Patent Publication Number: US-2021188717-A1

Title: Reinforced ceramic matrix composite and method of manufacture

Description:
BACKGROUND 
     Exemplary embodiments pertain to the art of ceramic matrix composites. 
     Ceramic matrix composites (CMC) can be formed by infiltrating a preform with a matrix material such as by chemical vapor infiltration to form the matrix. Ceramic matrix composites have high temperature capability and are light weight, and are an attractive material for various applications in which high temperature durability and light weight are desired. Based on these and other features, there remains a desire and need for alternate methods and materials for ceramic matrix composites. 
     BRIEF DESCRIPTION 
     A method of making a ceramic matrix composite is disclosed. According to the method, a first preform comprising fibers is formed, and a second preform including a helical surface portion is inserted into the first preform. The first preform with the inserted second preform is infiltrated with a matrix material comprising a ceramic to form the ceramic matrix composite. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first preform can comprise a three-dimensional woven fiber preform or a stacked fiber layup. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first preform can comprise a stacked layup of fibers including a Z-axis perpendicular to layers in the stacked layup, and the second preform is inserted with a helical axis of said helical surface portion arranged parallel to the Z-axis. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the method can further comprise rotating the second preform about an axis of the helix in a rotational direction that promotes advancement of the helical surface through the first preform. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, axial movement of the second member with said insertion can be equal to axial distance traveled by said helical surface portion in response to said rotation. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the helical surface portion can include a portion arranged as a screw. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the helical surface portion can include a helical portion arranged as a spring. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the second preform can be inserted into the first preform with the helical portion arranged as a spring being under tension, or under compression, or under neutral compression/tension. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the method can further comprise compressing the first preform and inserted second preform. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, compression of the inserted second preform can include a helical compression of the portion of the second preform arranged as a spring. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the second preform can comprise ceramic fibers and an organic polymer resin, and the method includes pyrolyzing the organic polymer resin after compression and before infiltrating the matrix material. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the method can further comprise compressing the first preform before infiltrating the matrix material. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the second preform can comprise ceramic fibers and an organic polymer resin, and the method includes pyrolyzing the organic polymer resin before infiltrating the matrix material. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the method can further include applying an interface coating to the first preform, or to the second preform, or to the first preform and the second preform before infiltrating the matrix material. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, infiltrating can comprise chemical vapor infiltration, atomic layer deposition, polymer infiltration and pyrolysis, and/or melt infiltration. 
     A ceramic matrix composite is also disclosed. The ceramic matrix composite includes a first portion including a matrix comprising a ceramic, and a reinforcement including fibers derived from the first preform in the matrix. A second portion including a helical surface portion interface with the first portion is disposed within the first portion. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the second portion can comprise a matrix comprising a ceramic, and a reinforcement comprising fibers in said matrix, said fibers derived from a second. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the helical surface portion can include a portion arranged as a screw. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the helical surface portion includes a helical portion arranged as a spring. 
     Also disclosed is a gas turbine engine component including a ceramic matrix composite that includes a first portion including a matrix comprising a ceramic, and a reinforcement including fibers derived from the first preform in the matrix. A second portion including a helical surface portion interface with the first portion is disposed within the first portion. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike: 
         FIG. 1  schematically shows a fiber preform; 
         FIG. 2  schematically shows a preform with a helical surface portion; 
         FIG. 3  schematically shows insertion of a helical preform into the preform of  FIG. 1 ; 
         FIG. 4  schematically shows the preform of  FIG. 1  with the helical preform fully inserted; 
         FIG. 5  schematically shows a preform with an inserted helical preform under compression; 
         FIG. 6  schematically shows a fiber preform after compression, ready for infiltration with a matrix material; 
         FIG. 7  schematically shows the fiber preform with infiltrated matrix material; and 
         FIG. 8  is a partial cross-sectional view of a gas turbine engine. 
     
    
    
     DETAILED DESCRIPTION 
     A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the Figures. 
     A ceramic matrix composite (“CMC”) can be made by infiltrating a preform using matrix material such as by chemical vapor infiltration. The fiber preform contributes beneficial mechanical properties to the composite by providing reinforcement for a matrix material. Mechanical properties of interest include but are not limited to interlaminar shear strength (“ILS”) and interlaminar tensile strength (“ITS”). CMC&#39;s can be used for high-temperature applications (e.g., 2200° F. and above), and may be designated as ultra-high temperature ceramic matrix composites (“UHT-CMC”). 
