroud body 104 extends at least partially between
the leading and trailing edges 46, 48 along the pressure side wall
52. In one embodiment, the pressure side shroud body 104 extends
along the pressure side wall 52 intermediate to the leading and
trailing edges 46, 48. In particular embodiments, the pressure side
shroud body 104 includes an inner or mating portion or surface 108
which is formed to substantially contour to a portion of the
pressure side wall 52. The inner mating portion 108 that contacts
with the airfoil 40 may have a crowned shape or distinct raised
areas in order to provide determinate contact between the airfoil
40 and the inner mating portion 108. This may be preferable when
the airfoil 40 is cast and thus not 100% repeatable from part to
part.
[0030] As shown in FIG. 3, the suction side shroud body 106 extends
or projects outwardly from the suction side wall 54. The suction
side shroud body 106 extends along the suction side wall 54 at
least partially between the leading and trailing edges 46, 48. In
one embodiment, the suction side shroud body 106 extends
substantially intermediate to the leading and trailing edges 46, 48
along the suction side wall 54. In one embodiment, as shown in FIG.
3, the suction side shroud body 106 includes an inner or mating
portion or surface 110 which is formed to substantially contour to
a portion of the suction side wall 54. The inner mating portion 110
that contacts the airfoil 40 may have a crowned shape or distinct
raised areas in order to provide determinate contact between the
airfoil 40 and the inner mating portion 110. Again, this may be
preferable when the airfoil 40 is cast and thus not 100% repeatable
from part to part.
[0031] In one embodiment, as shown in FIG. 3, the mid-span shroud
assembly 100 includes at least one spar 112 which extends through a
bore hole 56 defined by the airfoil 40. The bore hole 56 extends
through the pressure and suction side walls 52, 54 of the airfoil
40. The bore hole 56 is disposed or defined along the span of the
airfoil 40 intermediate to the root 42 and the tip portion 46. In
particular embodiments, the airfoil 40 defines a plurality of bore
holes 56 and the mid-span shroud assembly 100 includes a plurality
of spars 112 which each align with a corresponding bore hole 56. As
shown, the spar 112 may have a generally cylindrical cross
sectional shape. However, in other embodiments, the spar 112 may
have a generally non-cylindrical cross sectional shape.
[0032] In one embodiment, as shown in FIG. 3, the pressure side
shroud body 104 defines at least one fastener hole 114, the suction
side shroud body 106 defines at least one fastener hole 116 and the
airfoil 40 defines at least one fastener orifice 58. As shown, the
fastener orifice 58 aligns with the fastener hole 114 of the
pressure side shroud body 104 and with the fastener hole 116 of the
suction side shroud body 106.
[0033] In particular embodiments, as shown in FIG. 3, the mid-span
shroud assembly 100 includes at least one fastener 118 which
extends through the fastener holes 114, 116 and the fastener
orifice 58. The fastener 118 provides a clamping or inward force to
hold the pressure side shroud body 104 against the pressure side
wall 52 of the turbine blade 28 and the suction side shroud body
106 against the suction side wall 54 of the turbine blade 28. In
one embodiment, the mid-span shroud assembly 100 includes a
plurality of fastener holes 114, 116 and fastener orifices 58 and a
plurality of corresponding fasteners 118.
[0034] The fastener 118 may include any suitable fastener such as a
bolt, pin, rivet or the like. As shown in FIG. 3, the fastener 118
may include a head portion 120 which is disposed at one end of the
fastener 118. A second end of the fastener 118 may be formed with
threads and/or formed to flare outward to lock the fastener 118 in
place. In addition or in the alternative, the fastener 118 may be
welded or held in place by other suitable means such as by a nut
119 and/or by welding or the like.
[0035] FIG. 4 provides a cross sectional top view of a portion of
the airfoil 40 sectioned through the mid-span shroud assembly 100
at the spar 112, according to one embodiment of the present
invention. As shown in FIG. 4, a first spar pocket 122 is defined
by and/or within the pressure side shroud body 104 and a second
spar pocket 124 is defined by and/or within the suction side shroud
body 106. The spar pockets 122, 124 are generally aligned with the
bore hole 56 when the mid-span shroud assembly 100 is installed or
mounted on the turbine blade 28. In particular embodiments, the
pressure side and suction side shroud bodies 104, 106 may each
define a plurality of spar pockets 122, 124.
