Patent Publication Number: US-2019170013-A1

Title: Discontinuous Molded Tape Wear Interface for Composite Components

Description:
FIELD 
     The present subject matter relates generally to composite components for gas turbine engines. More particularly, the present subject matter relates to composite components having discontinuous molded tape wear interfaces and methods for manufacturing the same. 
     BACKGROUND 
     A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine general includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere. 
     More commonly, non-traditional high temperature materials, such as ceramic matrix composite (CMC) materials, are being used for various components within gas turbine engines. As CMC materials can withstand relatively extreme temperatures, there is particular interest in replacing components formed of traditional materials within the flow path of the combustion gases with CMC materials. For instance, nozzles, rotor blades, and shrouds of the turbine section of the gas turbine engine are more commonly being formed of CMC materials. As another example, combustion liners of the combustion section are also more commonly being formed of CMC materials. Such CMC components or laminates are generally formed of a plurality of unidirectional plies each formed of a reinforcement material (e.g., fibers) embedded within a ceramic matrix. 
     Within a gas turbine engine, certain CMC components may interface with components formed of other materials, such as e.g., metallic components. For example, a CMC shroud of the turbine section may interface directly with a metallic pin to couple the shroud with a hanger. As another example, a CMC shroud may interface with a metallic bushing or grommet that in turn interfaces with a metallic pin to couple the shroud with a hanger. Interfacing metallic components with CMC components presents a number of challenges. For instance, due to their laminate construction, CMC components can have anisotropic wear characteristics at the CMC-metallic interface, and thus, interface loads are generally not applied over the thickness of the CMC component. This may cause non-uniform wear along the interface, for example. Moreover, CMC components typically have relatively low laminar stress capability at their edges, making the underlying plies of the CMC susceptible to edge loaded chipping and ply delamination. Metallic bushing or grommets have been implemented to address these issues, but they drive a significant space claim and increase complexity and cost. Moreover, metallic bushings or grommets also typically interface directly with the underlying structural plies and thus many of the same challenges may persist. 
     Accordingly, composite components, such as e.g., CMC components, that include features that address one or more of the noted challenges would be useful. In particular, a composite component that includes features that improve the wear interface characteristics of the composite component, among other things, would be beneficial. 
     BRIEF DESCRIPTION 
     Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention. 
     In one exemplary embodiment, a component for a gas turbine engine is provided. The component includes a structural laminate formed of a plurality of plies comprised of reinforcement fibers embedded within a matrix material. Moreover, the component includes a discontinuous molded tape (DMT) attached to the structural laminate and comprising a plurality of reinforcement fragments embedded within a matrix material, wherein the DMT defines a wear interface. 
     In another exemplary embodiment, a method for manufacturing a composite component is provided. The method includes laying up one or more plies to form a structural laminate, the one or more plies comprised of reinforcement fibers embedded within a matrix material. In addition, the method includes attaching a discontinuous molded tape (DMT) to the structural laminate, the DMT comprised of a plurality of reinforcement fragments embedded within a matrix material, wherein the DMT defines a wear interface. 
     In another exemplary embodiment, a method for manufacturing a ceramic matrix composite (CMC) component for a gas turbine engine is provided. The CMC component includes a structural laminate comprised of one of more unidirectional plies. The method includes attaching a discontinuous molded tape (DMT) to the structural laminate, the DMT comprised of a plurality of reinforcement fragments embedded within a matrix material. Further, the method includes curing the structural laminate and the DMT so as to integrally form the DMT with the structural laminate, wherein the DMT defines a wear interface. 
     These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which: 
         FIG. 1  provides a schematic cross-sectional view of an exemplary gas turbine engine according to various embodiments of the present disclosure; 
         FIG. 2  provides an exemplary shroud hanger assembly of the gas turbine engine of  FIG. 1 ; 
         FIG. 3  provides a close up view of Section  3  of  FIG. 2  depicting a discontinuous molded tape defining a wear interface of a shroud of the shroud hanger assembly; 
         FIG. 4  an axial forward-looking-aft view of two wear interfaces depicted as bushings according to various embodiments of the present disclosure; 
         FIG. 5  provides a close up view of one exemplary embodiment of a wear interface defined by a discontinuous molded tape (DMT) according to various embodiments of the present disclosure; 
         FIG. 6  provides a close up view of another exemplary embodiment of a wear interface according to various embodiments of the present disclosure; 
         FIGS. 7 and 8  provide yet another example of a CMC component that includes a wear interface defined by a DMT that is configured to interface with a second component according to various embodiments of the present disclosure; 
         FIG. 9  provides a flow diagram of an exemplary method according to an exemplary embodiment of the present disclosure; and 
         FIG. 10  provides a flow diagram of another exemplary method according to an exemplary embodiment of the present disclosure. 
