Patent Description:
Turbine and compressor sections within an axial flow turbine engine generally include a rotor assembly comprising a rotating disk and a plurality of blades circumferentially disposed around the disk. Each blade can include a root, an airfoil, and a platform positioned in the transition area between the root and the airfoil. The roots of the blades are received in complementary shaped recesses within the disk. The platforms of the blades extend laterally outward and collectively form a flow path for fluid passing through the rotor stage. The forward edge of each blade is generally referred to as the leading edge and the aft edge as the trailing edge. Forward is defined as being upstream of aft in the gas flow through the engine.

During operation, blades may be excited into vibration by a number of different forcing functions. Variations in gas temperature, pressure, and/or density, for example, can excite vibrations throughout the rotor assembly, especially within the blade airfoils. Gas exiting upstream of the turbine and/or compressor sections in a periodic, or "pulsating" manner can also excite undesirable vibrations.

One concern in turbine operation is the tendency of the turbine blades to undergo vibrational stress during operation. In many installations, turbines are operated under conditions of frequent acceleration and deceleration. During acceleration or deceleration of the turbine, the blades are, momentarily at least, subjected to vibrational stresses at certain frequencies and in many cases to vibrational stresses at secondary or tertiary frequencies. When a blade is subjected to vibrational stress, its amplitude of vibration can readily build up to a point which may alter operations.

<CIT> relates to a gas turbine or turbine component partially or fully coated with a damping surface layer. The damping surface layer may have a thickness between <NUM> and <NUM> microns and may be capable of dissipating vibration or modifying a resonance frequency of the gas turbine or turbine component at ambient room temperatures including operational temperatures greater than <NUM>° F. , and the damping surface layer comprises at least one of (a) at least two layers comprising a first layer of at least one hard material and a second layer comprising at least one soft material, (b) a composite comprising a nickel alloy with a heat softenable chemistry, (c) a fine-grained nickel-based superalloy, or (d) a porous metallic coating, a porous metallic and ceramic coating, or a ceramic coating.

<CIT> relates to a rotor blade for use in a turbine of a combustion turbine engine. The rotor blade may include an airfoil that extends from a connection with a root. The rotor blade may further include a mid-span shroud configured to engage a corresponding mid-span shroud on at least one neighboring rotor blades during operation. Outboard of the mid-span shroud, the airfoil may include an outboard region that is substantially hollow, and inboard of the mid-span shroud, the airfoil may include an inboard region that is substantially solid.

<CIT> relates to a metal (e.g. titanium/titanium alloy) blade (such as a compressor or fan blade) defining a hollow interior, at least partly filled with a vibration damping and stiffening system involving varying material properties (such as elasticity, stiffness and density). The system may comprise a vibration damping layer (e.g. comprising a polymer blend) surrounding a rigid core (e.g. comprising a syntactic material). A plurality of damping layers may be provided. Details are disclosed of various materials and their respective moduli of elasticity, as well as the thicknesses of the layers involved. A method of manufacturing the blade involves the steps of forming two metal workpieces, applying stop off material to a preselected area, arranging the workpieces in a stack, heating and applying pressure to form a diffusion bond, heating and internally pressurising to form an aerofoil shape, cleaning the internal surface, bonding a vibration damping and stiffening system to the internal surface, and sealing the blade.

A first aspect of the disclosure provides a turbine blade, the turbine blade comprising: a base material; and a coating applied to the base material, wherein the coating includes a viscoelastic layer and a constraint layer. The viscoelastic layer is disposed on the base material and the constraint layer is disposed on the viscoelastic layer. The viscoelastic layer and the constraint layer join at an interface therebetween. The viscoelastic layer is partially bonded to the constraint layer. Cavities are formed at the interface partially bonding the viscoelastic layer to the constraint layer, the cavities enabling vibration and frictional interaction at the partially bonded interface between the partially bonded viscoelastic layer and the constraint layer. The cavities include particles. The particles are configured for vibration and frictional interaction at the partially bonded interface between the partially bonded viscoelastic layer and the constraint layer.

