Snap shroud configuration

Abstract

To provide apparatus and assemblies pertaining to turbine rotor blades having multiple snubber shrouds.SOLUTION: A blade (16) includes an airfoil (25) including: a platform (24); and a concave pressure sidewall (55) and a convex suction sidewall (56) that extend axially between corresponding leading and trailing edges and radially between a root (21) and an outboard tip (41). The blade (16)further includes: an angled outboard snubber shroud (52) positioned on the airfoil (25), the angled outboard snubber shroud (52) being provided at an angle α defined from a camber on the angled outboard snubber shroud (52) and the platform (24) of the blade (16); and an angled inboard snubber shroud (53) positioned on the airfoil (25), the angled inboard snubber shroud (53) being provided at an angle φ defined from a camber on the angled inboard snubber shroud (53) and the platform (24) of the blade (16). The angle α is set independently of the angle φ.SELECTED DRAWING: Figure 2

Description

Technical Field

[0001] This embodiment generally relates to an apparatus, method, and / or system for a turbine rotor blade. More specifically, without limitation, the present application relates to an apparatus and assembly for a turbine rotor blade having a plurality of shroud segments.

Summary of the Invention

[0002] This embodiment describes a blade for use in a turbine. The blade includes an airfoil that includes a platform and concave pressure sidewalls and convex suction sidewalls that extend axially between a corresponding leading edge and trailing edge and radially between a root and an outer tip. The blade further includes an angled outer shroud positioned on the airfoil, the angled outer shroud being provided at an angle α defined by a camber on the angled outer shroud and the platform of the blade, and an angled inner shroud positioned on the airfoil, the angled inner shroud being provided at an angle φ defined by a camber on the angled inner shroud and the platform of the blade. The angle α is set independently of the angle φ.

[0003] This application further describes a gas turbine engine having a turbine including a row of blades arranged at circumferential intervals. Each blade can include an airfoil portion including a platform and a concave pressure sidewall and a convex suction sidewall extending axially between a corresponding leading edge and trailing edge and radially between a root and an outer tip. The blade further includes an angled outer snub shroud positioned on the airfoil portion, the angled outer snub shroud provided at an angle α defined by a camber on the angled outer snub shroud and the platform of the blade, and an angled inner snub shroud positioned on the airfoil portion, the angled inner snub shroud provided at an angle φ defined by a camber on the angled inner snub shroud and the platform of the blade. The angle α is set independently of the angle φ.

[0004] These and other features of the present application will become apparent by considering the following detailed description of the preferred embodiments in conjunction with the drawings and the appended claims.

[0005] Exemplary aspects of the present disclosure are designed to solve the problems described herein and / or other problems not considered.

[0006] These and other features of the present embodiment will be more fully understood and appreciated by carefully considering the following more detailed description of the exemplary embodiments in conjunction with the following appended drawings.

Brief Description of the Drawings

[0007]

Figure 1

Figure 2

Figure 3

Best Mode for Carrying Out the Invention

[0008] It should be noted that the drawings of the present disclosure are not necessarily drawn to scale. The drawings are intended to illustrate only typical aspects of the present disclosure and should not be considered as limiting the scope of the present disclosure. In the drawings, like reference numerals represent like elements among the drawings.

[0009] The following examples of the present embodiment can be described with reference to a specific type of turbine engine. However, those skilled in the art will recognize that the present embodiment is not limited to such use and can be applicable to other types of turbine engines unless otherwise particularly limited. Furthermore, it will be recognized that specific technical terms may be used to refer to specific mechanical components within the gas turbine engine when describing the present embodiment. To the extent possible, general industrial technical terms are used and utilized in the same meaning as their accepted meaning. However, such technical terms should not be construed narrowly, as those skilled in the art will recognize that such specific mechanical components can often be referred to using different technical terms. In addition, what may be described herein as a single component may be referred to in another context as consisting of a plurality of components, or what may be described herein as including a plurality of components may be referred to elsewhere as a single component. Therefore, when understanding the scope of the present embodiment, attention should be paid not only to specific technical terms but also to the accompanying description, context, and the structure, configuration, function, and / or usage of the components, particularly what may be provided in the appended claims.

