Damping coating with constrained layer

Abstract

To provide a damping coating applied to a turbine component including a damping coating comprising a viscoelastic layer (71) and a constraining layer (72).SOLUTION: A coating is applied to a base (21) material. The coating includes a viscoelastic layer (71) and a constraining layer (72), the viscoelastic layer (71) is arranged on the base (21) material, and the constraining layer (72) is arranged on the viscoelastic layer (71).SELECTED DRAWING: Figure 8

Description

[Technical Field]

[0001] The present disclosure relates generally to damping coatings applied to turbine components, such as, but not limited to, blades, with or without part-span shrouds, to provide beneficial damping effects. More particularly, the present disclosure relates generally to damping coatings applied to turbine components with a damping coating including a viscoelastic layer and a constraining layer. [Background technology]

[0002] The turbine and compressor sections in an axial turbine engine typically include a rotor assembly including a rotating disk and a plurality of blades circumferentially arranged about the disk. Each blade may include a root, an airfoil, and a platform positioned in a transition area between the root and the airfoil. The blade root is received in a complementary shaped recess in the disk. The blade platforms extend laterally outward and collectively form a flowpath for fluid passing through the rotor stage. The forward edge of each blade is commonly referred to as the leading edge, and the aft edge is referred to as the trailing edge. Forward is defined as aft, upstream, of the gas flow through the engine.

[0003] During operation, blades can be excited into vibration by several different forcing functions. For example, fluctuations in gas temperature, pressure, and / or density can excite vibration throughout the rotor assembly, particularly within the blade airfoils. Gases exiting the upstream turbine and / or compressor sections in a cyclical or "pulsating" manner can also excite undesirable vibrations.

[0004] One concern in turbine operation is the tendency of turbine blades to be subjected to vibratory stresses during operation. In many installations, turbines are operated under conditions of frequent acceleration and deceleration. During turbine acceleration or deceleration, the blades are exposed, at least momentarily, to vibratory stresses at specific frequencies, often at secondary or tertiary frequencies. When blades are exposed to vibratory stresses, the amplitude of their vibrations can easily increase to a point that could alter their operation. Summary of the Invention

[0005] A first aspect of the present disclosure provides a coating on a base material, the coating including a viscoelastic layer and a constraining layer, the viscoelastic layer being disposed on the base material and the constraining layer being disposed on the viscoelastic layer.

[0006] A second aspect of the present disclosure provides a turbine blade having a coating on a base material, the turbine blade defining the base material. The coating includes a viscoelastic layer and a constraining layer. The viscoelastic layer is disposed on the base material, and the constraining layer is disposed on the viscoelastic layer.

[0007] A third aspect of the present disclosure provides a gas turbine system, the gas turbine system including a turbine blade, the turbine blade including a base material and a coating on the base material. The coating includes a viscoelastic layer and a constraining layer. The viscoelastic layer is disposed on the base material, and the constraining layer is disposed on the viscoelastic layer, the viscoelastic layer and the constraining layer being bonded at an interface therebetween. The viscoelastic layer is partially bonded to the constraining layer. A defect is formed at the interface partially bonding the viscoelastic layer to the constraining layer, the defect enabling vibration and frictional interaction at the partially bonded interface between the partially bonded viscoelastic layer and the constraining layer.

[0008] The exemplary aspects of the present disclosure are designed to solve the problems described herein and / or other problems not discussed.