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

It is noted that the drawings of the disclosure are not to scale.

As an initial matter, in order to clearly describe the current technology it will become necessary to select certain terminology when referring to and describing relevant machine components within turbine engines. To the extent possible, common industry terminology will be used and employed in a manner consistent with its accepted meaning. Unless otherwise stated, such terminology should be given a broad interpretation consistent with the context of the present application and the scope of the appended claims. Those of ordinary skill in the art will appreciate that often a particular component may be referred to using several different or overlapping terms. What may be described herein as being a single part may include and be referenced in another context as consisting of multiple components. Alternatively, what may be described herein as including multiple components may be referred to elsewhere as a single part.

In addition, several descriptive terms may be used regularly herein, and it should prove helpful to define these terms at the onset of this section. These terms and their definitions, unless stated otherwise, are as follows. As used herein, "downstream" and "upstream" are terms that indicate a direction relative to the flow of a fluid, such as the working fluid through the turbine engine or, for example, the flow of air through the combustor or coolant through one of the turbine's component systems. The term "downstream" corresponds to the direction of flow of the fluid, and the term "upstream" refers to the direction opposite to the flow. The terms "forward" and "aft," without any further specificity, refer to directions, with "forward" referring to the front or compressor end of the engine, and "aft" referring to the rearward or turbine end of the engine.

It is often required to describe parts that are disposed at differing radial positions with regard to a center axis. The term "radial" refers to movement or position perpendicular to an axis. For example, if a first component resides closer to the axis than a second component, it will be stated herein that the first component is "radially inward" or "inboard" of the second component. If, on the other hand, the first component resides further from the axis than the second component, it may be stated herein that the first component is "radially outward" or "outboard" of the second component. The term "axial" refers to movement or position parallel to an axis. Finally, the term "circumferential" refers to movement or position around an axis. It will be appreciated that such terms may be applied in relation to the center axis of the turbine.

It will be further understood that the terms "comprises" and/or "comprising," when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components but do not preclude the presence or addition of one or more other features, integers, steps, operations, elements, components, and/or groups thereof.

Where an element or layer is referred to as being "on," "engaged to," "connected to" or "coupled to" another element or layer, it may be directly on, engaged, connected or coupled to the other element or layer, or intervening elements or layers may be present. In contrast, when an element is referred to as being "directly on," "directly engaged to," "directly connected to" or "directly coupled to" another element or layer, there may be no intervening elements or layers present.

In a combustion turbine engine, air pressurized in a compressor is used to combust a fuel in a combustor to generate a flow of hot combustion gases, whereupon such gases flow downstream through one or more turbines so that energy can be extracted therefrom. In accordance with such a turbine, generally, rows of circumferentially spaced blades extend radially outwardly from a supporting rotor disc. Each blade typically includes a blade root, e.g., a dovetail, which permits assembly and disassembly of the blade in a corresponding slot in the rotor disc, as well as an airfoil that extends radially outwardly from the blade mount and interacts with the flow of the working fluid through the engine. The airfoil has a concave pressure side and convex suction side extending axially between corresponding leading and trailing edges, and radially between a root and a tip. It will be understood that the blade tip is spaced closely to a radially outer stationary surface for reducing leakage therebetween of the combustion gases flowing downstream between the turbine blades.

Shrouds at the tip of the airfoil or "tip shrouds" on aft stage blades provide a point of contact at the tip, manage bucket frequencies, enable a damping source (i.e., by connecting the tips of neighboring rotor blades), and reduce the over-tip leakage of the working fluid. Given the length of the blades in the aft stages, the damping function of the tip shrouds provides a significant benefit to durability. However, taking full advantage of the benefits is difficult considering the weight that the tip shroud adds to the assembly and the other configuration criteria, which include enduring hours of operation exposed to high temperatures and extreme mechanical loads. Thus, while large tip shrouds are desirable because of the effective manner in which they seal the gas path and robust connection they may form between neighboring blades, large tip shrouds may not be ideal because of the increased pull loads on the disc, particularly at the base of the airfoil because it must support the entire load of blade.