[0010] In this specification, several descriptive terms can be used in a regular manner, and it may be useful to define these terms at the beginning of this section. Therefore, these terms and their definitions are as follows, unless otherwise stated. As used herein, "downstream" and "upstream" are terms indicating directions with respect to a fluid flow, such as the flow of a working fluid through the compressor, combustor, and turbine sections of a gas turbine, or the flow of a coolant through one of the component systems of an engine. The term "downstream" corresponds to the direction of the fluid flow, and the term "upstream" refers to a direction opposite or contrary to the direction of the fluid flow. The terms "front" and "rear" refer to directions with respect to the orientation of the gas turbine, unless otherwise specified, where "front" refers to the front of the engine or the compressor end, and "rear" refers to the rear of the engine or the turbine end. Additionally, considering the configuration of a gas turbine engine around a central axis, as well as this same type of configuration in several component systems, terms for describing positions relative to the axis are likely to be used. In this context, it will be recognized that the term "radial" refers to movement or position perpendicular to the axis. In connection with this, it may be necessary to describe the relative distance from the central axis. In this case, for example, if a first component is positioned closer to the central axis than a second component, then in this specification, it is stated that the first component is "radially inward" or "inside" the second component. On the other hand, if a first component is positioned farther from the axis than a second component, then in this specification, it can be stated that the first component is "radially outward" or "outside" the second component. Additionally, it will be recognized that the term "axial" refers to movement or position parallel to the axis. And finally, the term "circumferential" refers to movement or position around the axis.

[0011] In a combustion turbine engine, compressed air from a compressor is used to combust fuel in a combustor to produce a flow of hot combustion gases, and such gases flow downstream through one or more turbines from which energy can be extracted. According to such turbines, generally, rows of blades spaced circumferentially extend radially outward from a support rotor disk. Each blade typically includes a blade mount, such as a dovetail, that enables the assembly and disassembly of the blade in a corresponding slot in the rotor disk, and an airfoil that extends radially outward from the blade mount and interacts with the flow of working fluid through the engine. The airfoil has a concave pressure side and a convex suction side that extend axially between a corresponding leading edge and trailing edge and radially between a root and a tip. The blade tip is closely spaced from a radially outer stationary surface, and it will be appreciated that leakage of combustion gases flowing downstream between the turbine blades therebetween is reduced.

[0012] A shroud at the tip of the airfoil or a "tip shroud" on a downstream stage or rotor blade provides a contact point at the tip, manages bucket frequency, activates a damping source (i.e., by connecting the tips of adjacent rotor blades), and reduces leakage of the working fluid from the tip. Considering the length of the rotor blade in the downstream stage, the damping function of the tip shroud provides a significant advantage for durability. However, considering the weight added to the assembly by the tip shroud, as well as other design criteria including long-term operation exposed to high temperatures and extreme mechanical loads, it is difficult to maximize the advantages. Thus, a large tip shroud is desirable for an effective way to seal the gas path and for a strong connection that can be formed between adjacent blades, but a large tip shroud may not be ideal because the tensile load on the disk, particularly at the base of the airfoil, increases because it has to support the full load of the blade.

[0013] Another consideration is that the output and efficiency of a gas turbine engine improve as the size of the engine increases, and more specifically, as the amount of air that can pass through the engine increases. However, the size of the engine can be limited by the operable length of the turbine blades, and the longer the turbine blades, the more the flow path through the engine can be expanded. However, as the blades get longer, the mechanical load increases, which can place additional requirements on the blades and the disks that hold the blades. As the blades get longer, the natural vibration frequency of the blades during operation also decreases, and the vibration response of the blades increases. This additional vibration load places further significant requirements on the blade configuration, limits the life of the components, and in some cases, can cause vibration loads on the turbine engine. One way to address the vibration load of long blades is to use shrouds that connect adjacent blades to each other.

[0014] One way to modify the blades in consideration of the loads on the blades is to position the shroud at a lower location on the airfoil of the blade. That is, instead of adding the shroud to the tip of the blade, the shroud is positioned near the radial portion at the center of the airfoil. As used herein, such a shroud is referred to as a "snubber shroud". At this lower (or more inner) radius, the mass of the shroud reduces the stress levels on the blade. However, this type of shroud can leave a portion of the airfoil of the blade unconstrained (i.e., the portion of the airfoil that extends outside of the snubber shroud). This cantilevered portion of the airfoil can result in lower frequency vibrations and increased vibration loads. Therefore, blade configurations that reduce or limit the loads are valuable.

