Ceramic matrix composite component repair and joining

Abstract

A method for joining ceramic matrix composite (CMC) components includes placing a plurality of bond layers onto a particular region such that a respective fiber direction of each of the plurality of bond layers is not in the same plane as a fiber direction of one or more composite layers of the particular region. A respective fiber direction of each of the plurality of bond layers is different from a respective fiber direction of each adjacent bond layer of the plurality of bond layers.

Description

Technical Field

[0001] The present disclosure relates to ceramic matrix composite (CMC) components and their repair. Background Art

[0002] Gas turbine engines have multiple components made of CMC. CMC provides high temperature capabilities that can be used in aerospace applications. Examples of CMC components include turbine blades, nozzles, and shrouds. Inside a turbine engine, a nozzle is formed by multiple airfoils joined to a band. Generally, CMC materials include ceramic fibers embedded in a matrix material such as silicon carbide (SiC), silicon, silica, alumina, or combinations thereof. Layers of CMC material can be placed to form a preform component, which can then be heat treated, such as cured or burned out to produce a high carbon residue in the preform, and then chemically treated, such as melt infiltrated with silicon, to obtain a component formed of CMC material having a desired chemical composition.

[0003] CMC structures are prone to damage during operation. For example, a CMC structure may come into contact with a rigid object and may be pierced or otherwise damaged. Another example is that engine environmental conditions may cause wear of the CMC structure. There is a need for CMC repair products and methods for repairing CMC components. Brief Description of the Drawings

[0004] A complete and enabling disclosure of the present disclosure, including its best mode, is set forth in the specification for those of ordinary skill in the art, with reference to the drawings, in which:

[0005] Figure 1 is a cross-sectional view of an exemplary gas turbine engine.

[0006] Figure 2 is Figure 1 a cross-sectional view of a nozzle of the gas turbine engine.

[0007] Figures 3A - 3B is Figure 2 an enlarged view of a repair area of the nozzle.

[0008] Figure 4 is a perspective view of a layer stack applied to the repair area.

[0009] Figures 5A - 5C is a cross-sectional view of a bonding layer of the layer stack.

[0010] Figure 6 is an enlarged view of another nozzle.

[0011] Figure 7 is a block diagram of a method for repairing a CMC component such as Figure 2 the nozzle. Detailed implementation manners

[0012] Reference will now be made in detail to the current embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and alphabetical names to refer to features in the drawings. Similar or like names in the drawings and the specification are used to refer to similar or like parts of the present disclosure.

[0013] As used herein, the term "exemplary" means "serving as an example, instance, or illustration". Any embodiment described herein as "exemplary" should not necessarily be construed as more preferred or advantageous than other embodiments. Additionally, unless otherwise specifically indicated, all embodiments described herein should be considered exemplary.

[0014] The singular forms "a", "an", and "the" include plural referents unless the context clearly dictates otherwise.

[0015] In the context of terms such as "at least one of A, B, and C", the term "at least one" refers to only A, only B, only C, or any combination of A, B, and C.

[0016] The phrases "from X to Y" and "between X and Y" each refer to a numerical range that includes the endpoints (i.e., a numerical range that includes X and Y).

[0017] "Fiber direction" is an angle defined between the layer fibers and a reference line (such as the default axes in a two-dimensional coordinate system). When the angular difference between two fiber directions is non-zero, the fiber direction of the first layer and the other fiber direction of the second layer are "off-plane".

[0018] Unless otherwise specified, the terms "first", "second", etc. are used herein only as labels and are not intended to impose ordinal, positional, or hierarchical requirements on the items referred to by these terms. Additionally, the mention of, for example, a "second" item does not require or preclude the existence of, for example, a "first" or lower-numbered item or a "third" or higher-numbered item.

[0019] As used herein, a ceramic matrix composite or “CMC” refers to a class of materials that includes a reinforcement (e.g., reinforcing fibers) surrounded by a ceramic matrix phase. Generally, the reinforcing fibers provide structural integrity to the ceramic matrix. Some examples of the matrix materials of CMCs can include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, alumina (Al2O3), silica (SiO2), aluminosilicates, or mixtures thereof), or mixtures of them. 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) can also be included in the CMC matrix. Some examples of the matrix materials of CMCs can include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, alumina (Al2O3), silica (SiO2), aluminosilicates, or mixtures thereof), or mixtures of them. 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) can also be included in the CMC matrix.

[0020] Some examples of the reinforcing fibers of CMCs can include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), non-oxide carbon-based materials (e.g., carbon), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, alumina (Al2O3), silica (SiO2), aluminosilicates such as mullite, or mixtures thereof), or mixtures of them.