     
    
    
     Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present invention. 
     DETAILED DESCRIPTION 
     Reference will now be made in detail to present embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the invention. As used herein, the terms “first,” “second,” and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows and “downstream” refers to the direction to which the fluid flows. “HP” denotes high pressure and “LP” denotes low pressure. 
     Exemplary aspects of the present disclosure are directed to composite components that include features that provide improved wear characteristics at the interface between the composite component and a second component. As one example, a ceramic matrix composite (CMC) component can include an integrally formed discontinuous molded tape (DMT) that defines a wear interface between the CMC component and a second component, such as e.g., a pin formed of a metallic material. The DMT can be formed of a plurality of discontinuous reinforcement fragments embedded within a matrix material. The wear interface defined by the DMT may provide improved durability of the CMC component and may facilitate more uniform wear at the interface, among other benefits. Methods for manufacturing composite components having discontinuous molded tape wear interfaces are also provided. 
       FIG. 1  is a schematic cross-sectional view of a gas turbine engine  100  in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of  FIG. 1 , the gas turbine engine  100  is an aeronautical, high-bypass turbofan jet engine configured to be mounted to an aircraft, such as in an under-wing configuration or tail-mounted configuration. As shown in  FIG. 1 , the gas turbine engine  100  defines an axial direction A (extending parallel to or coaxial with a longitudinal centerline  102  provided for reference), a radial direction R, and a circumferential direction C (i.e., a direction extending about the axial direction A; not depicted in  FIG. 1 ). In general, the gas turbine engine  100  includes a fan section  104  and a core turbine engine  106  disposed downstream from the fan section  104 . 
     The exemplary core turbine engine  106  depicted generally includes a substantially tubular outer casing  108  that defines an annular inlet  110 . The outer casing  108  encases, in serial flow relationship, a compressor section  112  including a booster or LP compressor  114  and an HP compressor  116 ; a combustion section  118 ; a turbine section  120  including an HP turbine  122  and a LP turbine  124 ; and a jet exhaust nozzle section  126 . An HP shaft or spool  128  drivingly connects the HP turbine  122  to the HP compressor  116 . An LP shaft or spool  130  drivingly connects the LP turbine  124  to the LP compressor  114 . The compressor section, combustion section  118 , turbine section, and jet exhaust nozzle section  126  together define a core air flowpath  132  through the core turbine engine  106 . 
     Referring still the embodiment of  FIG. 1 , the fan section  104  includes a variable pitch fan  134  having a plurality of fan blades  136  coupled to a disk  138  in a circumferentially spaced apart manner. As depicted, the fan blades  136  extend outwardly from disk  138  generally along the radial direction R. Each fan blade  136  is rotatable relative to the disk  138  about a pitch axis P by virtue of the fan blades  136  being operatively coupled to a suitable actuation member  140  configured to collectively vary the pitch of the fan blades  136 , e.g., in unison. The fan blades  136 , disk  138 , and actuation member  140  are together rotatable about the longitudinal centerline  102  by LP shaft  130  across a power gear box  142 . The power gear box  142  includes a plurality of gears for stepping down the rotational speed of the LP shaft  130  to a more efficient rotational fan speed. 
     Referring still to the exemplary embodiment of  FIG. 1 , the disk  138  is covered by rotatable front nacelle  144  aerodynamically contoured to promote an airflow through the plurality of fan blades  136 . Additionally, the exemplary fan section  104  includes an annular fan casing or outer nacelle  146  that circumferentially surrounds the fan  134  and/or at least a portion of the core turbine engine  106 . Moreover, for the embodiment depicted, the nacelle  146  is supported relative to the core turbine engine  106  by a plurality of circumferentially spaced outlet guide vanes  148 . Further, a downstream section  150  of the nacelle  146  extends over an outer portion of the core turbine engine  106  so as to define a bypass airflow passage  152  therebetween. 
     During operation of the gas turbine engine  100 , a volume of air  154  enters the gas turbine engine  100  through an associated inlet  156  of the nacelle  146  and/or fan section  104 . As the volume of air  154  passes across the fan blades  136 , a first portion of the air  154  as indicated by arrows  158  is directed or routed into the bypass airflow passage  152  and a second portion of the air  154  as indicated by arrow  160  is directed or routed into the LP compressor  114 . The pressure of the second portion of air  160  is then increased as it is routed through the high pressure (HP) compressor  116  and into the combustion section  118 . 