[0015] Figure 1 shows an exemplary combustion turbine engine in which embodiments of the present application can be used. It will be understood by those skilled in the art that the present embodiments are not limited to this type of usage. As described, the present embodiments can be used in combustion turbine engines such as engines used in power generation and aircraft, steam turbine engines, and other types of rotary engines. FIG. 1 is a schematic diagram of a combustion turbine engine 10. Generally, a combustion turbine engine operates by extracting energy from a stream of pressurized hot gas generated by fuel combustion in a stream of compressed air. As shown in FIG. 1, the combustion turbine engine 10 can be configured to have an axial compressor 11 that is mechanically coupled to a downstream turbine section, namely turbine 13, by a common shaft, and a combustor 12 positioned between the compressor 11 and the turbine 13. A diffuser 14 may be provided in the transition area between the exhaust of the turbine 13 and the exhaust system 15. The diffuser 14 guides the exhaust gas flow from the turbine 13 to the exhaust system 15 through a flow path. It is desirable to make the flow in the diffuser 14 (sometimes called the diffuser profile) free of turbulence to reduce impact (or balance the impact), vibration, and increase the efficiency of the turbine engine 10.

[0016] Figures 2 and 3 respectively show a side view and a perspective view of an exemplary turbine blade 16 including snub shrouds 52 and 53 according to the present embodiment. The snub shrouds 52 and 53 can be connected and joined to adjacent blades 16 with complementary snub structures. The connection of adjacent blades 16 may be made between interfaces 54 between the snub shrouds where the positive pressure side surface 55 of the airfoil and the negative pressure side surface 56 of the airfoil (opposite the airfoil 25 in FIG. 2) contact each other. The interfaces between the snub shrouds shown herein are schematic and not shown in a specific configuration. The illustrated interfaces between the snub shrouds are merely exemplary and are not intended to limit the embodiments in any way. Such blade connection may tend to increase the natural frequency of the assembly and damp out operating vibrations, which means that the mechanical stress that the blade 16 experiences during operation is smaller and the degradation may be slower if it is stressed and / or degraded at any time.

[0017] Figures 2 and 3 show aspects of the present embodiment. As shown in FIG. 2, the present embodiment describes a blade 16 having an airfoil portion 25 with a dual snub shroud, an outer snub shroud 52, and an inner snub shroud 53. The airfoil portion 25 extends from the root 21, and the platform 24 is a planar platform having essentially a fillet (not shown, transitioning from the base 21 to the airfoil portion 25). Each of the outer snub shroud 52 and the inner snub shroud 53 is provided on the blade 16 at angles α and φ, respectively. The respective angles α and φ of the outer snub shroud 52 and the inner snub shroud 53 are specified with respect to the platform 24 of the blade 16. There are several advantages to this snub shroud arrangement. Since a portion of its mass is repositioned near the axis of rotation, the tip mass is overall reduced, reducing the mechanical stress on the airfoil portion. Also, the angled snub shroud configuration embodied herein provides a flow in the diffuser 14 with reduced or essentially no turbulence by the outer snub shroud 52 and the inner snub shroud 53, reducing mechanical stress, shock (or balancing any shock of the turbine engine 10), and vibration (all of which can potentially reduce the operating output and efficiency of the turbine engine 10). Thus, the outer snub shroud 52 and the inner snub shroud 53 can increase the efficiency of the turbine engine 10 at angles α and φ, respectively.

[0018] In addition, the reduction of the mechanical stress of the blade and blade vibration as a result of the angled snub shroud configuration, as embodied herein, may be able to reduce the initial gap clearance between the tip 41 of the blade and its adjacent stationary structure 58 because the airfoil portion 25 should experience less elongation during operation. By reducing the tensile force of mechanical elongation on the blade 16, the life of the blade can also be extended.

[0019] The angled inner snub shroud 53 can be configured as a circumferentially extending protrusion that projects from one or both of the pressure side wall 55 and the suction side wall 56 of the airfoil portion 25. Similarly, the angled outer snub shroud 52 can be configured as a circumferentially extending protrusion that projects from one or both of the pressure side wall 55 and the suction side wall 56 of the airfoil portion 25. As described above, each of the angled snub shrouds can be configured to engage an adjacent snub shroud formed on one or both of the adjacent blades during installation. It will be appreciated that the dual contact points enabled by the present embodiment can advantageously limit the vibration response of the blade during operation.

[0020] The snub shrouds embodied by the present disclosure are provided at an angle with respect to the platform 24. For example, the angled outer snub shroud 52 is shown at an angle α in FIG. 2. The angle α is defined between the platform 24 of the blade and the longitudinal axis which is the camber line of the angled outer snub shroud 52. The angle α is provided in a fixed geometric shape. The angle α is provided in the range of about 0 degrees to about 17 degrees, such as about 5 degrees to about 15 degrees, and in some cases about 7 degrees to about 11 degrees. Preferably, the angle α is about 9 degrees in order to balance the load on the angled outer snub shroud 52 as embodied herein.

