US20160040537A1

US20160040537A1 - Turbine blade mid-span shroud assembly

Info

Publication number: US20160040537A1
Application number: US14/453,914
Authority: US
Inventors: David Randolph Spracher; Kevin Leon Bruce; John Wesley Harris, JR.; Brian Denver Potter; William Scott Zemitis
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2014-08-07
Filing date: 2014-08-07
Publication date: 2016-02-11
Also published as: DE102015111999A1; CN205025506U; JP2016037965A

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.

Description

FIELD OF THE INVENTION

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

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.
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.
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.
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.
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.
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

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.
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.
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.
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.
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

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:

FIG. 1 illustrates a functional diagram of an exemplary gas turbine as may incorporate at least one embodiment of the present invention;

FIG. 2 is a perspective view of an exemplary turbine blade according to at least one embodiment of the present invention;

FIG. 3 is an exploded perspective view of an exemplary turbine blade according to at least one embodiment of the present invention;

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

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

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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.

Claims

What is claimed is:

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 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

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.