Abstract:
The integral layer provides a ductile interface for attachment locations of a turbine engine component where a metallic surface is adjacent the attachment location. The ductile layer provides a favorable load distribution through the composite at the attachment location, and eliminates the need for a metallic shim.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    This application is a divisional of and claims priority to U.S. Utility application Ser. No. 11/938,349, Attorney Docket No. 121500 (07783-0173-DIV1), filed on Nov. 12, 2007, entitled “CERAMIC COMPOSITE WITH INTEGRATED COMPLIANCE/WEAR LAYER”, which is a divisional of and claims priority to U.S. Utility application Ser. No. 11/025,222, Attorney Docket No. 121500 (07783-0173), filed on Dec. 29, 2004, entitled “CERAMIC COMPOSITE WITH INTEGRATED COMPLIANCE/WEAR LAYER”, which is now U.S. Pat. No. 7,329,101, issued Feb. 12, 2008, both of which are incorporated herein by reference in their entirety. 
     
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
       [0002]    This invention was made with government support under Contract No. N00421-00-3-0536. The United States government may have certain rights to the invention. 
     
    
     FIELD OF THE INVENTION 
       [0003]    The present invention relates generally to ceramic matrix turbine engine components, and more particularly, to an interface layer integrated into the ceramic composite during manufacture to reduce wear and provide a more favorable load distribution. 
       BACKGROUND OF THE INVENTION 
       [0004]    A number of techniques have been used in the past to manufacture turbine engine components, such as turbine blades using ceramic matrix composites (CMC). One method of manufacturing CMC components, set forth in U.S. Pat. Nos. 5,015,540; 5,330,854; and 5,336,350; incorporated herein by reference and assigned to the Assignee of the present invention, relates to the production of silicon carbide matrix composites containing fibrous material that is infiltrated with molten silicon, herein referred to as the Silcomp process. The fibers generally have diameters of about 140 micrometers or greater, which prevents intricate, complex shapes, such as turbine blade components, to be manufactured by the Silcomp process. 
         [0005]    Another technique of manufacturing CMC turbine blades is the method known as the slurry cast melt infiltration (MI) process. A technical description of such a slurry cast MI method is described in detail in U.S. Pat. No. 6,280,550 B1, which is assigned to the Assignee of the present invention and which is incorporated herein by reference. In one method of manufacturing using the slurry cast MI method, CMCs are produced by initially providing plies of balanced two-dimensional (2D) woven cloth comprising silicon carbide (SiC)-containing fibers, having two weave directions at substantially 90° angles to each other, with substantially the same number of fibers running in both directions of the weave. 
         [0006]    Generally, such turbine components require attachment to adjoining metallic hardware and/or metallic surfaces. Two disadvantages associated with attaching a CMC to metallic hardware are the wear of the metallic hardware by the hard, abrasive ceramic material surface, and the lack of load distribution in the CMC. Load distribution is critical in blade dovetail/disk interfaces. Typically, metallic shims or ceramic cloth have been interposed between the CMC and metallic surfaces to improve load distribution. Wear is typically handled by the application of coatings to the metallic hardware or coatings to the blade attachment surfaces as set forth in U.S. Pat. No. 5,573,377, incorporated herein by reference and assigned to the Assignee of the present invention. 
         [0007]    What is needed is a method of manufacturing CMC turbine engine components that provides an interface layer on the CMC to improve load distribution within the CMC and reduce metallic wear. A favorable method would apply the interface layer during densification of the CMC. 
       SUMMARY OF THE INVENTION 
       [0008]    Improvements in manufacturing technology and materials are the keys to increased performance and reduced costs for many articles. As an example, continuing and often interrelated improvements in processes and materials have resulted in major increases in the performance of aircraft gas turbine engines, such as the improvements of the present invention. The present invention is directed to a method for manufacturing a turbine engine component made from a ceramic matrix composite (CMC) by incorporating an interface layer into the attachment contacting surfaces of the CMC and the resulting turbine engine component. The present invention produces a component that does not require attachment hardware to be wear coated and eliminates the need for load distributing shims during manufacture, thereby improving the functionality of the component. 
         [0009]    The present invention is directed to a turbine engine component including a body composed of a ceramic matrix material. The body includes an attachment location, the attachment location defining at least a portion of a surface of the component, and the attachment location being configured to be adjacent a second turbine engine component during operation of a turbine engine. An interface layer defines at least a portion of the surface of the attachment location and has a matrix side and an outer side. The interface layer is interposed between the body and the second turbine engine component during operation of a turbine engine, the matrix side of the interface layer being metallurgically bonded to the body. 
         [0010]    The present invention is also directed to a method of manufacturing a turbine engine component that includes the steps of: providing a plurality of ceramic fibers; coating a preselected portion of the fibers with a layer of boron nitride to form a plurality of coated fibers, wherein the coating includes using chemical vapor infiltration; laying up the plurality of fibers in a preselected arrangement to form a component preform; partially infiltrating the component preform using a carbon-containing slurry; further infiltrating the component preform with at least silicon to form a ceramic matrix composite aircraft engine component; and bonding an interface layer to an attachment location defined by a surface of the ceramic matrix composite aircraft engine component, wherein the interface layer contains at least silicon. 