[0021] Similarly, the angled inner snub shroud 53 is at an angle φ as shown in FIG. 2. The angle φ is provided in a fixed geometric shape. The angle φ is defined between the platform 24 of the blade and the longitudinal axis which is the camber line of the angled inner snub shroud 53. The angle φ is provided in the range of about 0 degrees to about 17 degrees, such as about 3 degrees to about 14 degrees, and in some cases about 5 degrees to about 10 degrees. Preferably, the angle φ is about 7 degrees in order to balance the load on the angled inner snub shroud 53.

[0022] Accordingly, when the angle α = 9 degrees and the angle φ = 7 degrees as described above, the angle α is not equal to the angle φ, and the angle α is greater than the angle φ.

[0023] The angled outer snub shroud 52 and the angled inner snub shroud 53 can be formed integrally with the blade 16. For example, without limiting the present embodiment, each of the angled outer snub shroud 52 and the angled inner snub shroud 53 can be cast with the blade 16, and the angled outer snub shroud 52 and the angled inner snub shroud 53 are fixed at a desired angle. Of course, if necessary, the angled outer snub shroud 52 and the angled inner snub shroud 53 can be formed separately from the blade 16 and then attached to the airfoil portion 25, and, without limitation, can be formed into an integral configuration by means such as welding, brazing, or other such attachment means known currently or developed in the future.

[0024] The angles of the outer snub shroud 52 and the inner snub shroud 53 are set independently of each other. As used herein, independent and independently set (used interchangeably) means that the angled outer snub shroud 52 can have its angle α at one value and the angle φ of the angled inner snub shroud 53 can be a different value. The angles φ and α of the outer snub shroud 52 and the inner snub shroud 53 need not be the same and can in fact be different. The angles φ and α of the outer snub shroud 52 and the inner snub shroud 53 are set to balance the loads on the airfoil portion 25 and may be capable of reducing vibrations and mitigating impacts in and on the airfoil portion 25.

[0025] The independent angles φ and α of the outer snub shroud 52 and the inner snub shroud 53 can also regulate and smooth the flow around the airfoil 25 over the angled outer snub shroud 52 and the angled inner snub shroud 53. Thus, the flow downstream of the airfoil 25 with the angled outer snub shroud 52 and the angled inner snub shroud 53 as embodied herein can have an overall pressure profile at the inlet to the diffuser 14 of the turbine engine 10 in order to reduce the effects of loading, stress, vibration, and other potentially harmful operations. The resulting overall pressure profile of the airfoil 25 with the angled outer snub shroud 52 and the angled inner snub shroud 53 can balance the upper and lower impacts of the angled outer snub shroud 52 and the angled inner snub shroud 53 and the smoothing of the flow passing therethrough. The balancing of the impacts can also potentially manage a reduction in efficiency (also known as flow energy) since the partial span shroud provides a smooth flow to the diffuser 14. This smooth flow does not impede the diffuser 14 and the exhaust section 15, thus preventing the rough flow from degrading the operating energy from the turbine engine 10. The partial span shroud configuration is also set to obtain a balance between the losses of the turbine and the downstream diffuser. The circumferential overall pressure profile entering the diffuser affects the recovery rate of the turbine. The angles of the dual partial span shrouds can be used to configure the profile in addition to the blade throat setting. For example, increasing the upper angle can result in a stronger tip profile as more flow is directed thereto, and the lower partial span shroud angle can bias the bottom of the blade pressure profile. The optimal profile angle for performance balances the impact losses near the shroud and sets the desired radial pressure profile of the diffuser over a range of turbine operating conditions.

[0026] As shown in FIG. 2, the outer snub shroud 52 can be positioned near the outer tip 41 of the airfoil portion 25, and the inner snub shroud 53 can be positioned near the radially central region of the airfoil portion 25. In an alternative embodiment, the outer snub shroud 52 is positioned immediately inside the outer tip 41 of the airfoil portion 25, and the inner snub shroud 53 is positioned at approximately the midpoint of the airfoil portion 25. In another embodiment, the radial positioning of the outer and inner snub shrouds 52, 53 is defined within a range of heights defined with respect to the airfoil portion 25. In such an embodiment, the inner snub shroud 53 can be positioned within a range of radial heights defined between an inner boundary at 25% of the radial height of the airfoil portion 25 and an outer boundary at 75% of the radial height of the airfoil portion 25, and the outer snub shroud 52 can be positioned outside the inner boundary at 60% of the radial height of the airfoil portion 25. In an alternative embodiment, the inner snub shroud 53 can be positioned within a range of radial heights defined between an inner boundary at 40% of the radial height of the airfoil portion 25 and an outer boundary at 60% of the radial height of the airfoil portion 25, and the outer snub shroud 52 can be positioned within a range of radial heights defined between an inner boundary at 75% of the radial height of the airfoil portion 25 and an outer boundary at 95% of the radial height of the airfoil portion 25. In another preferred embodiment, the inner snub shroud 53 can be positioned within a range of radial heights defined between an inner boundary at 40% of the radial height of the airfoil portion 25 and an outer boundary at 60% of the radial height of the airfoil portion 25, and the outer snub shroud 52 can be positioned outside the inner boundary at 90% of the radial height of the airfoil portion 25.