     Exemplary CMC materials can include silicon-containing, or oxide containing matrix and reinforcing materials. Some examples of CMCs include, but are not limited to, materials having a matrix and reinforcing fibers comprising non-oxide silicon-based materials such as silicon carbide, silicon nitride, silicon oxycarbides, silicon oxynitrides, silicides, and mixtures thereof. Examples include, but are not limited to, CMCs with a silicon carbide matrix and silicon carbide fiber; silicon nitride matrix and silicon carbide fiber; and silicon carbide/silicon nitride matrix mixture and silicon carbide fiber. Furthermore, CMCs can have a matrix and reinforcing fibers including oxide ceramics. Specifically, the oxide-oxide CMCs may include a matrix and reinforcing fibers comprising oxide-based materials such as aluminum oxide (Al 2 O 3 ), silicon dioxide (SiO 2 ), yttrium aluminum garnet (YAG), aluminosilicates, or mixtures comprising any of the foregoing. Aluminosilicates can include crystalline materials such as mullite (3Al 2 O 3  2SiO 2 ), as well as glassy aluminosilicates. Other ceramic composite materials in addition to or in combination with silicon or oxygen may be used, including carbon, carbides (e.g., zirconium carbide, hafnium carbide, boron carbide), nitrides, or other ceramic materials, alone or in combinations including any of the materials noted above. 
     Referring to  FIG. 1 , a first preform  100  is shown according to various embodiments. As shown in  FIG. 1 , the preform  100  includes a plurality of fiber tows or plies  110 ). Each fiber tow or ply  110  can include a plurality of individual fibers (e.g., from 100 to 500 fibers per tow). In various aspects, the fiber tows or plies  110  can be arranged in a 1-D tape or a 2-D fabric, and layers of tape or woven or unwoven fabric can be stacked in layup (including stacked layups of 2-D fabrics such as a woven fabric and also stacked layups of 1-D tape or 2-D non-woven fabric) in order to form the preform  100 . Alternatively, a 3-D fabric can be woven or otherwise arranged from continuous or non-continuous fibers. The fibers can include any material typically used in CMC processing, such as carbon, alumina, silicon carbide, silicon nitride-carbide, zirconia, zirconium carbide, boron carbide, glass, or mullite. In some aspects, such as for high-temperature applications like gas turbine engine hot section components, fibers capable of withstanding specified temperatures can be used, including but not limited to ceramic fibers such as silicon carbide, alumina, zirconia, boron carbide, zirconium carbide, hafnium carbide. In some aspects, the fibers can include an interface coating (e.g., a boron nitride coating on a silicon carbide fiber). However, in other aspects the fibers can be bare fibers, in which case the fibers would not include an interface coating. 
     With reference now to  FIG. 2 , a second preform including a helical surface portion is shown in the form of a helical compression spring  120 . Of course, the helical compression spring  120  is an example embodiment, and the second preform can include other types of helical surface portions. For example, the helical compression spring  120  includes a curvilinear-extending spring member  130  that extends along a helical path about an axis  140 . The compression spring includes a hollow space at and about the axis  140 , but the second preform could be solid at the axis  140  or include a solid annulus about the axis  140 , with the helical surface portion arranged in the form of a shaft twisting spiral rectangular shaft, or arranged as helical threads extending from a solid linear shaft or annulus. In some aspects, the helical surface portion of the second member can be arranged to function as a screw. As used herein, the term “screw” refers to the simple machine its operation that converts between linear motion and rotational motion, and between rotational force (i.e., torque) and linear force. The compression spring  120  is an example of a screw, and can be inserted into the preform  100  by rotating the compression spring  120  so that leading tip  150  follows a helical path into the preform. 
     The compression spring  120  can be formed from materials to provide a target resiliency for conversion of stress and strain. In some aspects, the compression spring  120  can be formed from fibers and an elastomeric or flexible matrix polymer. The matrix polymer can be any type of polymer, including but not limited to epoxy, polyurethane, a polystyrene, polypropylene, polyethylene, etc. The fibers can be any of the fibers disclosed above, and can be the same as or of different composition compared to the fibers used in the first preform. The fibers in the second preform can be chopped and dispersed in the matrix polymer, or they can be extruded as continuous fibers extending co-linearly along the helical extension of the spring member  130 . 
     In some aspects, the compression spring  120  can be inserted in a Z-direction into the preform  100  as schematically shown in  FIG. 3  (e.g., by screwing in with rotation in the direction of arrow  160 ). In some aspects, this rotational insertion can provide a technical effect of promoting reduced breakage of fibers during insertion compared to other reinforcement techniques such as z-pinning. In some aspects, the compression spring  120  can be inserted into the preform  100  so that compression spring is under tension. In some aspects, the compression spring  120  can be inserted into the preform  100  so that the compression spring is under compression. In some aspects, the compression spring  120  can be inserted into the preform  100  so that compression spring is under neither compression nor tension (i.e., neutral compression/tension). 