         [0011]    The present invention is further directed to a method of manufacturing a ceramic matrix composite aircraft engine component. The method includes the steps of: providing a plurality of unidirectional prepreg ceramic fiber plies, the plies comprising coated prepreg ceramic fiber tows; laying up the plurality of prepreg ceramic fiber plies in a preselected arrangement to form a component shape; overlying at least a portion of the outside surface of the component shape with at least one matrix ply; heating the component shape to form a preform; infiltrating the preform with at least silicon to form a ceramic matrix composite component, wherein the component includes an attachment area defining a surface of the component; and forming an interface layer on the attachment area, wherein the interface layer includes at least silicon, and the interface layer and the attachment area are integrated. 
         [0012]    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. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0013]      FIG. 1  is an exemplary partial sectional perspective view of a composite blade of the present invention assembled in a dovetail slot of a gas turbine engine rotor. 
           [0014]      FIG. 2  is a partial sectional view of the blade and rotor of  FIG. 1  of the present invention. 
           [0015]      FIG. 3  is an enlarged partial sectional view of the blade and rotor of  FIG. 1 , illustrating the interface layer, exaggerated for clarity, of the present invention. 
           [0016]      FIG. 4  is similar to  FIG. 3 , but illustrating a prior art wear coating and shim. 
           [0017]      FIG. 5  is a flow chart illustrating a slurry cast MI method of manufacture of the present invention to produce a CMC turbine blade with an integral interface layer. 
           [0018]      FIG. 6  is a flow chart illustrating an alternate slurry cast MI method of manufacture of the present invention to produce a CMC turbine blade with an integral interface layer. 
           [0019]      FIG. 7  is an exemplary enlarged sectional view of a gas turbine engine shroud incorporating an embodiment of the invention. 
           [0020]      FIG. 8  is an exemplary enlarged sectional view of a gas turbine engine combustor liner incorporating an embodiment of the invention. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0021]      FIG. 1  depicts an exemplary aircraft engine blade  10 . In this illustration, a turbine blade  10  is constructed of a ceramic matrix composite material. The turbine blade  10  is mounted to a turbine disk  12  in a dovetail slot  14 . The turbine blade  10  includes an airfoil  16 , against which a flow of hot exhaust gas is directed, and a dovetail  18 , also referred has a root or splayed base, that extends from the airfoil  16  and engages the dovetail slot  14 . 
         [0022]    Referring now to  FIG. 2 , which is an example of an enlarged sectional view of a CMC blade  10  and disk  12 , the contacting surfaces thereon are described in greater detail. The blade  10  includes a plurality of plies,  20  and  22 , which have been, bonded together, such as by processes well known in the art. Plies  22  are bonded to a root core  24 . The lower end of blade  10  is defined in part by a root surface  30  and a matrix surface  32 . Dovetail slot  14  of disk  12  is defined by a mating surface  34 . 
         [0023]      FIG. 3  illustrates a portion of the blade  10  and disk  12  portions of  FIG. 2  to include an interface layer  38  defined by an outer surface  40 , or side, and a matrix surface  42 , or side. The matrix surface  42  faces the surface of blade  10 . Interface layer  38  is interposed between blade  10  and disk  12  such that outer surface  40  and mating surface  34  of disk  12  bindingly contact. Interface layer  38  is preferably integrated to dovetail  18  of the blade  10 , as discussed herein. The interface layer  38  may be composed of silicon. Alternatively, the interface layer  38  can include a gradient layering of SiC and silicon. Preferably, the concentration of silicon toward the outer surface  40  is increased in comparison to the matrix surface  42 , silicon being of a lesser degree of hardness than the blade  10 . Preferably, outer surface  40  is composed of up to about 80 percent silicon. However, the concentration of silicon can range from about 20 to about 80 percent by volume, and can additionally include other materials, such as boron nitride (BN), silicon carbide (SiC), molybdenum disilicide (MoSi 2 ) to comprise the remainder. The methods by which these materials are applied and their constructions will be discussed in additional detail below. 
         [0024]    As shown in  FIG. 4 , a prior art assembly of a blade and disk is illustrated to include a wear coat  50  and a shim  52 . Typically, a wear coat  50  is applied to the dovetail mating portion of a blade  10  due to the differences in surface hardness of blades and disks that promote wear on the softer surface. A shim  52  may also be present to reduce wear or to provide a more favorable load distribution between blade  10  and disk  12 . 