[0027] According to some of the above-described embodiments, it will be appreciated that the present embodiment provides a method that can reduce the vibration response of a turbine blade, thereby limiting the vibratory mechanical loads that cause damage and also enabling an improvement in aerodynamic / leakage prevention performance. That is, according to the present embodiment, the natural frequency of the blade structure can be increased, harmful vibration responses can be avoided, thereby enabling longer turbine blades, and using these, a larger turbine engine with greater output and efficiency can be enabled. Additionally, the reduction in tip mass and resulting mechanical tension enabled by the present embodiment can allow for a smaller clearance between the blade and the surrounding stationary structure.

[0028] All elements of means-plus-function or step-plus-function in the following claims, corresponding structures, materials, acts, and equivalents thereof, are intended to include any structure, material, or act for performing the function in combination with other claimed elements specifically recited for performing that function. The description of the present disclosure has been presented for purposes of illustration and description and is not intended to be exhaustive or to limit the disclosure to the forms disclosed. Many modifications and variations will be apparent to those of ordinary skill in the art without departing from the scope and spirit of the present disclosure. The present embodiment has been chosen and described in order to best explain the principles of the disclosure and the practical application, and to enable others of ordinary skill in the art to understand the disclosure in various embodiments with various modifications as are suited to the particular use contemplated.

Description of Reference Numerals

[0029] 10 Combustion turbine engine 11 Axial flow compressor 12 Combustor 13 Turbine 14 Diffuser 15 Exhaust system, exhaust section 16 Turbine blade 21 Root, base 24 Platform 25 Airfoil 41 Outer tip 52 Angular outer snap shroud 53 Angular inner snap shroud 54 Interface 55 Positive pressure side surface, positive pressure side wall 56 Negative pressure side surface, negative pressure side wall 58 Stationary structure α Angle φ Angle

Claims

1. A blade (16), wherein the blade (16) has a platform (24), and an airfoil portion (25) including a concave pressure side wall (55) and a convex suction side wall (56) that extend between a leading edge and a trailing edge corresponding in the axial direction and also extend between a root (21) and an outer tip (41) in the radial direction, and the blade (16) has an angled outer snub shroud (52) positioned on the airfoil portion (25), the angled outer snub shroud (52) being provided at an angle α defined by a camber on the angled outer snub shroud (52) and the platform (24) of the blade (16), and an angled inner snub shroud (53) positioned on the airfoil portion (25), the angled inner snub shroud (53) being provided at an angle φ defined by a camber on the angled inner snub shroud (53) and the platform (24) of the blade (16), further including, wherein the angle α is set independently of the angle φ and the angle α is greater than the angle φ, the blade (16).

2. The blade (16) according to claim 1, wherein the angle α is provided in a range of about 0 degrees to about 17 degrees.

3. The blade (16) according to claim 1, wherein the angle α is about 9 degrees.

4. The blade (16) according to claim 1, wherein the angle φ is provided in a range of about 0 degrees to about 17 degrees.

5. The blade (16) according to claim 1, wherein the angle φ is about 7 degrees.

6. The blade (16) according to claim 1, wherein the angle α of the angled outer snub shroud (52) and the angle φ of the angled inner snub shroud (53) are fixed.

7. The angled inner snub shroud (53) includes a snub shroud disposed within a first range of a radial height defined on the airfoil portion (25), the first range including an inner boundary at 40% of the radial height of the airfoil portion (25) and an outer boundary at 60% of the radial height of the airfoil portion (25), 2. The blade of claim 1, wherein the angled outer snubber shroud comprises a snubber shroud disposed within a second range of radial height defined on the airfoil, the second range including an inner boundary at 75% of the radial height of the airfoil and an outer boundary at 95% of the radial height of the airfoil.

8. 8. A gas turbine engine (10) having a turbine (13) with a row of circumferentially spaced blades (16), each of the blades (16) being a blade (16) according to any one of claims 1 to 7.

Images