[0009] These and other features of the present disclosure will be more readily understood from the following detailed description of the various aspects of the disclosure, taken in conjunction with the accompanying drawings which illustrate various embodiments of the present disclosure. [Brief explanation of the drawings]

[0010] [Figure 1] 1 illustrates a gas turbine engine with locations where blades of the present embodiments may be used; [Figure 2] FIG. 1 illustrates an exemplary blade according to aspects of the present disclosure. [Figure 3] FIG. 1 illustrates an exemplary blade with a part span shroud, according to aspects of the present disclosure. [Figure 4] FIG. 1 illustrates an exemplary blade with a partially damping coating, according to aspects of the present disclosure. [Figure 5] FIG. 1 illustrates an exemplary blade with a part-span shroud having a partial damping coating, in accordance with aspects of the present disclosure. [Figure 6] FIG. 1 illustrates an exemplary blade with a complete blade attenuation coating, according to aspects of the present disclosure. [Figure 7] FIG. 1 illustrates an exemplary blade with a part span shroud having a complete blade damping coating, in accordance with aspects of the present disclosure. [Figure 8] FIG. 1 is a schematic diagram of an attenuating coating according to an aspect of the present disclosure. [Figure 9] FIG. 1 is a schematic diagram of a partially bonded attenuating coating according to an aspect of the present disclosure. [Figure 10] FIG. 10 is a schematic diagram of a partially bonded attenuating coating and particles within a cavity, according to an aspect of the present disclosure. DETAILED DESCRIPTION OF THE INVENTION

[0011] It should be noted that the drawings of the present disclosure are not to scale. The drawings are intended to illustrate only typical aspects of the present disclosure and therefore should not be considered limiting of the scope of the present disclosure. In the drawings, like numbers represent like elements between the drawings.

[0012] As an initial matter, a clear explanation of the state of the art necessitates the selection of specific terminology when referring to and describing relevant mechanical components within a turbine engine. Wherever possible, common industry terminology is used and utilized consistent with its accepted meaning. Unless otherwise noted, such terminology should be given a broad interpretation consistent with the context of this application and the scope of the appended claims. Those skilled in the art will recognize that in many cases, a particular component may be referred to using several different or overlapping terms. What may be described herein as a single part may include, and be referred to in other contexts as consisting of multiple components. Alternatively, what may be described herein as comprising multiple components may be referred to elsewhere as a single part.

[0013] Additionally, several descriptive terms may be used regularly herein, and it will prove useful to define these terms at the beginning of this section. These terms and their definitions are as follows, unless otherwise stated: As used herein, "downstream" and "upstream" are terms that indicate a direction relative to the flow of a working fluid through a turbine engine, or a fluid such as, for example, the flow of air through a combustor or a coolant through one of the turbine's component systems. The term "downstream" corresponds to the direction of fluid flow, and the term "upstream" refers to the opposite direction of flow. The terms "forward" and "aft," unless otherwise specified, refer to directions, with "forward" referring to the forward or compressor end of the engine and "aft" referring to the rear or turbine end of the engine.

[0014] It is often desired to describe components located at different radial positions relative to the central axis. The term "radial" refers to movement or position perpendicular to the axis. For example, if a first component is located closer to the axis than a second component, the first component may be described herein as being "radially inward" or "inward" of the second component. Conversely, if a first component is located farther from the axis than the second component, the first component may be described herein as being "radially outward" or "outward" of the second component. The term "axial" refers to movement or position parallel to the axis. Finally, the term "circumferential" refers to movement or position around the axis. It will be understood that such terms may be applied relative to the central axis of the turbine.

[0015] Additionally, as noted below, certain descriptive terms may be used herein in a conventional manner: the terms "first," "second," and "third" may be used interchangeably to distinguish one component from another, and are not intended to denote the location or importance of the individual components.

[0016] The terminology used herein is for the purpose of describing particular embodiments only and is not intended to limit the disclosure. As used herein, the singular forms "a," "an," and "the" are intended to include the plural forms unless the context clearly dictates otherwise. It will be further understood that the terms "comprise" and / or "comprising," as used herein, specify the presence of stated features, integers, steps, operations, elements, and / or components, but do not exclude the presence or addition of one or more other features, integers, steps, operations, elements, components, and / or sets thereof. "Optional" or "optionally" means that the subsequently-stated event or circumstance may or may not occur, and the description is meant to include instances in which the event occurs and instances in which it does not occur.

