Method for repairing composite parts using support members

Abstract

A method for repairing a composite component includes positioning one or more repair fiber layers relative to a repair interface of the composite component such that a butt joint is formed at the repair interface between the one or more repair fiber layers and the composite component. In addition, the method includes forming a cavity within the composite component such that the cavity extends through the repair interface. In addition, the method includes installing a support member within the cavity. In this regard, the support member extends across the butt joint such that a first portion of the support member is positioned within the composite component and a second portion of the support member is positioned within the one or more repair fiber layers.

Description

Technical Field

[0001] The present disclosure relates generally to methods for repairing components and, more particularly, to methods for securing new composite materials to worn or damaged existing composite components. Background Art

[0002] In recent years, the use of non-traditional high-temperature materials, such as ceramic matrix composite (CMC) materials, in gas turbine engines has grown dramatically. In particular, there is a strong interest in replacing metal alloy components in the combustion and turbine sections of gas turbine engines with CMC components. CMC materials can withstand higher operating temperatures than metal alloys. In turn, higher operating temperatures increase the efficiency of gas turbine engines. In addition, CMC components require less cooling than metal components. In addition, CMC materials are lighter than metal components and can reduce the structural requirements of the engine.

[0003] However, gas turbine components formed from CMC materials can be very expensive. In this regard, when a CMC gas turbine component is worn or damaged, it may be necessary to repair rather than replace the component. Therefore, methods of repairing CMC components have been developed. Specifically, the worn or damaged portion of the CMC component can be removed and replaced with new CMC material. In this method, the new CMC material is typically joined to the existing component via a scarf joint. However, when repairing the edges of certain components (e.g., blades), there is not enough material (e.g., the edge is too thin) to form a scarf joint.

[0004] Therefore, an improved method of repairing composite parts would be welcome. Summary of the invention

[0005] 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.

[0006] In one aspect, the present subject matter relates to a method for repairing a composite component. The method includes positioning one or more repair fiber layers relative to a repair interface of the composite component such that a butt joint is formed at the repair interface between the one or more repair fiber layers and the composite component. In addition, the method includes forming a cavity within the composite component such that the cavity extends through the repair interface. In addition, the method includes installing a support member within the cavity. In this regard, the support member extends across the butt joint such that a first portion of the support member is positioned within the composite component and a second portion of the support member is positioned within the one or more repair fiber layers.

[0007] On the other hand, the present subject matter relates to a method for repairing a composite turbine component. The method includes preparing a repair interface of the composite turbine component by removing worn or damaged composite material. In addition, the method includes positioning one or more repair fiber layers at the repair interface in a transverse orientation relative to the original fiber layers forming the composite turbine component, so that a butt joint is formed at the repair interface between the one or more repair fiber layers and the composite turbine component. In addition, the method includes forming a cavity in the composite turbine component so that the cavity extends through the repair interface. In addition, the method includes installing a support member in the cavity. In this regard, the support member extends across the butt joint so that a first portion of the support member is positioned within the composite turbine component and a second portion of the support member is positioned within the one or more repair fiber layers.

[0008] 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 present invention and, together with the description, serve to explain the principles of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0009] This specification sets forth a complete and enabling disclosure of the present invention, including the best mode thereof, to one of ordinary skill in the art with reference to the accompanying drawings, in which:

[0010] Figure 1 is a schematic cross-sectional view of one embodiment of a gas turbine engine;

[0011] Figure 2 is a side view of one embodiment of a shroud block for a gas turbine engine;

[0012] Figure 3 is a flow chart of one embodiment of a method for repairing a composite component;

[0013] Figure 4 is a top view of an embodiment of a composite component, specifically showing the component before repair;

[0014] Figure 5 yes Figure 4 A top view of the composite component shown in , specifically showing a repair interface of the component after preparation for repair;

[0015] Figure 6 yes Figure 4 and 5 A top view of the composite component shown in , specifically showing the repair fiber layer positioned relative to the repair interface of the component to form a butt joint;