     Referring still to  FIG. 1 , the compressed second portion of air  160  from the compressor section mixes with fuel and is burned within the combustion section  118  to provide combustion gases  162 . The combustion gases  162  are routed from the combustion section  118  along the hot gas path  174 , through the HP turbine  122  where a portion of thermal and/or kinetic energy from the combustion gases  162  is extracted via sequential stages of HP turbine stator vanes  164  that are coupled to the outer casing  108  and HP turbine rotor blades  166  that are coupled to the HP shaft or spool  128 , thus causing the HP shaft or spool  128  to rotate, thereby supporting operation of the HP compressor  116 . The combustion gases  162  are then routed through the LP turbine  124  where a second portion of thermal and kinetic energy is extracted from the combustion gases  162  via sequential stages of LP turbine stator vanes  168  that are coupled to the outer casing  108  and LP turbine rotor blades  170  that are coupled to the LP shaft or spool  130 , thus causing the LP shaft or spool  130  to rotate, thereby supporting operation of the LP compressor  114  and/or rotation of the fan  134 . 
     The combustion gases  162  are subsequently routed through the jet exhaust nozzle section  126  of the core turbine engine  106  to provide propulsive thrust. Simultaneously, the pressure of the first portion of air  158  is substantially increased as the first portion of air  158  is routed through the bypass airflow passage  152  before it is exhausted from a fan nozzle exhaust section  172  of the gas turbine engine  100 , also providing propulsive thrust. The HP turbine  122 , the LP turbine  124 , and the jet exhaust nozzle section  126  at least partially define a hot gas path  174  for routing the combustion gases  162  through the core turbine engine  106 . 
     It will be appreciated that the exemplary gas turbine engine  100  depicted in  FIG. 1  is by way of example only, and that in other exemplary embodiments, the gas turbine engine  100  may have other suitable configurations. Additionally, or alternatively, aspects of the present disclosure may be utilized with any other suitable aeronautical gas turbine engine, such as a turboshaft engine, turboprop engine, turbojet engine, etc. Further, aspects of the present disclosure may further be utilized with industrial and marine gas turbine engines, and/or auxiliary power units. 
     Various components of the gas turbine engine  100  can be formed of a composite material. In particular, components within hot gas path  174 , such as components of the combustion section  118 , HP turbine  122 , and/or LP turbine  124 , can be formed of a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability. Exemplary CMC materials utilized for such components can include silicon carbide, silicon, silica, or alumina matrix materials and combinations thereof. Ceramic fibers can be embedded within the matrix, such as oxidation stable reinforcing fibers including monofilaments like sapphire and silicon carbide (e.g., Textron&#39;s SCS-6), as well as rovings and yarn including silicon carbide (e.g., Nippon Carbon&#39;s NICALON®, Ube Industries&#39; TYRANNO®, and Dow Corning&#39;s SYLRAIVIIC®), alumina silicates (e.g., Nextel&#39;s 440 and 480), and chopped whiskers and fibers (e.g., Nextel&#39;s 440 and SAFFIL®), and optionally ceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite). CMC materials may have coefficients of thermal expansion in the range of about 1.3×10 −6  in/in/° F. to about 3.5×10 −6  in/in/° F. in a temperature range of approximately 1000-1200° F. 
     Some CMC components of the gas turbine engine  100  may interface with components formed of other materials, such as e.g., metallic materials. For instance, CMC shrouds may interface with metallic pins, CMC airfoils may interface with a metallic band, portions of CMC nozzle segments may interface with a metallic support ring, CMC combustor liners may interface with metallic rings, among other possible CMC-metallic interfaces. When a metallic component interfaces directly with a CMC component, the CMC component can experience aggressive anisotropic wear, edge loaded chipping, and inter-ply delamination at the interface, which can directly impact the integrity and durability of the CMC component. In accordance with exemplary embodiments of the present disclosure, CMC components that include features that provide improved wear characteristics at such CMC-metallic wear interfaces are disclosed. Various examples are provided below. 
       FIG. 2  provides an exemplary shroud hanger assembly  176  of the gas turbine engine  100  of  FIG. 1 . The shroud hanger assembly  176  includes a hanger  178  and a shroud  190 . The hanger  178  can be formed of any suitable material, such as e.g., a high temperature nickel-based alloy. The shroud  190  can be formed of a suitable composite material. For this embodiment, the shroud  190  is formed of a CMC material. As shown, the hanger  178  includes a forward hanger  180 , a mid hanger  182 , and an aft hanger  184 . The forward hanger  180  includes a forward arm  186  and the aft hanger  184  includes an aft arm  188 . The forward arm  186  and the aft arm  188  are spaced apart from one another along the axial direction A and are each configured to secure the shroud hanger assembly  176  with a casing (not shown) of the gas turbine engine  100  ( FIG. 1 ). Although the hanger  178  is shown as separate pieces in the depicted embodiment of  FIG. 2 , in alternative exemplary embodiments, the hanger  178  can be a single piece. 