[0021] Generally, a particular CMC can be referred to by the combination of its fiber type / matrix type. For example, C / SiC represents carbon fiber reinforced silicon carbide; SiC / SiC represents silicon carbide fiber reinforced silicon carbide, SiC / SiN represents silicon carbide fiber reinforced silicon nitride; SiC / SiC-SiN represents silicon carbide fiber reinforced silicon carbide / silicon nitride matrix mixture, etc. In other examples, a CMC can consist of a matrix and reinforcing fibers, and the reinforcing fibers include oxide-based materials such as alumina (Al2O3), silica (SiO2), aluminosilicates and mixtures thereof. Aluminosilicates can include crystalline materials such as mullite (3Al2O3 2SiO2), as well as vitreous aluminosilicates.

[0022] In some embodiments, the reinforcing fibers may be bundled or coated before being incorporated within a matrix. For example, fiber bundles may be formed into reinforcing tapes, such as unidirectional reinforcing tapes. Multiple tapes may be placed together to form a preform component. The fiber bundles may be impregnated with a slurry composition either before or after forming the preform. The preform may then undergo a heat treatment, such as curing or burnout to produce a high carbon residue in the preform, and subsequent chemical treatment, such as melt infiltration with silicon, to obtain a component formed from a CMC material having a desired chemical composition.

[0023] Such materials, as well as certain monolithic ceramics (i.e., ceramic materials without reinforcing materials), are particularly suitable for high temperature applications. Additionally, compared to superalloys, these ceramic materials are lightweight but can still provide strength and durability to the components made therefrom. Accordingly, such materials are currently being considered for use in many gas turbine components used in the high temperature sections of gas turbine engines, such as airfoils (e.g., turbines and vanes), burners, shrouds, and other similar components that would benefit from the lighter weight and higher temperature capabilities that these materials can provide.

[0024] The terms “upstream” and “downstream” refer to the relative direction with respect to the flow of fluid in a fluid path. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction towards which the fluid flows.

[0025] The terms “low,” “high,” or their respective comparatives (e.g., lower, higher, where applicable) refer to the relative speed or pressure within the engine, unless otherwise specified. For example, a “low pressure turbine” operates at a pressure that is generally lower than that of a “high pressure turbine.” Alternatively, unless otherwise specified, the above terms may be understood in their superlative form. For example, a “low pressure turbine” may refer to the turbine within the turbine section having the lowest maximum pressure, while a “high pressure turbine” may refer to the turbine within the turbine section having the highest maximum pressure. The engines of the present disclosure may also include a medium pressure turbine, such as an engine having three spools.

[0026] The present disclosure generally relates to the repair and manufacture of CMC components (such as nozzles) in gas turbine engines. CMC materials are more frequently used in various high temperature applications. For example, since CMC materials can withstand relatively extreme temperatures, particular attention is paid to replacing components within the combustion gas flow path of a gas turbine engine with components made from CMC materials. Generally, CMC materials include ceramic fibers embedded within a matrix material (such as silicon carbide (SiC), silicon, silica, alumina, or combinations thereof). Layers of CMC material may be placed to form a preform component, which may then be heat treated (e.g., cured or burned out to produce a high carbon residue in the preform), and then chemically treated (e.g., melt infiltrated with silicon), to obtain a component formed from a CMC material having a desired chemical composition.

[0027] During operation, CMC components may need to be repaired because some parts of the CMC components wear. Typically, to repair these parts, the layers are folded around a radius to address the change in the fiber direction of the layers. In this configuration, when the layers do not extend along the folding direction, they may fall off, creating gaps on the concave geometry. The gaps may reduce the layering, increasing the variability of the structural strength during infiltration.

[0028] These parts are made by folding a joining layer into the repair area, with the joining layer positioned such that the fiber direction in the joining layer is different from the fiber direction of the other parts of the CMC component. These non-coplanar fiber directions in the joining layer improve the layering and order and create an additional retention mechanism to improve the repair of the CMC component. Additionally, the different fiber directions allow for improved infiltration of the repair area, brazing or green-state joining in the repair area for subsequent repair operations.

[0029] Referring now to the drawings, where like numerals represent like elements in all the figures, Figure 1 is a cross-sectional view of an exemplary gas turbine engine 10. The gas turbine engine 10 defines a centerline axis 12. The gas turbine engine 10 includes a turbine 14 and a fan section 16 positioned upstream thereof. The turbine 14 includes a generally tubular housing 18 that defines an annular inlet 20. The housing 18 further encloses a booster compressor 22 for increasing the pressure of the air entering the turbine 14.