         [0025]      FIG. 5  is a flow chart illustrating a slurry cast MI method of manufacture of the present invention to produce a component for use in a heated environment, such as a CMC turbine blade. Heated environment, as used herein refers to temperatures in excess of at least 1,000° F. The initial step  100  of the process preferably includes laying up a preselected number of biased SiC-containing cloth plies of preselected geometry in a preselected arrangement to form a turbine blade shape. In a preferred embodiment, there are a preselected number of fiber tows woven in the weft direction, the woven tows being sufficient to allow the SiC cloth to be handled and laid up without falling apart. 
         [0026]    Once the plies are laid up, the next step  110  includes rigidizing the turbine blade shape by applying boron nitride (BN), using a chemical vapor infiltration (CVI) process as is known in the art, forming a rigid coated turbine blade preform. In an alternate embodiment, an additional layer(s) of silicon doped boron nitride (Si-doped BN) or siliconized BN, silicon nitride (Si 3 N 4 ) and silicon carbide (SiC) layers can be applied over the BN layer. 
         [0027]    The next step  120  includes infiltrating the coated turbine blade preform by introducing a carbon-containing slurry, typically including a polymer which is a carbon yielding polymer, carbide powder and other powders as is known in the art, into the porosity of the coated turbine blade preform. The next step  130  includes further infiltrating the turbine blade preform with at least silicon, and preferably boron doped silicon, through an MI process, as known in the art, forming a SiC/SiC CMC turbine blade. In step  140 , additional silicon is built up on the attachment location areas to construct the interface layer  38 . Preferably, interface layer  38  is about 2 to about 16 mils thick, and even more preferably, interface layer  38  is about 2 to about 4 mils thick. This build up of silicon is preferably accomplished by melt infiltrating additional silicon to the desired attachment locations which yields SiC. 
         [0028]    Alternately, plies or tapes can be constructed, referred to as matrix tapes or matrix plies, which are applied to portions of the plies forming the preform to obtain the desired surface finish of the resulting component. In one embodiment of the present invention that is directed to turbine blades, referring to  FIG. 6 , step  200  includes both steps  100  and  120  of  FIG. 5  as previously discussed. Preferably, in step  200  a preselected number of unidirectional prepreg ceramic fiber plies comprising coated prepreg ceramic fiber tows are used. Once step  200  is performed, step  205  includes laying up a matrix ply or plies over selected portions of the turbine blades to form wear surfaces of the turbine blade shape, although the matrix plies can be applied to overlay the entire outer surface of the turbine blade shape, if desired. Further, step  230  includes further infiltrating the turbine preform with at least silicon to form a SiC/SiC CMC turbine blade. However, due to the matrix ply construction, which can include silicon carbide powder and molybdenum powder or a combination thereof, the molybdenum and silicon react to form molybdenum disilicide (MoSi 2 ). The MoSi 2  is easier to machine, is less rigid, having a lower Young&#39;s modulus, and is more wear compatible. Alternately, in place of the silicon carbide powder in the matrix tape or ply, BN, Si-doped BN, Si 3 N 4  or a combination thereof can be used. Due to the lubricity of BN, the wear surface is improved as the proportion of BN is increased. With these modified matrix plies, once processing has been completed, about 15 percent by volume is SiC, about 5 percent silicon by volume, with the remainder up to about 80 percent by volume of silicon metal, MoSi 2 , BN and alloys thereof. In one embodiment, a gradient of silicon metal, MoSi 2 , BN is achieved along the wear surface of the interface layer, and preferably a higher concentration of SiC faces toward the matrix side when compared to the outer side of the interface layer of the component. 
         [0029]    It is to be understood that the above method can be used with existing CMC melt infiltration components. 
         [0030]    In this manner, interface layer  38  is provided for components, such as composite turbine engine components. The silicon of interface layer  38  has a degree of hardness that is more compatible with adjoining metallic hardware or disks, thereby reducing wear. The silicon of interface layer  38  also provides a thickness of material that is compliant, thereby distributing the loadings between a large area of adjoining surfaces, such as from the mating surface  34  of disk  12  to the plies of a ceramic composite. 
         [0031]    It is appreciated that the interface coating  38 , as described herein, reduces the relative wear experienced at contacting surfaces, and hence, reduces the need for wear coatings, although the dovetail slot  14  may be wear coated at mating surface  34  to provide a sacrificial layer to reduce wear on disk  12 . The present invention 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. An exemplary gas turbine shroud  70  is shown at  FIG. 7 . As can be seen in  FIG. 7 , the shroud  70  includes an interface layer  72  at the attachment point between the shroud  70  and a second gas turbine component  74  An exemplary gas turbine combustor liner  80  is shown at  FIG. 8 . As can be seen in  FIG. 8 , the combustor liner  80  includes an interface layer  82  at the attachment point between the combustor liner  80  and another gas turbine component  84 . Additionally, the wear surfaces can in addition to contact surfaces between different components, but can also include lining apertures used for structural fasteners. 
         [0032]    While the invention has been described with reference to a preferred embodiment, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this invention, but that the invention will include all embodiments falling within the scope of the appended claims.