Another consideration is that the output and efficiency of gas turbine engines improve as the size of the engine and, and more specifically, the amount of air able to pass through it increase. The size of the engine, however, may be limited by the operable length of the turbine blades, with longer turbine blades enabling enlargement of the flow path through the engine. Longer blades, though, incur increased mechanical loads, which may place further demands on the blades and the disc that holds them. Longer blades also decrease natural vibrational frequencies of blades during operation, which increases the vibratory response of the blades. This additional vibratory load places even greater demands on blade configuration, which may limit component life and, in some cases, may cause vibratory loads in the turbine engine. Another way, other than damping, to address the vibratory load of longer blades is through the use of shrouds that connect adjacent blades to each other.

One way to modify a blade in light of loads thereon is to position a shroud lower on the airfoil of the blade, i.e., at a position that is closer to the middle or base of the blade. Instead of adding the shroud to the tip of the blade, the shroud may be positioned near the middle radial portion of the airfoil. As used herein, such a shroud will be referred to as a "part-span shroud. " At this lower (or more inboard) radius, the mass of the shroud causes a reduced level of stress to the blade root. This type of part-span shroud may leave a portion of the airfoil of the blade unrestrained (i.e., that portion of the airfoil that extends outboard of the part-span shroud towards the tip). This portion of the airfoil is cantilevered and can result in lower frequency vibration and increased vibratory loads. Accordingly, reducing or limiting loads in a blade may increase life of a blade and associated systems.

With reference to <FIG>, a turbomachine <NUM> in the form of a combustion turbine or gas turbine (GT) system <NUM> (hereinafter `GT system <NUM>') is illustrated. GT system <NUM> includes a compressor <NUM> and a combustor <NUM>. Combustor <NUM> includes a combustion region <NUM> and a fuel nozzle assembly <NUM>. GT system <NUM> also includes a turbine <NUM> and a common compressor/turbine shaft <NUM> (hereinafter referred to as `rotor <NUM>'). In one embodiment, GT system <NUM> is a 7HA. <NUM> engine, commercially available from General Electric Company, Boston, MA. A set of stationary vanes or nozzles <NUM> cooperate with a set of rotating blades <NUM> to form each stage of turbine <NUM>, and to define a portion of a flow path through turbine <NUM>.

<FIG> illustrates aspects of the present embodiments. As illustrated in <FIG>, the present embodiments provide a blade <NUM> with an airfoil <NUM> having a part-span shroud <NUM>. The airfoil <NUM> extends from a root <NUM>, where the platform <NUM> is essentially a planar platform with a fillet <NUM> (best illustrated in <FIG>) transitioning from the base <NUM> to the airfoil <NUM>, and includes a pressure side surface <NUM> of the airfoil and a suction side surface face <NUM> of the airfoil (opposite side of the airfoil <NUM> in <FIG>).

<FIG> illustrates further aspects of the present embodiments. Like parts are indicated with similar reference numbers. As illustrated in <FIG>, the present embodiments describe a blade <NUM> with an airfoil <NUM> having a part-span shroud <NUM>. The part-span shroud <NUM> are capable of linking and/or connecting to adjacent blades at complementary adjacent part-span shroud structures on those adjacent blades. Benefits of this part-span shroud <NUM> arrangement are several, including an overall reduced tip mass as some of that mass is relocated closer to the axis of rotation, which reduces mechanical loads and mechanical stress of the blade root. Also, the part-span shroud configuration, as embodied herein, can provide a reduced turbulent flow to reduce mechanical stress, shock (or balance any shock in the turbine system <NUM>), vibrations, all of which can decrease operational output and efficiency of the turbine system <NUM>. Therefore, each part span shroud <NUM> respectively, can enhance efficiency of the turbine system <NUM>.