     The shroud  190  is shown in  FIG. 2  positioned outward of the turbine rotor blade  166  of the HP turbine  122  ( FIG. 1 ) along the radial direction R. The shroud  190  is operatively coupled with the hanger  178 . More particularly, the shroud  190  includes a forward flange  192  and an aft flange  194  spaced apart from the forward flange  192  along the axial direction A. The forward flange  192  and the aft flange  194  each define an opening  196 . For this embodiment, the openings  196  are through holes. As shown in  FIG. 2 , metallic pins  198  extend through the openings  196  and couple the shroud  190  with the hanger  178 . More specifically, one of the metallic pins  198  extends through the opening  196  of the forward flange  192  and couples the forward flange  192  with the forward hanger  180  and the mid hanger  182 . Further, one of the metallic pins  198  extends through the opening  196  of the aft flange  194  and couples the aft flange  194  with the mid hanger  182  and the aft hanger  184 . Although not shown, the forward and aft flanges  192 ,  194  may define other openings through which metallic pins may extend to further couple the hanger  178  with the shroud  190 , e.g., at a location spaced apart from the metallic pins  198  shown in  FIG. 2  along the circumferential direction C. 
       FIG. 3  provides a close up view of Section  3  of the shroud hanger assembly  176  of  FIG. 2 . More particularly,  FIG. 3  provides a close up view of the CMC-metallic interface between a composite component  200  and a second component  260 . For this embodiment, the composite component  200  is the shroud  190  and the second component  260  is the metallic pin  198  that couples the forward flange  192  with the forward hanger  180  and the mid hanger  182 . Moreover, in FIG.  3 , portions of the forward flange  192  radially outward of and radially inward of the opening  196  are cutaway so that the underlying structural laminate  210  that forms the forward flange  192  of the shroud  190  may be viewed. 
     As shown in  FIG. 3 , for this embodiment, the structural laminate  210  is formed of a plurality of plies  212  that each include a reinforcement material embedded within a matrix. More specifically, for this embodiment, each ply  212  of the structural laminate  210  is a unidirectional ply having a plurality of SiC filaments or fibers  214  bundled in tows  216  and embedded or encased within a SiC ceramic matrix  218  along a single direction. As depicted in  FIG. 3 , the structural laminate  210  is constructed with unidirectional plies  212  that have alternating fiber orientations. More specifically, the plies  212  include first plies  220  that have fibers  214  oriented along the radial direction R and second plies  222  that have fibers  214  oriented along the circumferential direction C. The second plies  222  alternate with the first plies  220  to form the structural laminate  210 . In this way, structural laminate  210  has a bidirectional laminate construction. 
     As further shown in  FIG. 3 , a discontinuous molded tape (DMT)  240  is integrally formed with the structural laminate  210  and defines a wear interface  242 . For this embodiment, the DMT  240  also further defines opening  196 , as the added volume of DMT  240  decreases the diameter of the opening  196 . The wear interface  242  is configured to interface with the second component  260 , which as noted above, is metallic pin  198 . The wear interface  242  is shown positioned between the structural laminate  210  and the second component  260  and is shown extending circumferentially about one or more surfaces  224  of the structural laminate  210  and along a depth D of the opening  196 . In this way, for this embodiment, the wear interface  242  is a DMT bushing in this embodiment. It will be appreciated that the composite component  200  can include more than one DMT bushing. For instance,  FIG. 4  provides an axial forward-looking-aft view of two wear interfaces  242  defined by DMTs  240  depicted as bushings. 
     Moreover, as shown in the depicted embodiment of  FIG. 3 , the DMT  240  is formed of a plurality of reinforcement fragments  246  embedded within a matrix material  248 . For instance, the reinforcement fragments  246  can be ceramic fiber fragments, such as SiC fiber fragments, and the matrix material  248  of the DMT  240  can be any suitable matrix material, such as a SiC ceramic matrix material. Preferably, the matrix material  248  of the DMT  240  is formed of a matrix material that is compliant with the matrix material  218  of the plies  212  of the structural laminate  210 . Although the reinforcement fragments  246  are shown as discontinuous fragments randomly arranged in matrix material  248  in  FIG. 3 , in alternative exemplary embodiments, the reinforcement fragments  246  can be discontinuous fibers arranged randomly within the matrix material  248 . Further, in some embodiments, the reinforcement fragments  246  of the DMT  240  can include both discontinuous fragments and discontinuous fibers. The discontinuous fragments and fibers provide strength to the wear interface  242  yet still allow the wear interface  242  to have isotropic wear characteristics, which may lead to more uniform wear over the interface, among other benefits. 