[0017] When an element or layer is referred to as "on," "engaged," "connected," or "coupled" to another element or layer, it can be directly on, engaged, connected, or coupled to the other element or layer, or intervening elements or layers may be present. Conversely, when an element is referred to as "directly on," "directly engaged," "directly connected," or "directly coupled" to another element or layer, there may not be intervening elements or layers. Other terms used to describe relationships between elements should be interpreted similarly (e.g., "between" versus "directly between," "adjacent to" versus "directly adjacent to," etc.). As used herein, the term "and / or" includes any and all combinations of one or more of the associated listed items.

[0018] In a combustion turbine engine, air compressed in a compressor is used to combust fuel in a combustor to generate a flow of hot combustion gases, which flow downstream through one or more turbines from which energy can be extracted. Such turbines generally include a row of circumferentially spaced blades extending radially outward from a supporting rotor disk. Each blade typically includes a blade root, e.g., a dovetail, that enables 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 corresponding leading and trailing edges and radially between the root and tip. It will be appreciated that the blade tips are closely spaced from a radially outer stationary surface to reduce leakage of combustion gases flowing downstream between the turbine blades therebetween.

[0019] Shrouds at the tips of airfoils, or "tip shrouds" on later-stage blades, provide a contact point at the tips, manage bucket frequencies, provide a damping source (i.e., by connecting the tips of adjacent rotor blades), and reduce tip leakage of working fluid. Given the length of the blades in later stages, the damping function of the tip shroud offers a significant durability benefit. However, maximizing the benefit can be difficult when considering other design criteria, including the weight the tip shroud adds to the assembly and extended operation exposed to high temperatures and extreme mechanical loads. Therefore, while large tip shrouds are desirable for their effective gas path sealing and the strong connection they can form between adjacent blades, they may not be ideal because they must support the full blade load, increasing tensile loads on the disk, especially at the base of the airfoil.

[0020] Another consideration is that the power output and efficiency of a gas turbine engine improves with engine size, or more specifically, the amount of air that can pass through the engine. However, engine size may be limited by the operational length of the turbine blades; longer turbine blades allow for a larger flow path through the engine. However, longer blades increase mechanical loads, which may place additional demands on the blades and the disks that hold them. Longer blades also reduce the blade's natural vibration frequency during operation, increasing the blade's vibratory response. This additional vibratory load places greater demands on the blade configuration, limiting component life and, in some cases, causing vibratory loads on the turbine engine. Another method, other than damping, to address the vibratory loads of long blades is the use of shrouds that interconnect adjacent blades.

[0021] One way to modify a blade to account for the loads it places on it is to position the shroud lower on the blade's airfoil, i.e., closer to the center or base of the blade. Instead of adding a shroud to the tip of the blade, the shroud can be positioned closer to the central radial portion of the airfoil. As used herein, such a shroud is referred to as a "partial span shroud." At this lower (or more inboard) radius, the mass of the shroud reduces stress levels at the blade root. This type of partial span shroud can leave a portion of the blade's airfoil unconstrained (i.e., the portion of the airfoil extending outward from the partial span shroud toward the tip). This portion of the airfoil is cantilevered, which can result in lower frequency vibrations and increased vibratory loads. Therefore, reducing or limiting the loads on the blade can extend the life of the blade and associated systems.

[0022] Referring to FIG. 1 , a turbomachine 90 in the form of a combustion turbine or gas turbine (GT) system 100 (hereinafter “GT system 100”) is shown. The GT system 100 includes a compressor 102 and a combustor 104. The combustor 104 includes a combustion section 105 and a fuel nozzle assembly 106. The GT system 100 also includes a turbine 108 and a common compressor / turbine shaft 110 (hereinafter “rotor 110”). In one embodiment, the GT system 100 is a 7HA.03 engine commercially available from General Electric Company, Boston, Massachusetts. A set of stationary vanes or nozzles 112 cooperate with a set of rotating blades 114 to form each stage of the turbine 108 and define a portion of the flow path through the turbine 108.