     During some aspects of processing (e.g., compression of the preforms prior to and during matrix infiltration and consolidation), the compression spring  120  may undergo helical deformation in response to an application of linear stress. As used herein, the term “helical deformation” means changes in the parameters of the helix such as pitch (axial distance corresponding to one complete turn of the helix), arc length, curvature, torsion to accommodate stress applied along the axis  140  and resulting deformation of a spring or spring-like structure in response to axial stress. In some aspects, the second preform can utilize an elastomeric matrix polymer to provide a resilient compression spring  120 . However, in some aspects, one-way deformation during processing, but absorption of compressive load by a resilient spring including shape recovery may not be needed if the performance of the preform is satisfied by placing reinforcing fibers along the helical insertion path of the second member. In such cases, the polymer matrix for the second member need not be elastomeric, but can instead provide a level of flexibility (in combination with the mechanical properties of the fibers in the second preform), that is sufficiently low to maintain its shape during insertion and to prevent deflection of the leading tip  150 , and also sufficiently high to tolerate deformation such as in response to compression along the axis  140 .  FIG. 4  shows the compression spring  120  fully inserted in the Z-direction into the preform  100 . 
     In  FIG. 5 , the preform  100  with the inserted second preform in the form of the compression spring  120  is shown disposed between tooling  170  and subjected to a compression load in the direction of arrows  180 . Compression of a preform before matrix infiltration/consolidation can beneficially promote high fiber volume in final composite, but can cause problems for conventional z-pinning schemes, including but not limited to breaking of fibers, and crimping or other damage to fibers, which can adversely impact the mechanical and performance properties of the composite. However, the compression spring  120  accommodates the stress from the compressive load  180  with helical deformation including a change in helical pitch with shorter axial distances between each turn of the helix as shown in  FIG. 5 , which can promote reduced susceptibility to breaking of fibers, crimping, or other damage to fibers. A final preform structure after compression, with three compression springs  120  inserted into the preform  100 , is shown in  FIG. 6 . 
     Any polymer resin in the second preform (and also any polymer resin in the first preform) can be pyrolyzed thermally decomposed, also known as burn-out, before infiltration of the matrix material. In cases where a fiber matrix interface coating is applied to the preforms, any polymer resin can be thermally decomposed before application of the fiber matrix interface coating. Thermal decomposition of polymer resin can be performed at temperatures of 200-700° C. in an oxidizing, inert, or vacuum environment. 
     In some aspects, a fiber/matrix interface coating can optionally be applied before matrix consolidation. In some aspects, a fiber/matrix interface coating can promote reduction in formation or propagation of cracks by allowing the fiber to slide in the interface coating at the fiber-coating interface or by allowing the coated fiber to move in the matrix by sliding at the coating-matrix interface. The choice of material for a fiber/matrix interface coating depends on the materials of the fibers and the matrix. For example, in the case of silicon carbide fibers and a silicon carbide matrix, the fibers may be coated with boron nitride as a fiber/matrix interface coating. A matrix/fiber interface coating can be applied by known means such as chemical vapor infiltration or atomic layer deposition, which can promote a uniform thickness coating and can infiltrate fine spaces in the preform(s). 
     With any polymers thermally decomposed, and a fiber/matrix interface coating applied (if desired), the preforms are ready for matrix consolidation. Matrix consolidation is typically performed by infiltration of the preforms with a matrix material. For example, in some aspects, the preforms can be treated with chemical vapor infiltration (CVI) in which a gas comprising a matrix material precursor (e.g., CH 3 SiCl 3 —H 2 ) in a carrier gas at elevated temperature is infiltrated into the void space in the preforms and consolidates there to form a ceramic matrix material (e.g., SiC). The resultant composite, including fiber tows or plies  110  derived from the preform  100 , fiber reinforcements derived from second preforms in the form of compression springs  120 , and matrix material  190  is shown in  FIG. 7 . 
     The method described herein can be used to prepare a variety of components comprising ceramic matrix composites such as components in the aviation industry, marine industry and energy industry. Exemplary components include components for gas turbine engines, such as in high pressure compressors (“HPC”), fans, boosters, high pressure turbines (“HPT”), and low pressure turbines (“LPT”). More specifically exemplary components include combustion liners, shrouds, nozzles, and blades. In addition to the above-referenced technical effect of reduced fiber breakage, the reinforcement provided to the composite by the second preform can promote additional technical effects, including but not limited to promotion of improved interlaminar shear strength (“ILS”) and/or improvement of interlaminar tensile strength (“ILT”). 
       FIG. 8  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include other systems or features. The fan section  22  drives air along a bypass flow path B in a bypass duct, while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     The exemplary engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided, and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a low pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  and high pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . An engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The engine static structure  36  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
     The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (TSFC′)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)] 0.5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 m/sec). 
     The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. 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, element components, and/or groups thereof. 
     While the present disclosure has been described with reference to an exemplary embodiment or embodiments, 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 present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.