     As further illustrated in  FIG. 3 , the wear interface  242  defined by the DMT  240  includes a wear surface  244  that is configured to interface with a wear surface  262  of the second component  260 , which as noted above, is metallic pin  198  in this exemplary embodiment. During operation of the gas turbine engine  100  ( FIG. 1 ), the metallic pin  198  and the forward flange  192  move relative to one another along the axial direction A and the circumferential direction C. As a result, the wear surface  244  of the wear interface  242  defined by the DMT  240  interfaces with the wear surface  262  of the metallic pin  198 . Over time, the metallic pin  198  can wear the forward flange  192 . 
     The wear interface  242  defined by the DMT  240  functions as a buffer or bridge between the metallic pin  198  and the laminate structure  210  of the forward flange  192  of the CMC shroud  190 . Accordingly, the laminate structure  210  does not directly interface with the metallic pin  198 , which protects the underlying laminate structure  210  and reduces the stress on the plies  212 . In addition, as the DMT  240  that defines the wear interface  242  has isotropic wear properties, or wear properties that are the same or substantially the same in all directions, the wear interface  242  distributes the interface load across the thickness of the structural laminate  210  (in this embodiment, the thickness of the structural laminate extends along the axial direction A) and provides uniform multi-directional wear characteristics. In this way, the occurrence of edge-loading driven chipping and inter-ply delamination may be reduced. 
     More specifically, as shown in  FIG. 3 , as noted previously, the fibers  214  of the first plies  220  are oriented along the radial direction R and the fibers  214  of the second plies  222  are oriented along the circumferential direction C. Thus, the fibers  214  of the first and second plies  220 ,  222  are both orthogonal to the axial direction A. As such, without the DMT  240 , the laminate structure  210  would be supported in the axial direction A only by matrix material  218 . The DMT wear interface  242  provides additional structural support to the laminate structure  210  in the axial direction A, and due to its isotropic wear properties, the DMT wear interface  242  distributes the load over the axial thickness of the structural laminate  210 , as opposed a point load at a single ply. Accordingly, the DMT wear interface  242  reduces the occurrence of inter-ply delamination and ply chipping. 
     Moreover, as the wear interface  242  defined by the DMT  240  is integrally formed with the laminate structure  210 , the space claim is reduced and the CMC-metallic wear interface coupling is simplified. Stated differently, as the wear interface  242  defined by the DMT  240  has advantageous wear properties, a metallic bushing or grommet need not be inserted into opening  196  to support the metallic-CMC interface. Thus, the additional space needed for such metallic bushing or grommet to fit into the opening  196  is not needed. 
     In some embodiments, for improved integration or bonding of the DMT  240  with the structural laminate  210 , the opening  196  can be laid up or machined such that the surface area to which the DMT  240  can attach is increased. For example, as shown in  FIG. 3 , the one or more surfaces  224  of the structural laminate  210  to which the DMT  240  is integrally formed can include a first inclined surface  230 , a second inclined surface  232 , and a tip portion  234  connecting the first and second inclined surfaces  230 ,  232  along its circumferential cross section. The first and second inclined surfaces  230 ,  232  converge at the tip portion  234 . Tip portion  234  is shown extending generally parallel to the wear surface  244  of the wear interface  242 . Accordingly, for the depicted embodiment of  FIG. 3 , the wear interface  242  has a circumferential cross section that has an hour glass or butterfly wing shape. In some embodiments, the tip portion  234  may be rounded or pointed, for example, as shown in  FIGS. 5 and 6 , respectively. In yet other embodiments, the wear interface  242  defined by DMT  240  may have a circumferential cross section having other suitable geometries. By laying up or machining the surfaces  224  of the structural laminate  210  such that the structural laminate  210  has first and second inclined surfaces  230 ,  232 , the surface area to which the DMT  240  can attach is increased, and thus, the integration or bonding of the DMT wear interface  242  to the structural laminate  210  may be improved. 