[0030] A high-pressure, multi-stage, axial-flow compressor 24 receives the pressurized air from the booster compressor 22 and further increases the pressure of the air. The pressurized air flows into a combustor 26 where fuel is injected into the pressurized air stream to increase the temperature and energy level of the pressurized air. The high-energy combustion products flow from the combustor 26 to a first high-pressure (HP) turbine 28 for driving the high-pressure compressor 24 via a first HP drive shaft, and then to a second low-pressure (LP) turbine 32 for driving the booster compressor 22 and the fan section 16 via a second LP drive shaft coaxial with the first HP drive shaft. The HP turbine 28 includes HP stationary nozzles 34. The LP turbine 32 includes LP stationary nozzles 35.

[0031] A rotor disk is located downstream of the nozzles, which rotates about the centerline axis 12 of the gas turbine engine 10 and carries an array of airfoil-shaped turbine blades 36. Shrouds 29, 38 having a plurality of arcuate shroud segments are arranged to surround and closely encircle the turbine blades 27, 36, thereby defining an outer flow path boundary along the radial direction R of the gas turbine engine 10 for the hot gas flow passing over the turbine blades 27, 36. After driving each turbine 28, 32, the combustion products exit the turbine 14 through an exhaust nozzle 40.

[0032] The fan section 16 includes a rotatable axial - flow fan rotor 30 and a plurality of fan rotor blades 46, and the fan rotor blades 46 are surrounded by an annular fan housing 42. It will be appreciated that the annular fan housing 42 is supported from the turbine 14 by a plurality of outlet guide vanes 44 that extend generally radially and are circumferentially spaced apart. In this way, the fan housing 42 encloses the fan rotor 30 and the plurality of fan rotor blades 46.

[0033] From a flow perspective, it will be appreciated that an initial air flow 50 (represented by arrows) enters the gas turbine engine 10 through an inlet 52. The air flow 50 passes through the fan rotor blades 46 and divides into a first compressed air flow 54 (represented by arrows) and a second compressed air flow 56 (represented by arrows). The first compressed air flow 54 moves through the fan housing 42, and the second compressed air flow 56 enters the booster compressor 22. The pressure of the second compressed air flow 56 increases and enters the HP compressor 24, as shown by arrow 58. After being mixed with fuel and burned in the combustor 26, the combustion products 48 leave the combustor 26 and flow through the HP turbine 28. Then, the combustion products 48 flow through the LP turbine 32 and escape via the exhaust nozzle 40, thereby providing thrust for the gas turbine engine 10.

[0034] However, it should be understood that Figure 1 the exemplary gas turbine engine 10 shown is for example only, and in other exemplary embodiments, the gas turbine engine 10 can have any other suitable configuration. For example, although the shown gas turbine engine 10 is configured as a ducted gas turbine engine (i.e., includes a fan housing 42), in other embodiments, the gas turbine engine 10 can be a non - ducted gas turbine engine (such that the fan is a non - ducted fan and the outlet guide vanes 44 project cantilevered from the outer casing 18). Additionally or alternatively, although the shown gas turbine engine 10 is configured as a direct - drive gas turbine engine (the fan is directly driven by the LP turbine 32) and a fixed - pitch gas turbine engine (the fan rotor blades 46 cannot rotate about their respective pitch axes), in other embodiments, the gas turbine engine 10 can additionally or alternatively be configured as a geared gas turbine engine (including a reduction gearbox between the LP turbine 32 and the fan), a variable - pitch gas turbine engine (such that the fan includes fan rotor blades 46 that can rotate about the pitch axes) or both. It should also be understood that in other exemplary embodiments, aspects of the present disclosure can be incorporated into any other suitable gas turbine engine. For example, in other exemplary embodiments, aspects of the present disclosure can (as appropriate) be incorporated into, for example, a turboprop gas turbine engine, a turboshaft gas turbine engine or a turbojet gas turbine engine.

[0035] Reference Figure 2, which shows a cross-sectional view of the CMC component. The CMC component can be, as a non-limiting example, the nozzle 60, such as one of the HP fixed nozzles 34 or LP fixed nozzles 35 as described above. In other non-limiting examples, the CMC component can be a shroud, a bushing, or any other suitable component formed of CMC material.

[0036] The nozzle 60 includes an airfoil 62 and a band 64. Generally, the nozzle 60 includes a plurality of airfoils 62 and a plurality of bands 64, and for clarity, Figure 2 one of them is shown. The airfoil 62 defines a flow path F for air flowing through the nozzle 60. The band 64 secures the airfoil 62 to another structure of the gas turbine engine 10, providing load-bearing support for the airfoil 62. An opening 66 is defined between the band 64 and the airfoil 62, and this opening is configured to allow a cooling medium to flow through.