Additionally, the reduction in mechanical loads and stress, and blade vibration as a result of part span shroud configuration, as embodied herein, can reduce initial gap clearance between the blade at the tip <NUM> and its adjacent stationary structure <NUM> because the airfoil <NUM> should experience less elongation during operation. The reduced mechanical elongation pull on the blade <NUM> may also prolong blade life.

As noted above, longer blades, though, incur increased mechanical loads, which may place further demands on the blades and the disc that holds them. Longer blades also decrease natural vibrational frequencies of blades during operation, which increases the vibratory response of the blades. This additional vibratory load places even greater demands on blade configuration, which may limit component life and, in some cases, may cause vibratory loads in the turbine engine. Even if provided with part-span shroud configurations, vibratory loads may still be encountered, including above the part-span shroud configurations away from the base of the airfoil.

Accordingly, the present disclosure is directed to a layered damping coating applied to a blade, where the blade either includes or does not include part-span shrouds, to provide beneficial damping effects. Additional mass of the layered damping coating may inhibit vibrations on structures during operation of a machine, and thus a layered damping coating on a blade may inhibit vibrations, and damp the blade.

Therefore, the disclosure provides a layered "sandwich structure" damping coating <NUM> (hereinafter "layered coating") that damps a turbine blade, where the turbine blade forms a base material of the layered damping coating. Layered coating <NUM>, as embodied by the disclosure, can be applied to the entirely of airfoil <NUM> (e.g., at least on suction and pressure sides <NUM>, <NUM>, which will be the application of layered coating <NUM> herein unless otherwise disclosed) of turbine blade <NUM> with part-span shrouds <NUM>, or applied to an airfoil <NUM> of turbine blade <NUM> that does not include part-span shrouds, as described hereinafter. Moreover, the layered coating <NUM>, as embodied by the disclosure, can be applied to substantially an entire airfoil <NUM> of a turbine blade <NUM> with part-span shrouds <NUM> or applied only to portions of a turbine blade <NUM> that does not include part-span shrouds.

With reference to <FIG>, a layered coating <NUM>, as embodied by the disclosure, is applied to a turbine blade <NUM> that does not include part-span shrouds. As illustrated, the layered coating <NUM> extends from tip <NUM> over only an upper portion of airfoil <NUM> of blade <NUM>. While layered coating <NUM> is illustrated as only applied to airfoil <NUM> at an upper portion, aspects of the embodiment include the layered coating <NUM> applied to airfoil <NUM> from tip <NUM> to any part of airfoil <NUM> and blade <NUM> extending to the platform <NUM>. Alternatively, layered coating <NUM> can extend the entire length of airfoil <NUM> of blade <NUM> (see <FIG> described below).

Moreover, as illustrated in <FIG>, layered coating <NUM> can be applied to airfoil <NUM> of blade <NUM> that includes part span shrouds <NUM>. In <FIG>, layered coating <NUM> is applied from tip <NUM> to airfoil <NUM> stopping before part span shrouds <NUM>. However, this extent of layered coating <NUM> on airfoil <NUM> from tip <NUM> to the part span shrouds <NUM> is merely illustrative and not intended to limit the embodiments in any manner. Layered coating <NUM> can extend from tip <NUM> on the airfoil <NUM> stopping before the part span shrouds <NUM> (as illustrated) or conversely layered coating <NUM> can extend past the part span shrouds <NUM> on the airfoil <NUM> towards the platform <NUM> (see <FIG> described below). Additionally, layered coating <NUM> can be optionally applied from tip <NUM> and coated on part span shrouds <NUM> to provide damping effect at those locations, if layered coating <NUM> extends past the part span shrouds <NUM>.

<FIG> illustrates layered coating <NUM> on a blade <NUM> extending the entire airfoil <NUM> from tip <NUM> towards the platform <NUM>. Moreover, layered coating <NUM> can extend from tip <NUM> onto platform <NUM> for further enhanced damping effect. In this aspect of the disclosure, layered coating <NUM> can overlie fillet <NUM> that is formed at an intersection of airfoil <NUM> at platform <NUM>, again affording further damping of blade <NUM>.