[0023] An aspect of this embodiment is shown in Figure 2. As shown in Figure 2, this embodiment provides a blade 114 with an airfoil 25 having a part-span shroud 53. The airfoil 25 extends from a root 21, and the platform 24 is essentially a planar platform with a fillet 26 (best shown in Figure 7) that transitions from the base 21 to the airfoil 25, including a pressure side 55 of the airfoil and a suction side 56 of the airfoil (opposite the airfoil 25 in Figure 2).

[0024] FIG. 3 illustrates further aspects of this embodiment. Similar parts are designated with like reference numerals. As shown in FIG. 3 , this embodiment describes a blade 214 with an airfoil 25 having a part-span shroud 53. The part-span shroud 53 can be linked and / or connected to adjacent blades in complementary adjacent part-span shroud structures on those adjacent blades. This part-span shroud 53 arrangement has several advantages, including an overall reduction in tip mass due to a relocation of a portion of its mass closer to the axis of rotation, reducing mechanical loads and stresses at the blade root. Additionally, the part-span shroud configuration, as embodied herein, can reduce turbulence to reduce mechanical stresses, shocks (or balancing shocks within the turbine system 100), and vibrations (all of which can reduce the operating power and efficiency of the turbine system 100). Thus, each part-span shroud 53 can enhance the efficiency of the turbine system 100.

[0025] Additionally, the reduced mechanical loads and stresses and blade vibrations resulting from the part-span shroud configuration, as embodied herein, may reduce the initial gap clearance between the blade at tip 41 and its adjacent stationary structure 58 because the airfoil 25 should experience less stretch during operation. The reduced mechanical stretch pull on the blade 214 may also increase the life of the blade.

[0026] However, as noted above, longer blades can place additional demands on the blades and the disks that support them due to increased mechanical loads. Longer blades also reduce the blade's natural vibration frequency during operation, increasing the blade's vibratory response. This additional vibratory load can place greater demands on the blade configuration, limit component life, and potentially cause vibratory loads on the turbine engine. Even when a partial span shroud configuration is provided, vibratory loads can still occur, including above the partial span shroud configuration away from the base of the airfoil.

[0027] Thus, the present disclosure relates to a laminar damping coating applied to a blade, which may or may not include a part-span shroud, to provide a beneficial damping effect. The additional mass of the laminar damping coating can suppress vibrations relative to the structure during operation of the machine, and thus the laminar damping coating on the blade can suppress vibrations and damp the blade.

[0028] Thus, the present disclosure provides a layered "sandwich" damping coating 70 (hereinafter "layered coating") for damping a turbine blade, the turbine blade forming the base material of the layered damping coating. The layered coating 70, as embodied by the present disclosure, can be applied to the entire airfoil 25 of a turbine blade 214 with a part-span shroud 53 (e.g., at least on the suction side and pressure side 55, 56, which is referred to herein as application of the layered coating 70 unless otherwise disclosed), or, as described below, can be applied to the airfoil 25 of a turbine blade 114 that does not include a part-span shroud. Furthermore, the layered coating 70, as embodied by the present disclosure, can be applied to substantially the entire airfoil 25 of a turbine blade 214 with a part-span shroud 53, or only to a portion of the turbine blade 114 that does not include a part-span shroud.

[0029] 4, a layered coating 70 is applied to a turbine blade 114 that does not include a part-span shroud, as embodied by the present disclosure. As shown, the layered coating 70 extends from the tip 41 across only the upper portion of the airfoil portion 25 of the blade 114. While the layered coating 70 is shown applied only to the airfoil portion 25 at the upper portion, aspects of the embodiment include the layered coating 70 being applied to the airfoil 25 from the tip 41 to any portion of the airfoil portion 25 and the blade 114 extending to the platform 24. Alternatively, the layered coating 70 can extend the entire length of the airfoil portion 25 of the blade 114 (see FIG. 7, described below).