[0016] Figure 7 yes Figure 4-6 The composite part shown in Figure 6A cross-sectional view taken along line 7-7 in FIG. 1 , specifically showing the repair fiber layer oriented transversely relative to the fiber layer of the component;

[0017] Figure 8 yes Figure 4-7 , specifically showing holes formed in the repair fiber layer and the composite component; and

[0018] Fig. 9 yes Figure 4-7 A top view of the composite component shown in , specifically showing the grooves formed in the repair fiber layer and the composite component;

[0019] Fig.10 yes Figure 4-8 A top view of the composite component shown in , specifically showing a pin positioned within one of the holes formed in the repair fiber layer and the composite component;

[0020] Fig.11 yes Figure 4-7 and a top view of the composite component shown in 9, specifically showing the key or plate positioned within the groove formed in the repair fiber layer and the composite component;

[0021] Fig.12 is a flow chart of another embodiment of a method for repairing a composite component;

[0022] Fig.13 yes Figure 4 and 5 A top view of the composite component shown in , specifically showing a hole formed in the composite component;

[0023] Fig.14 yes Figure 4 and 5 A top view of the composite component shown in , specifically showing a groove formed in the composite component;

[0024] Fig.15 yes Figure 4 , 5 and a top view of the composite component shown in 13, specifically showing the pin installed in the formed hole;

[0025] Fig.16 yes Figure 4 , 5 and a top view of the composite component shown in 14, specifically showing the plate installed in the formed hole;

[0026] Fig.17 yes Figure 4 , 5 , 13 and 15 are top views of the composite components, specifically showing the repair fiber layers positioned relative to the repair interface of the components so as to form a butt joint; and

[0027] Fig.18 yes Figure 4 , 5 , 14 and 16 are top views of the composite components, specifically showing the repair fiber layers positioned relative to the repair interface of the components so as to form a butt joint.

[0028] Repeat use of reference characters in the present specification and drawings is intended to represent same or analogous features or elements of the invention. DETAILED DESCRIPTION

[0029] Reference will now be made in detail to exemplary embodiments of the subject matter of the present disclosure, one or more examples of which are shown in the accompanying drawings. Each example is provided by way of illustration and should not be construed as limiting the present disclosure. In fact, it will be apparent to those skilled in the art that various modifications and variations may be made in the present disclosure without departing from the scope or spirit of the present disclosure. For example, a feature shown or described as part of one embodiment may be used together with another embodiment to produce yet another embodiment. Therefore, the present disclosure is intended to encompass these modifications and variations within the scope of the appended claims and their equivalents.

[0030] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the position or importance of a single component.

[0031] In addition, the terms "upstream" and "downstream" refer to relative directions relative 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 toward which the fluid flows.

[0032] In addition, unless otherwise specified, the terms "low", "high" or their respective comparatives (e.g., lower, higher, as applicable) refer to relative speeds within the engine. For example, a "low-pressure turbine" operates at a pressure that is generally lower than a "high-pressure turbine". Alternatively, unless otherwise specified, the above terms may be understood as their highest level. For example, a "low-pressure turbine" may refer to the lowest maximum pressure turbine within a turbine section, while a "high-pressure turbine" may refer to the highest maximum pressure turbine within a turbine section.

[0033] In general, the present subject matter relates to methods for repairing composite components. In several embodiments, the disclosed methods can be used to repair various turbine components, such as ceramic matrix composite (CMC) gas turbine engine blades, buckets, shroud blocks, bushings, and / or the like. More specifically, during operation, composite components may wear or become damaged. For example, in some cases, the damaged portion of the composite component may even break off or become separated from the remainder of the composite component. Therefore, according to the disclosed methods, a new composite material is formed to replace the worn / damaged portion of the composite component.