<FIG> illustrates layered coating <NUM> on a blade <NUM> extending the entire airfoil <NUM> from tip <NUM> towards platform <NUM>. Moreover, layered coating <NUM> can extend from tip <NUM> of airfoil <NUM> onto platform <NUM> for further enhanced damping effect. In this aspect of the disclosure, layered coating <NUM> can overlie, and be applied to part span shrouds <NUM> to provide enhanced damping effect at those locations. Further, layered coating <NUM> can overlie fillet <NUM> that is formed at an intersection of airfoil <NUM> at platform <NUM>, again affording further damping of blade <NUM>.

It will be understood that <FIG> show illustrative configurations of blade <NUM>, <NUM>. As embodied by the disclosure, any blade can be provided with layered coating <NUM>, as described herein. For example, and not intended to limit the embodiments, a blade as embodied herein can include a blade with a top shroud, a blade with multiple part-span shrouds, blades that are solid, blades that include cooling passages, blades for steam turbines, blades for gas turbines, blades for compressors, blades that are driven by various motive forces, or any blade configuration, now known or hereinafter developed.

<FIG> illustrate layered coating <NUM>. Layered coating <NUM> includes a viscoelastic layer <NUM> applied to base material of layered damping coating <NUM>, here airfoil <NUM> of turbine blade <NUM>, <NUM>. A constraint layer <NUM> is applied onto viscoelastic layer <NUM>. Constraint layer <NUM> is coextensively and conterminously applied onto viscoelastic layer <NUM> when applied to blade <NUM>, <NUM>. Constraint layer <NUM> can be fully bonded to viscoelastic layer <NUM> or partially bonded to viscoelastic layer <NUM>, as described hereinafter. Moreover, friction at interface <NUM> between viscoelastic layer <NUM> and constraint layer <NUM> acts to dampen and dissipate kinetic energy during blade vibration, excitation, and/or load.

Viscoelastic layer <NUM> can include viscoelastic material, such as being formed entirely of viscoelastic material or partially of viscoelastic material. Viscoelastic material can include but is not limited to one or more ceramic and/or high-temperature superalloy(s). The superalloy may be the material of airfoil <NUM>. Viscoelasticity is the property of materials that exhibit both viscous and elastic characteristics when undergoing deformation. Viscous materials resist shear flow and strain linearly with time when a stress is applied. Elastic materials strain when stretched and immediately return to their original state once the stress is removed. Viscoelastic materials exhibit an ability to creep, recover, undergo stress relaxation and absorb energy. Some examples of viscoelastic materials include amorphous polymers, semicrystalline polymers, biopolymers, metals at very high temperatures, and bitumen materials.

The constraint layer <NUM> can include a Pre-sintered Preforms (PSP), such as a PSP including superalloy constituents. Constraint layer <NUM> is preferably a thin hard metal layer, such as a Pre-sintered Preform, which enhances energy dispensation of the viscoelastic layer <NUM>. Preforms can be cut from a sintered plate, connected to a component (for example but not limited to by welding), and vacuum brazed thereto. Pre-sintered Preforms are a sintered powder metallurgy product including a homogeneous mixture of a superalloy base material and braze alloy powders. PSP braze materials may 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, PSP braze materials may include 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), in addition to the base metal. Minimal post-braze grinding or machining is needed with PSP components.

Because the PSP braze materials may possess mechanical and chemical properties (e.g., mechanical strength and high temperature oxidation resistance) that make braze alloys suitable for use in high temperature oxidative environments, PSP braze materials may facilitate manufacture of articles for high temperature mechanical systems in turbine components, which are then joined using the PSP braze materials. PSP braze materials, as embodied herein, may be easier to position in blade regions and result in a more uniform braze j oint.

Constraint layer <NUM>, as embodied by the disclosure, acts as a retaining shell and is applied over the viscoelastic layer <NUM> that is applied over base material, airfoil <NUM> of turbine blade <NUM>, <NUM>. As discussed above, layered coating <NUM> can be applied to a partial surface of the base material, airfoil <NUM> (<FIG>) or alternatively to the entire surface of base material, airfoil <NUM> (noting <FIG>).