[0037] During operation, the nozzle 60 may wear, for example, due to debris impact, high-temperature exposure, etc. The portion of the nozzle 60 worn by debris is the "repair area" 68 of the nozzle 60, that is, the portion of the nozzle 60 designated to be repaired or otherwise joined to another part of the nozzle 60. The repair area 68 can be disposed near the interface where the layer forming the airfoil 62 intersects with the layer forming the band 64.

[0038] Now referring to Figures 3A - 5C , which shows the repair area 68 of the nozzle 60 and an exemplary joining layer 70. The joining layer 70 is a composite material layer designed to connect two or more parts of the nozzle 60. For example, the first part of the nozzle 60 can be the airfoil 62, and the second part of the nozzle 60 can be the band 64. When the joining layer 70 is used to repair a CMC component (such as the nozzle 60), the joining layer 70 can be referred to as a "repair layer". Examples of materials for the joining layer 70 include, but are not limited to, prepreg composite layers, such as those including woven carbon fibers, adhesive materials, and coated SiC fibers. Other suitable materials for the joining layer 70 include oxides containing aluminum, zirconium, titanium, magnesium, silicon, mullite, spinel, etc. or combinations thereof; carbides containing silicon, boron, titanium, or combinations thereof; nitrides containing silicon and boron. Other known fibers suitable for use in the joining layer 70 include Nextel, Nicalon, hi-Nicalon, Tyranno, and Sylramic fibers.

[0039] The layer forming the airfoil 62 has a first fiber direction 72, while the layer forming the tape 64 has a second fiber direction 74 that is different from the first fiber direction 72 of the airfoil 62. For purposes of illustration in these figures, the second fiber direction 74 of the layer of the tape 64 is defined as 0 degrees and extends along axis A. The first fiber direction 72 of the layer of the airfoil 62 in FIG. 3 is 90 degrees and extends along axis B perpendicular to axis A. That is, generally, the first fiber direction 72 of the layer of the airfoil 62 and the second fiber direction 74 of the layer of the tape 64 can differ by 90 degrees. When the fiber direction 72 of the layer of the airfoil 62 is different from the fiber direction 74 of the layer of the tape 64, stress in the airfoil 62 and the tape 64 is more easily dissipated. Additionally, in forms not shown in the figures, the airfoil 62 or the tape 64 can include adjacent layers having alternating fiber directions, such as 0 degrees then 90 degrees in a 0 / 90 layer stack.

[0040] A surface layer 78 can be disposed on the bonding layer 70 to provide a smooth or flush surface with the airfoil 62 or the tape 64 or both. Specifically, the fiber direction of the surface layer 78 can be parallel or perpendicular to the fiber direction of the composite layer 76 of the repair region 68. That is, the fiber direction of the surface layer 78 can be equal to the fiber directions 72, 74 of the layer of the airfoil 62 or the layer of the tape 64. In the example of FIG. 3, the fiber direction of the surface layer 78 is equal to the fiber direction 74 of the tape 64. Alternatively, the fiber direction of the surface layer 78 can be selected to meet a particular shape, geometry, or clearance with one or more adjacent components.

[0041] The bonding layer 70 is placed such that the fiber direction 80 of the bonding layer 70 is not in the same plane as the fiber direction of the composite layer 76 of the repair region 68. That is, the repair region 68 includes the composite layer of the airfoil 62 or the tape 64 or both, and the fiber direction 80 of the bonding layer 70 is not in the same plane as the fiber direction 72, 74 of one or both of the airfoil 62 and the tape 64. In Figures 3A - 3B the exemplary embodiment, the fiber direction of the composite layer 76 of the repair region 68 is equal to the fiber direction 74 of the tape 64, defined as 0 degrees. More specifically, the fiber direction 80 of the bonding layer 70 can differ from the fiber direction of the composite layer 76 by at least 40 degrees and at most 140 degrees.

[0042] As Figure 3A shown, the bonding layer 70 can be placed on one of the composite layers 76 of the repair region 68. Such a configuration can occur when one of the composite layers 76 is a molded or repair surface.

[0043] Optionally, as Figure 3BAs shown, the bonding layer 70 can be placed between the airfoil 62 and the tape 64, and the composite layer 76 is adjacent to the bonding layer 70. This configuration may occur when the nozzle 60 is machined after the bonding layer 70 is cured to the repair area 68.