     Stated differently, as shown in  FIG. 3 , the opening  196  has a depth D that extends between a first end  226  and a second end  228 . The wear interface extends between a first side  236  and a second side  238  along at least a portion of the depth D of the opening  196 . For this embodiment, the wear interface  242  extends between the first end  226  and the second end  228  along substantially the entire depth D of the opening  196 . A midline M is defined midway between the first side  236  and the second side  238  of the wear interface  242 . Furthermore, as shown, the one or more surfaces  224  of the structural laminate  210  include one or more inclined surfaces that are inclined with respect to the wear surface  244  of the wear interface  242 . More particularly, as noted above for the depicted embodiment of  FIG. 3 , the one or more inclined surfaces include a first inclined surface  230  and a second inclined surface  232  that are inclined with respect to the wear surface  244  and converge proximate the midline M at tip portion  234 . 
     In some embodiments, the DMT wear interface  242  can have a greater thickness along certain portions of the depth D of the opening  196 , e.g., to better protect certain portions of the underlying laminate structure  210  from being damaged as the wear interface  242  wears over time. For instance, as shown in  FIG. 3 , the wear interface  242  has a thickness T extending between the one or more surfaces  224  of the structural laminate  210  and the wear surface  244  of the wear interface  242  configured to interface with the second component  260 . As shown, the thickness of the wear interface  242  is greater proximate at least one of the first side  236  and the second side  238  of the wear interface  242  than the thickness of the wear interface proximate the midline M. In this way, the wear interface  242  is reinforced with DMT  240  proximate at least one of its sides, which may be particularly advantageous where the wear interface  242  is likely to wear at its sides, such as in the present embodiment depicted in  FIG. 3 . As the metallic pin  198  wears the sides of the wear interface  242  over time, the increased thickness of the DMT wear interface  242  at one of its sides provides an extra buffer of protection so that the structural laminate  210  does not interface directly with the metallic pin  198 . In some embodiments, the thickness of the wear interface  242  is greater proximate both the first side  236  and the second side  238  of the wear interface  242  than the thickness of the wear interface proximate the midline M. 
       FIGS. 7 and 8  provide another example of a CMC component that includes a wear interface defined by a DMT that is configured to interface with a second component. More particularly, for this embodiment, the CMC component  200  is a combustion liner  270  and the second component  260  is a ring  272  ( FIG. 8 ) formed of a metallic material. Further, for this embodiment, the wear interface  242  is a ring contact face of the combustion liner  270  that has advantageous wear properties as described above. The ring contact face or wear interface  242  is integrally formed with one or more surfaces  224  of the structural laminate  210 . Moreover, in  FIG. 8 , a part of the combustor liner  270  is cutaway so that the underlying structural laminate  210  that forms the combustion liner  270  may be viewed. 
     During operation of a gas turbine engine (such as e.g., the gas turbine engine  100  of  FIG. 1 ), the wear interface  242 , or the ring contact face in this embodiment, interfaces with and moves relative to the metallic ring  272  or vice versa along the radial direction R and the circumferential direction C. As the wear interface  242  defined by the DMT  240  is positioned between the structural laminate  210  and the metallic ring  272 , the metallic ring  272  does not interface directly with the structural laminate  210 , thereby protecting the underlying plies  212  of the structural laminate  210 . Moreover, as the DMT  240  has isotropic wear properties, the DMT wear interface  242  distributes the load over the thickness of the structural laminate  210 , as opposed to a single ply. Accordingly, the DMT wear interface  242  reduces the occurrence of inter-ply delamination and ply chipping. Further, the wear interface  242  defined by the DMT  240  offers additional benefits, as described above. 
     Exemplary methods for manufacturing components having DMT wear interfaces will now be provided. In particular,  FIG. 9  provides a flow diagram of an exemplary method for manufacturing a composite component configured to interface with a second component. For instance, the component can be the shroud  190  of  FIGS. 2 and 3  or the combustion liner  270  of  FIGS. 7 and 8  illustrated and described herein. Moreover, the component can be another component of a gas turbine engine, such as e.g., the gas turbine engine  100  of  FIG. 1 . As yet another example, the component can be a CMC airfoil and the second component can be a metallic outer band. As a further example, the composite component can be formed of a polymer matrix composite (PMC) material and the second component can be a metallic component. 
     At ( 302 ), the method ( 300 ) includes laying up one or more plies to form a structural laminate, the one or more plies comprised of reinforcement fibers embedded within a matrix material. For instance, the one or more plies can be unidirectional plies. The reinforcement fibers can be bundled in tows. The matrix material can be a ceramic matrix material or another suitable matrix material, such as e.g., a polymer matrix material. As one example, the structural laminate can be one of the flanges  192 ,  194  of the shroud  190  of  FIGS. 2 and 3 . As another example, the structural laminate can be a portion of the combustion liner  270  of  FIGS. 7 and 8 . 