A further aspect of the disclosure is illustrated in <FIG>. Enhanced damping may be enabled by forming a zone of partial debonding zone <NUM>, at interface <NUM> of the viscoelastic layer viscoelastic layer <NUM> and constraint layer <NUM>. Partial debonding zone <NUM> can be made with insertion of at least one and preferably a plurality of artificial defect <NUM> during application of the layers. Defects <NUM> enable increased and enhanced vibration and frictional interaction between the viscoelastic layer as it is partially bonded to the constraint layer, and thus enhanced dissipation of vibrational/kinetic energy at the interface <NUM>. Artificial defect <NUM> can be formed in any shape or configuration, and although illustrated as a rectangle that is merely illustrative and not limiting of the embodiments, any polygonal configuration is within the scope of the embodiments.

According to the invention, artificial defects <NUM> are cavities filled with small particles <NUM>, as illustrated in <FIG>. Particles <NUM> can include particles <NUM> of any shape, composition, configuration, or materials known now or hereinafter developed. Particles <NUM> can include any materials including but not limited to materials that form viscoelastic layer <NUM>, constraint layer <NUM>, and base material <NUM>. As blade vibrates, blade kinetic energy can be dissipated as particles <NUM> impact each other and with the artificial defect <NUM> wall as well.

For example, if viscoelastic layer <NUM> is applied via an appropriate coating process (such as but not limited to plasma spraying, electron beam physical deposition, chemical vapor deposition, or any other appropriate coating process now know or hereinafter developed) defects <NUM> can be purposely introduced and formed by the coating process. For example, if viscoelastic layer <NUM> is sprayed, voids as defects <NUM> may be inherent with particular coating controls, and these defects <NUM> would be advantageously left for the partial debonding as described herein. Alternatively, defects <NUM> may be added to viscoelastic layer <NUM> during coating application of viscoelastic layer <NUM> as some defects <NUM> in viscoelastic layer <NUM> will get disposed at interface <NUM>, and thus be able to provide the desired partial debonding at the interface <NUM>. As a further alternatively, defects <NUM> can be added via a separate spray step, with the defects <NUM> disposed at interface <NUM>.

Accordingly, a value modified by a term or terms, such as "about," "approximately" and "substantially," are not to be limited to the precise value specified. Here and throughout the specification and claims, range limitations may be combined and/or interchanged; such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. "Approximately" as applied to a particular value of a range applies to both end values, and unless otherwise dependent on the precision of the instrument measuring the value, may indicate +/- <NUM>% of the stated value(s).

Claim 1:
A turbine blade (<NUM>), the turbine blade (<NUM>) comprising:
a base (<NUM>) material; and
a coating applied to the base (<NUM>) material, wherein the coating includes:
a viscoelastic layer (<NUM>); and
a constraint layer (<NUM>);
wherein the viscoelastic layer (<NUM>) is disposed on the base (<NUM>) material and the constraint layer (<NUM>) is disposed on the viscoelastic layer (<NUM>);
wherein the viscoelastic layer (<NUM>) and the constraint layer (<NUM>) join at an interface (<NUM>, <NUM>) therebetween, the viscoelastic layer (<NUM>) is partially bonded to the constraint layer (<NUM>);
characterized in that cavities (<NUM>) are formed at the interface (<NUM>, <NUM>) partially bonding the viscoelastic layer (<NUM>) to the constraint layer (<NUM>), the cavities (<NUM>) enabling vibration and frictional interaction at the partially bonded interface (<NUM>, <NUM>) between the partially bonded viscoelastic layer (<NUM>) and the constraint layer (<NUM>); and
wherein the cavities (<NUM>) include particles (<NUM>), the particles (<NUM>) configured for vibration and frictional interaction at the partially bonded interface (<NUM>, <NUM>) between the partially bonded viscoelastic layer (<NUM>) and the constraint layer (<NUM>).