[0044] Now referring Figure 4 , a layer stack of the bonding layer 70 is shown. To repair the nozzle 60, the bonding layer 70 is first placed on the composite layer 76 of the repair area 68. The bonding layer 70 can be folded or arranged into a layer stack before being placed on the repair area 68. A "layer stack" is a group of bonding layers 70 that are placed or folded on top of each other before being applied to the repair area 68. Placing the layer stack on the repair area 68 may be easier or faster compared to a single bonding layer 70. For example, the layer stack can be slid into an opening or gap within the repair area 68. Further, by way of example, the size or orientation of such a gap can be variable or include blind holes. The layer stack forms a ridge structure that absorbs mechanical loads from the airfoil 62 and the tape 64, improves interlayer interaction, and creates an additional retention mechanism to increase the strength of the nozzle 60. The bonding layers 70 of the layer stack can be tilted when shrinking to the repair area 68 to accommodate post-build machining and thickness variations during curing of the bonding layer 70. This tilting reduces defects (such as voids) that may occur during shrinkage.

[0045] Particularly referring Figure 5A , 5B and 5C, the fiber direction 80 of some bonding layers 70 can be different from the fiber direction 80 of other bonding layers 70. That is, the bonding layer 70 can include a first bonding layer 70a and a second bonding layer 70b placed on the first bonding layer 70a such that the first fiber direction 80a of the first bonding layer 70a is different from the second fiber direction 80b of the second bonding layer 70b, thereby forming a cross pattern. In one exemplary embodiment, the first bonding layer 70a can be a layer as Figure 5A shown, which has a fiber direction 80a of 45 degrees, while the second bonding layer 70n can be a layer as Figure 5B shown, which has a fiber direction 80b of -45 degrees. The second fiber direction 80b can differ from the first fiber direction 80a by at least 60 degrees and at most 120 degrees. For the sake of clarity of the present disclosure, Figures 5A - 5C the bonding layers shown as <�

[0046] are generally denoted by the number "70" and the fiber direction "80", and letters are used to indicate specific bonding layers 70 having a specific fiber direction 80.

[0046] As Figure 5CAs shown, the bonding layer 70 may further include a third bonding layer 70c disposed on the second bonding layer 70b, and a fourth bonding layer 70d disposed on the third bonding layer 70c. The third bonding layer 70c defines a third fiber direction 80c equal to the first fiber direction 80a, and the fourth bonding layer 70d defines a fourth fiber direction 80d equal to the second fiber direction 80b. Thus, adjacent bonding layers 70 may have different fiber directions 80, which further disperses the load and inhibits damage to the nozzle 60.

[0047] In addition to being out of the same plane as the fiber direction 72 of the airfoil 62, the fiber direction 80 of the bonding layer 70 may also be out of the same plane as the flow direction of the flow path F (see Figure 3) defined by the airfoil 62. The flow path F is perpendicular to the plane defined by the A-axis and the B-axis, in which the fiber directions 72, 74, 80 are defined, and the fiber direction 80 of the bonding layer 70 will be out of the same plane because a non-zero angle is defined between the fiber direction 80 and the flow direction of the flow path F perpendicular to the A-B plane.

[0048] When the fiber direction 80 of the bonding layer 70 is out of the same plane as the fiber directions 72, 74 of the airfoil 62 or the tape 64, delamination of the bonding layer 70 during heating is inhibited, thereby improving the fusion of the bonding layer 70 with the airfoil 62 and the tape 64. The out-of-plane fiber direction 80 also provides an auxiliary holding mechanism to help fix the bonding layer 70 to the repair area 68. In this way, the repaired nozzle 60 can resist further damage from debris during operation.

[0049] Once all the bonding layers 70 have been placed, the nozzle 60 is heated to fuse the bonding layers 70 to the repair area 68, thereby forming an integral CMC component. By fusing the bonding layer 70 to the repair area 68, the mechanical load received by the nozzle 60 can be smoothly transferred between the airfoil 62 and the tape 64. Specifically, the bonding layer 70 can form a ridge structure that absorbs at least part of the mechanical load, thereby inhibiting crack formation and forming an auxiliary holding mechanism.

[0050] The nozzle 60 can be heated using heating processes such as melt infiltration, chemical vapor infiltration, brazing, sintering, etc. For example, in the melt infiltration process, silicon or a silicon alloy is typically applied externally to a porous preform and melted, and the molten silicon or silicon alloy penetrates into the pores of the preform. A portion of the molten silicon reacts with the elemental carbon present in the porous preform, such as carbon black initially present as a precursor in the slurry, or any carbon char formed by the pyrolysis of an organic binder. The molten silicon and carbon black react to form additional silicon carbide that fills the pores, thereby producing the final shape of the nozzle 60.

[0051] As another example, in a chemical vapor infiltration process, the preform is heated in a vacuum or inert atmosphere to decompose the organic binder, forming ceramic carbon and a porous layer. Then, a gaseous source of silicon, such as silicon carbide, is provided from the outside as a chemical vapor to infiltrate the porous layer. During infiltration, the chemical vapor reacts with the inner surface of the porous layer, depositing silicon carbide therein.

