Patent Description:
Ceramic matrix composites (CMC) are often used in applications requiring high strength in elevated temperature environments, such as turbine engine components, including turbine blades.

Generally, such turbine components require attachment to adjoining metallic hardware and/or metallic surfaces, sometimes referred to as disks. Among disadvantages associated with attaching a CMC to metallic hardware is the wear of the metallic hardware by the hard, abrasive ceramic material surface. Under high contact stresses, damage to the ceramic material surface is also possible, usually due to matrix cracking and fiber breakage that lead to the formation of wear troughs.

In response, U. Publication No. <CIT> discloses a method for creating a fir tree dovetail attachment for a CMC airfoil using a secondary metallic member with multiple contact surfaces. The metallic member is intended to trap the CMC and transfer the airfoil loading into the metallic member which has features/bearing surfaces similar to a multi-tooth fir tree attachment. These surfaces are designed for load transfer and not to reduce friction or wear at the disk attachment interface. Design against wear is not discussed.

<CIT> discloses use of a vibration source with a tailored frequency to help remove debris particles that build-up on the contact surface in an effort to lessen wear rates.

<CIT> discloses incorporating a circumferential internal wear pocket and radial slots, but for purposes of reducing forced excitation due to fluid flow. <CIT> discloses a compliant sleeve for turbine blades. <CIT> discloses a compliant interface for ceramic turbine blades. <CIT>, relevant under Art <NUM>(<NUM>) EPC only, discloses a blade root shim having a contact relief region, for a CMC blade.

What is needed is an apparatus and method for reducing wear between CMC-to-metal surfaces during operation at elevated temperatures.

Improvements in manufacturing technology and materials are the keys to increased performance and reduced costs for many articles and apparatus. As an example, continuing and often interrelated improvements in processes and materials have resulted in major increases in the performance of gas turbine engines, such as the improvements of the present invention. In one embodiment, the present invention is directed to an apparatus, as defined in claim <NUM>, and method as described in claim <NUM>, for manufacturing a component made from a ceramic matrix composite (CMC), in which CMC-to-metal attachment and interface occurs at elevated temperatures. Insertion of a compliant metal layer having a low coefficient of friction between the attachment contacting surfaces of the CMC and a metal component reduces wear along the CMC-to-metal attachment.

The terms interface surface, interfacing surface and the like are intended to include contacting surfaces as well as attaching or interlocking surfaces.

The present invention is directed to an apparatus for use in a heated environment including a CMC component, as defined in claim <NUM>. A metal component has a surface interface with the CMC component. A metal layer is configured for insertion between the surface interface between the CMC component and the metal component. The surface interface of the metal layer is compliant relative to asperities of the surface interface of the CMC component. A coefficient of friction between the surface interface of the CMC component and the metal component is about <NUM> or less at an operating temperature between about <NUM> to about <NUM> and a limiting temperature of the metal component.

The present invention is also directed to a method to reduce wear and friction between CMC-to-metal attachment and interface, as defined in claim <NUM>. The method includes providing a metal layer configured for insertion between a surface interface between a CMC component and a metal component. The surface interface of the metal layer is compliant relative to asperities of the surface interface of the CMC component. A coefficient of friction between the surface interface of the CMC component and the metal component is about <NUM> or less at an operating temperature between about <NUM> to about <NUM> and a limiting temperature of the metal component. The method includes inserting the layer between the surface interface between the CMC component and the metal component. The method further includes operating the CMC component, the layer and the metal component at the operating temperature.

Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings which illustrate, by way of example, the principles of the invention, the scope of the latter being defined by the appended claims.

<FIG> depicts an exemplary gas turbine engine blade <NUM>. In this illustration, a turbine blade <NUM> is constructed of a ceramic matrix composite (CMC) material. Turbine blade <NUM> is mounted to a turbine disk <NUM> in a dovetail slot <NUM>. Turbine blade <NUM> includes an airfoil <NUM>, against which a flow of hot exhaust gas is directed, and a dovetail <NUM>, also referred as a root or splayed base, that extends from airfoil <NUM> and engages dovetail slot <NUM>.