     At ( 304 ), the method ( 300 ) includes attaching a discontinuous molded tape (DMT) to the structural laminate, the DMT comprised of a plurality of reinforcement fragments embedded within a matrix material. In some implementations, the matrix material of the DMT is a ceramic matrix material or a pre-melt infiltrate matrix material, e.g., a ceramic matrix precursor. The reinforcement fragments can be discontinuous carbon fiber or silicon carbide fragments or can be discontinuous fibers randomly dispersed within the matrix material of the DMT. 
     In some implementations, as shown at ( 306 ), attaching the DMT to the structural laminate includes piping or puttying the DMT along one or more surfaces of the structural laminate to form a desired shape of the wear interface. For instance, the desired shape can be a DMT bushing, such as e.g., the DMT bushings shown in  FIGS. 3 and 4 . Further, when the DMT is piped or puttied along the one or more surfaces of the structural laminate, the DMT can be in a paste form so that the DMT can be shaped to the desired shape. After attaching the DMT to the structural laminate, as shown at ( 308 ), the method further includes subjecting the structural laminate and the DMT to elevated temperatures and pressures in an autoclave. Thus, in such implementations, the paste-like DMT is applied to the structural laminate of the component prior to subjecting the component and the applied DMT to elevated temperatures and pressures in an autoclave. Accordingly, the DMT can be applied to the structural laminate when the structural laminate is a preform or in a preform state. 
     In some implementations, the DMT is formed from waste materials of one or more previously formed composite components. For instance, a cured composite component can be finish machined via a grinding process to form the component to a desired geometry. The composite chips or swarf produced by the grinding operation are mixed with various machining cooling and/or lubrication liquids during grinding, resulting in a slurry composition. The composite pieces are then separated from the waste machining liquids. The composite pieces can include reinforcement, matrix, and/or a combination of reinforcement and matrix materials. Next, the CMC pieces can be embedded within a matrix material to form the DMT. The matrix material can be, for example, a SiC ceramic matrix material. Additional additives can be added to the DMT to adjust its composition and properties. For example, various solvents, such as isopropanol, can be added to the DMT material to adjust the viscosity of the DMT to the desired viscosity and content. 
     In some implementations, as shown at ( 310 ), after laying up the plies to form the structural laminate at ( 302 ), the method ( 300 ) further includes subjecting the structural laminate to elevated temperatures and pressures in an autoclave. As a result, the structural laminate is in the green state. Thereafter, the DMT can be applied to the structural laminate as shown at ( 304 ). As shown at ( 312 ), attaching the DMT to the structural laminate includes nesting the DMT along one or more surfaces of the structural laminate. When the DMT is nested along the one or more surfaces of the structural laminate, the DMT can be a prefabricated member in a solid or semi-solid form and can include a plurality of reinforcement fragments embedded within a matrix material. As one example, the prefabricated member can be a slug shaped as a hollow cylindrical component that can be press fit or interference fit into an opening. In this way, the prefabricated member can be a DMT bushing. As another example, the prefabricated member can be an elongated ring or ring segment configured to be press fit into a recess or along a notch of the composite component. For instance, the prefabricated member can be an elongated ring segment configured to be press fit into a recess or notch of a composite combustion liner. 
     In some implementations, as shown at ( 314 ), nesting the DMT along one or more surfaces of the structural laminate includes drawing a vacuum to drive or pull the prefabricated member along the one or more surfaces of the structural laminate such that the prefabricated member is drawn into contact with the reinforcement material. During the vacuum draw or thereafter, the prefabricated member and the structural laminate can be subjected to elevated temperatures and pressures in the autoclave, as shown at ( 308 ). Further, in such implementations, the matrix material of the prefabricated member can be formed of a ceramic matrix material. In such implementations, when the prefabricated member is nested along the one or more surfaces of the structural laminate, the prefabricated member is a green state prefabricated member. That is, the prefabricated member has been subjected to elevated temperatures and pressures in an autoclave but has not undergone a firing process. 
     At ( 316 ), the method ( 300 ) includes curing the structural laminate and the DMT so as to integrally form the DMT with the structural laminate, wherein the DMT defines a wear interface configured to interface with the interface member. In some implementations, as shown at ( 318 ), curing includes burning out the laminate and the DMT and melt-infiltrating the structural laminate and the DMT so as to integrally form the DMT with the structural laminate. 