[0052] Since the bonding layer <70> may shrink during the heating process, the fiber direction <80> may be different from the original fiber direction <80> when the bonding layer <70> is placed. The bonding layer <70> can be arranged with additional material to accommodate this shrinkage during heating, such that the fiber direction <80> remains substantially the same after the heating process. Specifically, the bonding layer <70> can be arranged such that when the bonding layer <70> shrinks, the bonding layer <70> shrinks in a specific direction to form the profile surface of the nozzle <60> pointing along the flow path <F>.

[0053] In addition to repairing CMC components, the bonding layer <70> can also be used to join parts of the CMC components together. To perform the joining, i.e., to replace or join an entire airfoil <62>, the new cured airfoil <62> can be joined to the already cured tape <64> at a specific joining area using the bonding layer <70> described above. The joining area can be strategically placed to allow the overall CMC component to function after the joining is completed.

[0054] Now referring Figure 6 , another CMC component <90> is shown. Specifically, the CMC component <90> is a nozzle having an airfoil <92> and a tape <94>, and the airfoil <92> defines a cavity <96>. In this exemplary embodiment, the repair area <98> extends from the airfoil <92> to the tape <94>, including the cavity <96>. Since the cavity <96> is defined inside the nozzle, the bonding layer <100> is inserted into the cavity <96> such that the bonding layer <100> extends into the cavity <96>. Then, when the nozzle is heated, the bonding layer <100> fuses with the airfoil <92>, filling the cavity <96> and forming a monolithic structure to form the nozzle. Since the cavity <96> may be hidden by parts of the airfoil <92> (i.e., the cavity <96> is a blind cavity), arranging the bonding layer <100> such that some of the bonding layer <100> is inserted into the cavity <96> can improve the repair or manufacture of the CMC component <90>.

[0055] In addition to repairing the CMC component 90, the bonding layer 100 can also provide additional structure to the airfoil 92 or the strap 94. More specifically, some of the bonding layers 100 in the layer stack can form a ridge structure that, when fused to the repair region 98, forms an extension 102 extending from the repair region 98. The extension 102 can absorb energy from loads and debris impacts, thereby acting as a stiffening rib. The fiber direction of the bonding layer 100 forming the extension 102 can be different from the fiber direction of the bonding layers 100 in the rest of the layer stack. As an example, the extension 102 can include bonding layers 100 having a fiber direction of 0 degrees, 90 degrees, or a combination thereof, while the other bonding layers 100 can have a fiber direction of 45 degrees, -45 degrees, or a combination thereof. That is, the bonding layer 100 and the extension 102 can form a -45 / 45 / 0 / 90 layer stack.

[0056] Now referring Figure 7 , a flowchart of a method 200 for forming a CMC component in accordance with an exemplary aspect of the present disclosure is provided. Figure 7 The method 200 of Figure 1 can be used to repair or manufacture one or more of the CMC components described above with reference to

[0057] to FIG. 5. Accordingly, it should be understood that the method 200 can generally be used to repair the CMC components of the gas turbine engine 10 as described above. However, in other exemplary aspects, the method 200 can additionally or alternatively be used to form any other suitable CMC component, including newly manufactured components.

[0057] As shown, the method 200 includes arranging (202) the bonding layers in a layer stack such that the fiber directions of adjacent bonding layers are different. More specifically, as Figure 5C shown, the bonding layers can be arranged such that the fiber direction alternates between two specified fiber directions, such as the 45-degree and -45-degree fiber directions described above. It will be appreciated that this step can be omitted when the bonding layers are not arranged in the layer stack.

[0058] The method 200 includes placing (204) the bonding layer on the repair region of the CMC component. Placing can include, as a non-limiting example, placing the bonding layer as a layer stack or individually on the repair region. The bonding layer is placed such that the fiber direction of the bonding layer is not in the same plane as the fiber direction of the composite layer of the repair region. As an example, the fiber direction of the bonding layer can be 45 degrees and -45 degrees, which can be out of plane with the layers of the airfoil having a 0-degree fiber direction or with the layers of the strap having a 90-degree fiber direction. In an exemplary embodiment, the bonding layer extends into a cavity formed in the CMC component. As a further non-limiting example, placing can include inserting the layer stack into a gap, hole, blind opening, or repair region. This further includes that the layer stack can be slid or bucketed into such a gap, hole, blind opening, or repair region.