Referring now to <FIG>, which is an example of an enlarged sectional view of a CMC blade <NUM>, such as comprised of silicon carbide reinforcement fibers in a silicon carbide matrix, and disk <NUM>, such as comprised of a metal, such as a nickel alloy, the contacting or interfacing surfaces thereon are described in greater detail. The blade <NUM> includes a plurality of plies, <NUM> and <NUM>, which have been bonded together, such as by processes well known in the art. Plies <NUM> are bonded to a root core <NUM>. The lower end of blade <NUM> is defined in part by an end surface <NUM> and a root surface <NUM>. Dovetail slot <NUM> of disk <NUM> is defined by a mating surface <NUM>. A collective pair of interface or interfacing surfaces <NUM> are formed between opposed corresponding root surfaces <NUM> and mating surfaces <NUM>. Wear between root surfaces <NUM> of CMC blade <NUM> and mating surfaces <NUM> of disk <NUM> occur as a result of abrasive contact due to asperities inherent in processing CMC blade <NUM>, in combination with radially directed sliding contact between root surfaces <NUM> of CMC blade <NUM> and mating surfaces <NUM> of disk <NUM>, due to centrifugal forces generated during the high-speed rotational movement of the gas turbine engine during its operation. Superimposed, is the micro-motion due to airfoil highfrequency vibration which results in what is commonly known as fretting wear.

Although an exemplary embodiment of CMC is comprised of silicon carbide, technical ceramics such as alumina, aluminum nitride, silicon nitride or zirconia may also be used. Other available CMCs can include, for example, C/C, C/SiC, SiC/SiC and Al<NUM>O<NUM>/Al<NUM>O<NUM>. CMC materials composed of Carbon (C), special silicon carbide (SiC), alumina (Al<NUM>O<NUM>) and mullite (Al<NUM>O<NUM>-SiO<NUM>) fibers are most commonly used for CMCs. The matrix materials are usually the same; that is C, SiC, alumina and mullite).

<FIG> illustrate respective perspective and end views of an interface or interfacing layer <NUM> for insertion between CMC blade <NUM> and disk <NUM> (<FIG>). Interface or interfacing layer <NUM> includes an inner surface <NUM> and an outer surface <NUM> and layer portions <NUM>, <NUM>, <NUM> positioned between opposite layer portions <NUM>. In one embodiment, layer portions <NUM>, <NUM>, <NUM> may have optional openings <NUM> formed therein, such as for weight savings, so long as layer portions <NUM>, <NUM>, <NUM> maintain a predetermined spacing of opposed layer portions <NUM> relative to one another. That is, opposed layer portions <NUM> are contiguously interconnected by other layer portions, however configured, to form a unitary or one-piece construction. In one embodiment, the one-piece construction is symmetrical. Stated differently, the interface layer has opposed layer portions <NUM> that are symmetric, such as about a center line of an airfoil, which airfoil having symmetric interfacing surfaces about a center axis of the airfoil, such as interfacing surfaces <NUM> (<FIG>) that contact corresponding layer portions <NUM>. Applicants have determined that a one-piece construction of interface layer <NUM>, and notably a one-piece interface layer having a symmetrical construction provides improved stiffness. Such stiffness resists small amounts of motion, sometimes referred to as micro-motion, that can occur due to vibration. Separate interface surfaces, such as associated with separate shims (i.e., one shim per interface surface <NUM> (<FIG>)) has been shown to be especially susceptible to micro-motion, which micro-motion being capable of accelerating wear between the interface surfaces <NUM>, <NUM>. In another embodiment, layer portions <NUM>, <NUM>, <NUM> may resemble bands <NUM>, e.g., layer portion <NUM> and <FIG>, collectively extending in a direction <NUM> substantially transverse to an axial direction <NUM> between opposed layer portions <NUM>. Respective inner surfaces <NUM> of layer portions <NUM> of interfacing layer <NUM> form a collective interfacing surface <NUM> with root surface <NUM> of turbine blade <NUM>. Interfacing layer <NUM> is axially slid over dovetail or root <NUM> of turbine blade <NUM> and restrained radially using either a mechanical interlock, such as tab <NUM> or a secondary bond <NUM>, such as an adhesive layer or chemical bond. In one embodiment, interfacing layer <NUM> is sufficiently radially restrained to substantially prevent relative radial movement along interfacing surface <NUM> between inner surface <NUM> of layer portions <NUM> of interfacing layer <NUM> and corresponding root surfaces <NUM> of turbine blade <NUM>. In another embodiment, interfacing layer <NUM> is sufficiently radially restrained to reduce relative radial movement along interfacing surface <NUM> between inner surface <NUM> of layer portions <NUM> of interfacing layer <NUM> and corresponding root surfaces <NUM> of turbine blade <NUM>.