[0036] As shown in FIG. 4, the spar 112 extends through the bore
hole 56 and into each of the spar pockets 122, 124. In one
embodiment, a first end 126 of the spar 112 is non-rigidly situated
and/or extends within the spar pocket 122 of the pressure side
shroud body 104. In one embodiment, a second end 128 of the spar
112 is non-rigidly situated and/or extends within the spar pocket
124 of the suction side shroud body 106. In one embodiment, an
intermediate portion 130 of the spar 112 is non-rigidly situated or
extends within the bore hole 56 of the airfoil 40.
[0037] FIG. 5 provides a cross sectional top view of a portion of
the airfoil 40 including the mid-span shroud assembly 100,
according to one embodiment of the present invention. As shown in
FIG. 5, at least one of the spar pockets 122, 124 defines an
interlocking feature which is formed to interlock with
complementary interlocking features defined at the respective ends
of the spar 112. For example, in one embodiment the spar pocket 122
of the pressure side shroud body 104 defines interlocking feature
132 which extends at least partially around an inner surface 134 of
the spar pocket 122. The interlocking feature 132 may be formed as
a slot, groove or other surface indentation and/or as a rib, wall
or other projection which extends outward from the inner surface
134.
[0038] In addition or in the alternative, the spar pocket 124 of
the suction side shroud body 106 defines interlocking feature 136
which extends at least partially around an inner surface 138 of the
spar pocket 122. The interlocking feature 136 may be formed as a
slot, groove or other surface indentation and/or as a rib, wall or
other projection which extends outward from the inner surface
138.
[0039] As shown in FIG. 5, at least one end of the spar 112 may
include interlocking features 140 which are complementary to the
interlocking features 132, 134 of the corresponding spar pockets
122, 124. For example, the interlocking features 140 may include
spring fingers 142, 144 or other features which are formed to
interlock with the corresponding interlocking features 132, 134.
The interlocking features 140 may be used to hold the mid-span
shroud assembly 100, particularly the pressure and suction side
shroud bodies 104, 106, in place during installation and/or during
operation.
[0040] As described and illustrated herein, the present invention
provides various technical benefits over existing turbine blade
mid-span shroud technologies. For example, this mid-span shroud
assembly 100 creates a contact between adjacent turbine blades 28
for a full 360 degrees around the turbine disk 26 at a desired
percent span/percent chord of the given turbine blade 28. This
contact alters the natural frequencies and mode shapes of the
airfoil 40.
[0041] The mid-span shroud assembly 100 as provided herein is
attached using one or multiple fasteners and spars to retain the
pressure and suction side shroud bodies 104, 106 to the airfoil 40.
The fastener(s) 118 both clamp the pressure and suction side shroud
bodies 104, 106 to the airfoil 40 and to each other, while carrying
or taking the radial/shear loading of the pressure and suction side
shroud bodies 104, 106 during rotation of the turbine blades
28.
[0042] In addition, the bore hole(s) 56 and the fastener orifice(s)
58 can be positioned in relation to one another to provide a
shielding effect so as to minimize stress concentration effects
which may result from having the bore hole(s) 56 and the fastener
orifice(s) 58 within the airfoil 40. For example, stacking the bore
hole(s) 56 above the fastener orifice(s) 58 provides a better
stress state within the airfoil 40. In addition, having non-round
(ideally elliptical) shaped bore hole(s) 56 and/or fastener
orifice(s) 58 may further mitigate stress on the airfoil 40. In
addition, the spar(s) 112 may transfer the centrifugal loads of the
pressure side and suction side shroud bodies 104, 106 to the
airfoil 40, thereby reducing bending in the fastener 118. In
addition or in the alternative, the mid-span shroud assembly 100 as
presented herein may be incorporated into new OEM parts and/or may
be adapted to fit exiting turbine blade designs.
[0043] 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 and examples are intended to be within the
scope of the claims if they include 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 language of the claims.
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