[0059] Method 200 includes heating the CMC component and the bonding layer at (206) to fuse the bonding layer to the repair area. The heating can be a melt infiltration process, a chemical vapor infiltration process, a brazing process, or any other suitable heating process. During heating, the bonding layer fuses with the repair area and becomes a repaired portion of the CMC component. In an exemplary embodiment, the bonding layer forms an extension that becomes a reinforcing rib when fused to the repair area to increase the strength of the CMC component.

[0060] The disclosed method allows for the repair or remanufacture of CMC components such as nozzles, airfoils, and tapes in a manner that allows CMC infiltration into the repair area, CMC brazing in the repair area, or green state bonding of CMC, thereby allowing further repair operations. By having certain layers in the repair area have a different fiber orientation than the layers in the adjacent portion of the component, delamination can be reduced or inhibited, thereby increasing the mechanical strength of the CMC component.

[0061] The subject matter of the following items provides further aspects:

[0062] A method for joining a ceramic matrix composite (CMC) component, the method comprising: placing a plurality of bonding layers onto a repair area, wherein a fiber orientation of each of the plurality of bonding layers is not in the same plane as a fiber orientation of one or more composite layers of the repair area; and forming a monolithic component by bonding the plurality of bonding layers to the repair area.

[0063] The method according to any one of the preceding items, wherein the fiber orientation of the plurality of bonding layers differs from the fiber orientation of the one or more composite layers by at least 40 degrees and at most 140 degrees.

[0064] The method according to any one of the preceding items, wherein the plurality of bonding layers includes a first bonding layer and a second bonding layer placed on the first bonding layer, wherein a first fiber orientation of the first bonding layer is different from a second fiber orientation of the second bonding layer.

[0065] The method according to any one of the preceding items, wherein the plurality of bonding layers includes a third bonding layer placed on the second bonding layer, the third bonding layer defining a third fiber orientation equal to the first fiber orientation.

[0066] The method according to any one of the preceding items, wherein the plurality of bonding layers includes a fourth bonding layer placed on the third bonding layer, the fourth bonding layer defining a fourth fiber orientation equal to the second fiber orientation.

[0067] The method according to any one of the preceding clauses, wherein the second fiber direction differs from the first fiber direction by at least 60 degrees and at most 120 degrees.

[0068] The method according to any one of the preceding clauses, wherein the plurality of bonding layers form a ridge structure that forms an extension extending from the repair region when fused with the repair region.

[0069] The method according to any one of the preceding clauses, wherein heating the CMC component further comprises heating the CMC component by one of a melt infiltration process, a chemical vapor infiltration process, or a brazing process.

[0070] The method according to any one of the preceding clauses, further comprising: folding the plurality of bonding layers into a folded layer stack and placing the folded layer stack onto the repair region.

[0071] The method according to any one of the preceding clauses, wherein the repair region includes a cavity defined within the CMC component, and wherein the method further comprises: placing the plurality of bonding layers within the repair region such that the plurality of bonding layers extend into the cavity.

[0072] The method according to any one of the preceding clauses, wherein the bonding includes an infiltration process to form a bond between the plurality of bonding layers and the repair region.

[0073] The method according to any one of the preceding clauses, wherein the plurality of bonding layers bond a first portion to a second portion.

[0074] The method according to any one of the preceding clauses, wherein one or more composite layers of the repair region each have a fiber direction of 90 degrees, the first portion includes additional composite layers each having a fiber direction of 90 degrees, the second portion includes a plurality of composite layers each having a fiber direction of 0 degrees, and each of the plurality of bonding layers has a fiber direction different from 0 degrees or 90 degrees.

[0075] The method according to any one of the preceding clauses, wherein the one or more composite layers of the repair region each have a fiber direction of 90 degrees, the first portion includes additional composite layers having fiber directions alternating between 0 degrees and 90 degrees, and wherein the plurality of bonding layers have fiber directions alternating between -45 degrees and 45 degrees.

[0076] A method for joining ceramic matrix composite (CMC) components, comprising placing a plurality of joining layers onto a specific area such that the respective fiber direction of each of the plurality of joining layers is not in the same plane as the fiber direction of one or more composite layers of the specific area. Wherein the respective fiber direction of each of the plurality of joining layers is different from the respective fiber direction of each adjacent joining layer among the plurality of joining layers.

[0077] The method according to any one of the preceding clauses, wherein the respective fiber direction of each of the plurality of joining layers is at least 60 degrees and at most 120 degrees different from the respective fiber direction of each adjacent joining layer among the plurality of joining layers.

[0078] The method according to any one of the preceding clauses, wherein placing the plurality of joining layers onto the specific area further comprises placing a first joining layer in a first fiber direction and placing a second joining layer onto the first joining layer in a second fiber direction, the second fiber direction being at least 60 degrees and at most 120 degrees different from the first fiber direction.