In an alternate embodiment, <FIG> illustrates an end view of an interfacing layer <NUM> for insertion between CMC blade <NUM> (<FIG>) and disk <NUM>. Interfacing layer <NUM> includes an inner surface <NUM> and an outer surface <NUM> and layer portions <NUM>, <NUM>, <NUM> positioned between opposite layer portions <NUM>. In one embodiment, layer portions <NUM>, <NUM>, <NUM> may have optional openings (not shown) formed therein such as previously discussed for layer portions <NUM>, <NUM>, <NUM> of interfacing layer <NUM> (<FIG>), such as for weight savings, so long as layer portions <NUM>, <NUM>, <NUM> maintain a predetermined spacing of opposed layer portions <NUM> relative to one another. In another embodiment, layer portions <NUM>, <NUM>, <NUM> may resemble bands (not shown) such as previously discussed and shown for layer portion <NUM> of interfacing layer <NUM> (<FIG>), collectively extending in a direction substantially transverse to an axial direction between opposed layer portions <NUM>. Respective outer surfaces <NUM> of layer portions <NUM> of interfacing layer <NUM> form a collective interfacing surface <NUM> with mating surfaces <NUM> of disk <NUM>. Respective inner surfaces <NUM> of layer portions <NUM> of interfacing layer <NUM> form a collective interfacing surface <NUM> with root surfaces <NUM> of turbine blade <NUM> (<FIG>). Interfacing layer <NUM> is axially slid over root or dovetail slot <NUM> of disk <NUM> and restrained radially using either a mechanical interlock (e.g., similar to tab <NUM> in <FIG>) or a secondary bond (not shown). In one embodiment, interfacing layer <NUM> is sufficiently radially restrained to substantially prevent relative radial movement along interfacing surface <NUM> between outer surface <NUM> of layer portions <NUM> of interfacing layer <NUM> and corresponding mating surfaces <NUM> of disk <NUM>. In another embodiment, interfacing layer <NUM> is sufficiently radially restrained to reduce relative radial movement along interfacing surface <NUM> between inner surface <NUM> of layer portions <NUM> of interfacing layer <NUM> and corresponding root surfaces <NUM> of turbine blade <NUM>. Optionally, a layer <NUM> of a dry film lubricant of about <NUM> to about <NUM> may be applied along interfacing surface <NUM>. Layer <NUM> of dry film lubricant may include Tungsten disulfide, Graphite-based lubricants, Molybdenum disulfide, or any suitable combinations thereof.

In one embodiment, interfacing layer <NUM>, <NUM> is between about <NUM> and about <NUM>, between about <NUM> and about <NUM>, between about <NUM> and about <NUM>, between about <NUM> and about <NUM>, or any suitable range or sub-range thereof. In one embodiment, interfacing layer <NUM>, <NUM> is about <NUM>, about <NUM>, about <NUM>, about <NUM>, about <NUM>, about <NUM>, about <NUM>, about <NUM>, or any suitable sub-range thereof.