     As one example, for CMC components, after processing the structural laminate and the applied DMT in an autoclave to subject them to elevated temperatures and pressures to produce a compacted green state laminate, the green state laminate can be placed in a furnace to burn out excess binders or the like and then can be placed in a furnace with a piece or slab of silicon and fired to melt infiltrate the structural laminate with at least silicon. More particularly, heating (i.e., firing) the green state structural laminate and applied DMT in a vacuum or inert atmosphere decomposes the binders, removes the solvents, and converts the precursor to the desired pyrolyzed material. The decomposition of the binders results in a porous pyrolyzed body. The body may thereafter undergo densification, e.g., melt infiltration (MI), to fill the porosity. In one example, where the pyrolyzed component is fired with silicon, the component can undergo silicon melt-infiltration. However, densification can be performed using any known densification technique including, but not limited to, Silcomp, melt infiltration (MI), chemical vapor infiltration (CVI), polymer infiltration and pyrolysis (PIP), and oxide/oxide processes, and with any suitable materials including but not limited to silicon. In one embodiment, densification and firing may be conducted in a vacuum furnace or an inert atmosphere having an established atmosphere at temperatures above 1200° C. to allow silicon or other appropriate material or combination of materials to melt-infiltrate into the component. 
     Thereafter, the densified laminate and now integrally formed DMT can be finish machined as necessary. For instance, the laminate and/or the DMT can be grinded or otherwise machined, e.g., to bring the laminate and/or the DMT within tolerance and to the desired shape. It will be appreciated that other methods for curing the structural laminate and applied DMT are possible. 
     In some implementations, after laying up the one or more plies to form the structural laminate, the method ( 300 ) further includes machining an opening into the structural laminate, wherein the opening machined into the structural laminate is defined by one or more surfaces of the structural laminate, and wherein the opening has a depth extending between a first end and a second end. In such implementations, during attaching, the DMT is applied to the one or more surfaces of the structural laminate to further define the opening and such that the wear interface extends between a first side and a second side along at least a portion of the depth of the opening. Moreover, a midline is defined between the first side and the second side. In addition, the wear interface has a thickness extending between the one or more surfaces of the structural laminate and a wear surface of the wear interface, and wherein the thickness of the wear interface is greater proximate at least one of the first side and the second side than the thickness of the wear interface proximate the midline. In yet other implementations, the thickness of the wear interface is greater proximate both the first and second sides than the thickness of the wear interface proximate the midline. 
     In some implementations, after laying up the one or more plies to form the structural laminate, the method ( 300 ) further includes machining an opening into the structural laminate. The opening machined into the structural laminate is defined by one or more surfaces of the structural laminate and has a depth extending between a first end and a second end. In addition, in such implementations, during attaching, the DMT is applied to the one or more surfaces of the structural laminate to further define the opening and such that the wear interface extends between a first side and a second side along at least a portion of the depth of the opening and wherein the wear interface has a thickness extending between the one or more surfaces of the structural laminate and a wear surface of the wear interface configured to interface with the second component, and wherein the one or more surfaces of the structural laminate include on or more inclined surfaces that are inclined with respect to the wear surface of the wear interface. 
       FIG. 10  provides a flow diagram of an exemplary method for manufacturing a CMC component for a gas turbine engine. The CMC component is configured to interface with a second component. The CMC component includes a structural laminate that includes one of more unidirectional plies. For instance, the CMC component can be the shroud  190  of  FIGS. 2 and 3  or the combustion liner  270  of  FIGS. 7 and 8  illustrated and described herein. Moreover, the component can be another component of a gas turbine engine, such as e.g., the gas turbine engine  100  of  FIG. 1 . The second component can be a metallic component, for example, such as metallic pin  198  of  FIGS. 2 and 3 . 
     At ( 402 ), the method includes attaching a discontinuous molded tape (DMT) to the structural laminate, the DMT comprised of a plurality of reinforcement fragments embedded within a matrix material. In some implementations, the matrix material is a ceramic matrix material or a pre-melt infiltrate matrix material, e.g., a ceramic matrix precursor. The matrix material of the DMT can alternatively be a matrix material that is compliant with the ceramic matrix material of the CMC component. The reinforcement fragments can be carbon fiber or silicon carbide fragments or can be discontinuous fibers randomly dispersed within the matrix material of the DMT. 
     At ( 404 ), the method includes curing the structural laminate and the DMT so as to integrally form the DMT with the structural laminate, wherein the DMT defines a wear interface configured to interface with the interface member. In some implementations, curing includes burning out the laminate and the DMT and melt-infiltrating the structural laminate and the DMT so as to integrally form the DMT with the structural laminate. 
     In some implementations of method ( 400 ), various aspects discussed above with regard to method ( 300 ) can be implemented in method ( 400 ). 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.