[0079] The method according to any one of the preceding clauses, further comprising placing a third joining layer onto the second joining layer in the first fiber direction and then placing a fourth joining layer onto the third joining layer in the second fiber direction.

[0080] The method according to any one of the preceding clauses, wherein the fiber direction of at least one of the plurality of joining layers is at least 40 degrees and at most 140 degrees different from the fiber direction of the one or more composite layers of the specific area.

[0081] The method according to any one of the preceding clauses, further comprising heating the CMC component using one of a melt infiltration process, a chemical vapor infiltration process, or a brazing process.

[0082] A method for joining ceramic matrix composite (CMC) components, comprising placing a plurality of joining layers onto a specific area such that the respective fiber direction of each of the plurality of joining layers is not parallel to the fiber direction of one or more composite layers of the specific area, and forming an integral component by bonding the plurality of joining layers to the specific area of the CMC component.

[0083] The method according to any one of the preceding clauses, wherein the respective fiber direction of each of the plurality of joining layers is different from the respective fiber direction of each adjacent joining layer among the plurality of joining layers.

[0084] The method according to any one of the preceding clauses, wherein the specific area is a repair area, and the integral component formed by bonding is a repaired CMC component.

[0085] A method for joining ceramic matrix composite (CMC) components, the method comprising: placing a layer stack defined by a plurality of joining layers onto a specific area of a CMC component such that the respective fiber directions of each of the plurality of joining layers are not in the same plane as the fiber directions of one or more composite layers of the specific area of the CMC component.

[0086] The method according to any one of the preceding clauses, wherein the respective fiber direction of each of the plurality of joining layers is different from the respective fiber direction of each adjacent composite layer of the one or more composite layers of the specific area of the CMC component.

[0087] The method according to any one of the preceding clauses, further comprising forming a monolithic component by bonding the plurality of joining layers to the specific area of the CMC component.

[0088] The method according to any one of the preceding clauses, wherein the specific area is a repair area and the monolithic component formed by the bonding is a repaired CMC component.

[0089] The method according to any one of the preceding clauses, wherein the specific area of the CMC component includes a cavity and placing the layer stack includes inserting the layer stack into at least a portion of the cavity.

[0090] This written description uses examples to disclose the present disclosure, including the best mode, and also enables those skilled in the art to practice the present disclosure, including making and using any device or system and performing any combined method. The patentable scope of the present disclosure is defined by the claims and may include other examples that occur to those skilled in the art. If these other examples have structural elements that are not different from the literal language of the claims, or if they include equivalent structural elements that do not have a substantial difference from the literal language of the claims, then these other examples are intended to fall within the scope of the claims.

Claims

1. A method for joining ceramic matrix composite (CMC) components, characterized in that, The method includes: placing a plurality of joining layers onto a repair area, wherein a fiber direction of each of the plurality of joining layers is not in the same plane as a fiber direction of one or more composite layers of the repair area; and forming a monolithic component by bonding the plurality of joining layers to the repair area.

2. The method according to claim 1, wherein Wherein the fiber direction of the plurality of joining layers is at least 40 degrees and at most 140 degrees different from the fiber direction of the one or more composite layers of the repair area.

3. The method according to claim 1, wherein Wherein the plurality of joining layers includes a first joining layer and a second joining layer placed on the first joining layer, wherein a first fiber direction of the first joining layer is different from a second fiber direction of the second joining layer.

4. The method according to claim 3, characterized in that, Wherein the plurality of joining layers includes a third joining layer placed on the second joining layer, and the third joining layer defines a third fiber direction equal to the first fiber direction.

5. The method according to claim 4, characterized in that, Wherein the plurality of joining layers includes a fourth joining layer placed on the third joining layer, and the fourth joining layer defines a fourth fiber direction equal to the second fiber direction.

6. The method according to claim 3, wherein Wherein the second fiber direction is at least 60 degrees and at most 120 degrees different from the first fiber direction.

7. The method according to claim 1, wherein Wherein the plurality of joining layers forms a ridge structure that forms an extension extending from the repair area when fused with the one or more composite layers of the repair area.

8. The method according to claim 1, characterized in that, Wherein heating the CMC component further includes heating the CMC component by one of a melt infiltration process, a chemical vapor infiltration process, or a brazing process.

9. The method according to claim 1, wherein Further includes: folding the plurality of joining layers into a folded layer stack; and inserting the folded layer stack into the repair area.

10. The method according to claim 1, characterized in that, Wherein the repair area includes a cavity defined within the CMC component, and wherein the method further includes: placing the plurality of joining layers in the repair area such that the plurality of joining layers extend into the cavity.

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