[0030] Further, as shown in FIG. 5 , the layered coating 70 can be applied to the airfoil 25 of the blade 214, including the part-span shroud 53. In FIG. 5 , the layered coating 70 is applied from the tip 41 to the airfoil 25, stopping before the part-span shroud 53. However, this extent of the layered coating 70 on the airfoil 25 from the tip 41 to the part-span shroud 53 is merely exemplary and is not intended to limit the embodiment in any way. The layered coating 70 can extend from the tip 41 on the airfoil 25, stopping before the part-span shroud 53 (as shown), or conversely, the layered coating 70 can extend beyond the part-span shroud 53 on the airfoil 25 toward the platform 24 (see FIG. 8 , described below). Additionally, the layered coating 70 can be optionally applied from the tip 41 to coat the part-span shroud 53, providing a damping effect at those locations if the layered coating 70 extends beyond the part-span shroud 53.

[0031] 6 shows layered coating 70 on blade 114 extending across airfoil 25 from tip 41 toward platform 24. Additionally, layered coating 70 may extend from tip 41 onto platform 24, further enhancing the damping effect. In this aspect of the disclosure, layered coating 70 may cover fillet 26 formed at the intersection of airfoil 25 at platform 24, again providing additional damping for blade 114.

[0032] 7 illustrates a layered coating 70 on the blade 214 that extends the entire airfoil 25 from the tip 41 toward the platform 24. Additionally, the layered coating 70 may extend from the tip 41 of the airfoil 25 onto the platform 24, further enhancing the damping effect. In this aspect of the disclosure, the layered coating 70 may be applied over the part-span shroud 53 to provide enhanced damping at those locations. Additionally, the layered coating 70 may cover the fillet 26 formed at the intersection of the airfoil 25 at the platform 24, again providing further damping for the blade 214.

[0033] It will be appreciated that Figures 1-7 illustrate exemplary configurations of blades 114, 214. As embodied by the present disclosure, any blade may be provided with the layered coating 70 as described herein. For example, and not intended to be limiting, blades as embodied herein may include blades with top shrouds, blades with multiple part-span shrouds, solid blades, blades including cooling passages, blades for steam turbines, blades for gas turbines, blades for compressors, blades driven by various prime movers, and the like. is responsible The blade configuration may include any blade configuration.

[0034] 8 and 9 illustrate a layered coating 70. The layered coating 70 includes a viscoelastic layer 71 applied to the base material of the layered damping coating 70, here the airfoil portion 25 of the turbine blade 114, 214. A constraining layer 72 is applied over the viscoelastic layer 71. When applied to the blade 114, 214, the constraining layer 72 is applied coextensively and continuously over the viscoelastic layer 71. The constraining layer 72 can be fully bonded to the viscoelastic layer 71 or partially bonded to the viscoelastic layer 71, as described below. Additionally, friction at an interface 77 between the viscoelastic layer 71 and the constraining layer 72 acts to damp and dissipate kinetic energy during blade vibration, excitation, and / or loading.

[0035] The viscoelastic layer 71 can include, for example, a viscoelastic material formed entirely or partially of a viscoelastic material. The viscoelastic material can include, but is not limited to, one or more ceramics and / or high-temperature superalloys. The superalloy can be the material of the airfoil 25. Viscoelasticity is the property of a material that exhibits both viscous and elastic properties when deformed. Viscous materials resist shear flow when stressed and strain linearly with time. Elastic materials strain when stretched and instantly return to their original state after the stress is removed. Viscoelastic materials exhibit the ability to creep, recover, undergo stress relaxation, and absorb energy. Some examples of viscoelastic materials include amorphous polymers, semi-crystalline polymers, biopolymers, very high-temperature metals, and bituminous materials.