Interfacing layer <NUM>, <NUM> is composed of a ductile or compliant material as compared to a CMC in order to cover the asperities inherent in processed CMC components. In addition, interfacing layer <NUM>, <NUM> is thin, as discussed above, which reduces the effect associated with a reduction of load carrying capacity of the blade attachment for supporting disk lugs of interfacing layer <NUM>, <NUM> as a result of its geometry. That is, at least partly as a result of interfacing layer <NUM>, <NUM> being thin, being better matched for tribology with a metal disk, being in intimate contact with and/or attached or otherwise providing an interface between the CMC or the disk as discussed above, a reduced coefficient of friction is achieved.

Interfacing layer <NUM>, <NUM> may be composed of ductile alloys configured for use at a service temperature, or an ambient temperature of the components during operation, while providing a coefficient of friction of <NUM> or less, to prevent fretting, due to vibratory motion or as corrosion. For example, for a gas turbine engine, the service temperature is typically between about <NUM> to about <NUM> and a limiting temperature of the disk, e.g., disk-grade steel alloys, or a Nickel-based alloy having about a <NUM> limiting temperature, although use of other disk materials could be greater than <NUM>. The term limiting temperature is intended to refer to a maximum temperature at which a component may be used. Exemplary compositions of interfacing layer <NUM>, <NUM> include, but are not limited to ferrous alloys, especially high Chromium steels, Nickel-based alloys, such as Alloy <NUM>, Alloy <NUM>, Alloy <NUM>, etc., and Cobalt-based alloys. In one embodiment, depending upon the types of metal utilized, the coefficient of friction is between about <NUM> and <NUM>, between about <NUM> and <NUM>, between about <NUM> and <NUM>, between about <NUM> and <NUM>, or any suitable range or sub-range thereof. In one embodiment, the coefficient of friction is about <NUM>, about <NUM>, about <NUM>, about <NUM>, about <NUM>, about <NUM>, about <NUM>, or any suitable range or sub-range thereof.

It is to be understood that the interfacing layer <NUM>, <NUM> as described herein, may also be applied to attachment locations for ceramic composites such as shrouds or combustion liners, or any other appropriate location that would benefit from a compliant layer with the benefits described herein. Additionally, the wear surfaces can in addition to contact surfaces between different components, also include lining apertures used for structural fasteners.

It is to be understood that the interfacing layer of the present invention includes arrangements, such as a root or dovetail having a "fir tree" arrangement, or multiple interfacing surfaces between the root or dovetail surfaces and the disk slot surfaces, as is well known.

Claim 1:
An apparatus for use in a heated environment comprising:
a CMC turbine blade (<NUM>);
a metal turbine disk (<NUM>) having a surface interface with the CMC turbine blade (<NUM>); and
an interfacing layer (<NUM>,<NUM>), the interfacing layer (<NUM>, <NUM>) being metal and having an inner surface (<NUM>,<NUM>) and an outer surface (<NUM>,<NUM>), and the interfacing layer (<NUM>, <NUM>) having layer portions (<NUM>, <NUM>, <NUM>) positioned between opposite layer portions (<NUM>),
the interfacing layer (<NUM>, <NUM>) being configured for insertion between root surface (<NUM>) of the CMC turbine blade (<NUM>) and mating surface (<NUM>) of the metal turbine disk (<NUM>),
wherein respective outer surfaces (<NUM>, <NUM>) of opposite layer portions (<NUM>, <NUM>) of interfacing layer (<NUM>) form a first collective interfacing surface (<NUM>) with a mating surface (<NUM>) of metal turbine disk (<NUM>), and respective inner surfaces (<NUM>, <NUM>) of opposite layer portions (<NUM>, <NUM>) of interfacing layer (<NUM>) form a second collective interfacing surface (<NUM>) with root surfaces (<NUM>) of the CMC turbine blade (<NUM>),
wherein the surface interface of the metal interfacing layer (<NUM>, <NUM>) is compliant relative to asperities of the surface interface of the CMC turbine blade (<NUM>), a coefficient of friction between the surface interface of the CMC turbine blade (<NUM>) and the metal turbine disk (<NUM>) being about <NUM> or less at an operating temperature between about <NUM> to about <NUM> and a limiting temperature of the metal turbine disk (<NUM>),
wherein the metal interfacing layer (<NUM>) is between about <NUM> and about <NUM> thick.