[0036] The constraining layer 72 can include a pre-sintered preform (PSP), such as a PSP containing a superalloy component. The constraining layer 72 is preferably a thin, hard metal layer, such as a pre-sintered preform, that enhances the energy distribution of the viscoelastic layer 71. The preform can be cut from the sintered plate, connected to a component (for example, but not limited to, by welding), and vacuum brazed to the component. The pre-sintered preform is a sintered powder metallurgy product containing a homogeneous mixture of a superalloy base material and a braze alloy powder. The PSP braze material can include a superalloy, such as, but not limited to, an iron-based superalloy, a nickel-based superalloy, or a cobalt-based superalloy. In some examples, the PSP braze material may include, in addition to the base metal, at least one of aluminum (Al), titanium (Ti), chromium (Cr), tungsten (W), molybdenum (Mo), rhenium (Re), tantalum (Ta), silicon (Si), boron (B), or iron (Fe). PSP components require minimal post-braze grinding or machining.

[0037] Because the PSP braze material may have mechanical and chemical properties (e.g., mechanical strength and high-temperature oxidation resistance) that make the braze alloy suitable for use in high-temperature oxidizing environments, the PSP braze material may facilitate the manufacture of articles for high-temperature mechanical systems in turbine components, which are then joined using the PSP braze material. The PSP braze material, as embodied herein, may be easier to position in blade regions and may result in a more uniform braze joint.

[0038] The constraining layer 72, as embodied by the present disclosure, acts as a retaining shell and is applied over the viscoelastic layer 71 that is applied over the base material airfoil 25 of the turbine blade 114, 214. As discussed above, the layered coating 70 can be applied to a partial surface of the base material airfoil 25 (see Figures 4 and 5) or over the entire surface of the base material airfoil 25 (see Figures 6 and 7).

[0039] A further aspect of the present disclosure is illustrated in FIG. 9. Enhanced damping can be achieved by forming a partial delamination zone 75 at the interface 77 of the viscoelastic layer 71 and the constraining layer 72. The partial delamination zone 75 can be created by inserting at least one, and preferably multiple, artificial defects 76 during application of the layers. The defects 76 allow for increased and enhanced vibration and friction interaction between the viscoelastic layers as they are partially bonded to the constraining layer, thus enhancing the dissipation of vibration / kinetic energy at the interface 77. The artificial defects 76 can be formed in any shape or configuration, and while shown as rectangular, this is merely exemplary and not limiting of the embodiment, and any polygonal configuration is within the scope of the embodiment.

[0040] In a further aspect of the embodiment, the artificial defects 76 may be cavities filled with small particles 78, as shown in Figure 10. The particles 78 may be duty The blade may include particles 78 of any shape, composition, configuration, or material. The particles 78 may include any material, including, but not limited to, the materials that form the viscoelastic layer 71, the constraining layer 72, and the base material 25. As the blade vibrates, the particles 78 collide with each other and with the walls of the artificial defect 76, thereby dissipating the kinetic energy of the blade.

[0041] For example, the viscoelastic layer 71 may be formed by a suitable coating process, including but not limited to plasma spraying, electron beam physical vapor deposition, chemical vapor deposition, or the like. is responsibleWhen the viscoelastic layer 71 is applied via a thermal spray process (e.g., any other suitable coating process), the defects 76 can be intentionally introduced and formed by the coating process. For example, when the viscoelastic layer 71 is thermally sprayed, voids as defects 76 can be inherent to a particular coating process, and these defects 76 are advantageously left for partial delamination as described herein. Alternatively, defects 76 can be added to the viscoelastic layer 71 during coating application of the viscoelastic layer 71, such that some defects 76 in the viscoelastic layer 71 are located at the interface 77, thus providing the desired partial delamination at the interface 77. As a further alternative, the defects 76 can be added via a separate thermal spray step, with the defects 76 located at the interface 77.

[0042] As used herein throughout this specification and claims, approximation language can be applied to modify any quantitative expression that can reasonably vary without resulting in a change in the basic function involved. Thus, values ​​modified by terms such as "approximately," "about," and "substantially" are not limited to the exact value specified. In at least some instances, approximation language can correspond to the precision of the instrument used to measure the value. Here, and throughout this specification and claims, range limitations are combinable and / or interchangeable, and unless the context and language dictate otherwise, such ranges are identified and include all subranges encompassed therein. "About," as applied to a particular value in a range, applies to both endpoints and can indicate + / - 10% of the stated value, unless specifically dependent on the precision of the instrument used to measure the value.

[0043] The corresponding structure, material, acts, and equivalents of all means-plus-function or step-plus-function elements in the following claims are intended to encompass any structure, material, or acts for performing that function in combination with other specifically claimed claim elements. 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 precise form disclosed. Many modifications and variations will be apparent to those skilled in the art without departing from the scope and spirit of the disclosure. The present embodiments were chosen and described in order to best explain the principles and practical application of the disclosure and to enable others skilled in the art to understand the disclosure in its various embodiments with various modifications as suited to the particular uses envisioned. [Explanation of symbols]

[0044] 7 Interface 21 Root, base 24 Platform 25 Airfoil, base material 26 Fillet 41 Tip 53 Partial Span Shroud 55 Pressure side 56 Negative pressure side 58 Stationary structure 70 layer coating 71 Viscoelastic layer 72 Restraint layer 75 Partial detachment zone 76 Artificial Defects 77 Interface 78 particles 90 Turbomachinery 100 Gas Turbine (GT) System 102 Compressor 104 Combustor 105 Combustion Zone 106 Fuel nozzle assembly 108 Turbine 110 Compressor / turbine shaft, rotor 112 Stationary vane, nozzle 114 Turbine blades, rotating blades 214 Turbine Blade

Claims

1. A coating applied to an airfoil portion (25) of a turbine blade (214), the coating comprising: a viscoelastic layer (71); A constraining layer (72) and It is equipped with the viscoelastic layer (71) is disposed on the airfoil portion (25), and the constraining layer (72) is disposed on the viscoelastic layer (71); The viscoelastic layer (71) and the constraining layer (72) are bonded at an interface (7, 77) therebetween, and the viscoelastic layer (71) is partially bonded to the constraining layer (72); a cavity (76) is formed at the interface (7, 77) where the viscoelastic layer (71) and the constraining layer (72) are partially bonded, the cavity (76) allowing vibration and frictional interaction at the partially bonded interface (7, 77) between the partially bonded viscoelastic layer (71) and the constraining layer (72); The coating, wherein the cavities (76) contain particles (78), the particles (78) being configured to vibrate and frictionally interact at the partially bonded interface (7, 77) between the partially bonded viscoelastic layer (71) and the constraining layer (72).

2. A coating as described in claim 1, wherein the viscoelastic layer (71) and the constraining layer (72) are continuous and coextensive on the airfoil portion (25).

3. A coating as described in claim 1, wherein the coating extends at least partially on the airfoil portion (25) from the tip (41) toward the base (21).

4. The coating of claim 1 , wherein the constraining layer (72) comprises a pre-sintered preform (PSP).

5. A turbine blade (214), an airfoil (25); A coating applied to the airfoil (25) according to any one of claims 1 to 4; A turbine blade (214).

6. 6. The turbine blade of claim 5, wherein the turbine blade includes at least one part-span shroud on the airfoil, and the coating extends from the tip toward at least the at least one part-span shroud.

7. The turbine blade (214) of claim 6, wherein the at least one part span shroud (53) on the airfoil (25) includes the coating.

8. A turbine blade (214) as described in claim 5, wherein the coating extends completely on the airfoil portion (25) to the base (21).

9. 9. The turbine blade (214) of claim 8, wherein a fillet (26) connects the airfoil (25) to the base (21), and the coating is applied to the fillet (26) and the base (21).

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