[0034] The new composite material is joined to the composite component via a butt joint and is supported relative to the composite component via one or more support members (e.g., pins or plates). Specifically, in one embodiment, the repair fiber layer is positioned relative to the repair interface of the composite component so that a butt joint is formed at the repair interface between the repair fiber layer and the component. Thereafter, one or more cavities are formed in the repair fiber layer and the component so that the cavity extends through the repair interface. The support member is then installed in the cavity. In another embodiment, one or more cavities are formed in the composite component so that the cavity extends through the repair interface. The support member is then installed in each cavity. Thereafter, the repair fiber layer is pressed against the support member, thereby positioning the layer relative to the repair interface to form a butt joint. Regardless of whether the positioning of the repair fiber layer occurs before or after the installation of the support member, the result is that each support member extends across the butt joint so that the first portion of each support member is positioned in the composite component and the second portion of each support member is positioned in the repair fiber layer.

[0035] The use of a support member and a butt joint provides one or more technical advantages. More specifically, as described above, the edges of certain composite components (e.g., turbine blades) do not have sufficient material to form a miter joint with a repair fiber layer. In this regard, the disclosed method forms a butt joint between a fiber layer and a repair fiber layer of a composite component that requires less material than a miter joint. However, unsupported butt joints are weak and generally cannot meet the operational requirements of the component. In this regard, a support member (such as a pin or plate) supports the butt joint between the fiber layer of the component and the repair fiber layer. Thus, the support member reinforces the butt joint so that the butt joint is strong enough to support the operational loads of the component.

[0036] Now referring to the accompanying drawings, Figure 1 is a schematic cross-sectional view of one embodiment of a gas turbine engine 10. In the illustrated embodiment, the engine 10 is configured as a high bypass turbofan engine. However, in alternative embodiments, the engine 10 may be configured as a propfan engine, a turbojet engine, a turboprop engine, a turboshaft gas turbine engine, or any other suitable type of gas turbine engine.

[0037] like Figure 1 As shown, the engine 10 defines a longitudinal direction L, a radial direction R, and a circumferential direction C. Generally, the longitudinal direction L extends parallel to an axial centerline 12 of the engine 10 , the radial direction R extends orthogonally outward from the axial centerline 12 , and the circumferential direction C extends generally concentrically about the axial centerline 12 .

[0038] In general, the engine 10 includes a fan 14, a low pressure (LP) spool 16, and a high pressure (HP) spool 18, at least partially surrounded by an annular nacelle 20. More specifically, the fan 14 may include a fan rotor 22 and a plurality of fan blades 24 (one shown) coupled to the fan rotor 22. In this regard, the fan blades 24 are spaced apart from one another along a circumferential direction C and extend outwardly from the fan rotor 22 along a radial direction R. In addition, the LP and HP spools 16, 18 are positioned downstream of the fan 14 along the axial centerline 12 (i.e., in the longitudinal direction L). As shown, the LP spool 16 is rotatably coupled to the fan rotor 22, thereby allowing the LP spool 16 to rotate the fan 14. In addition, a plurality of outlet guide vanes or struts 26 spaced apart from one another in the circumferential direction C extend between an outer casing 28 surrounding the LP and HP spools 16, 18 and the nacelle 20 along a radial direction R. Thus, the struts 26 support the nacelle 20 relative to the outer shell 28 such that the outer shell 28 and the nacelle 20 define a bypass airflow passage 30 therebetween.

[0039] The outer casing 28 generally surrounds or encloses the compressor section 32, the combustion section 34, the turbine section 36, and the exhaust section 38 in a serial flow order. For example, in some embodiments, the compressor section 32 may include a low pressure (LP) compressor 40 of the LP spool 16 and a high pressure (HP) compressor 42 of the HP spool 18, with the HP compressor 42 positioned downstream of the LP compressor 40 along the axial centerline 12. Each compressor 40, 42 may in turn include one or more rows of stator vanes 44 interleaved with one or more rows of compressor rotor blades 46. Further, in some embodiments, the turbine section 36 includes a high pressure (HP) turbine 48 of the HP spool 18 and a low pressure (LP) turbine 50 of the LP spool 16, with the LP turbine 50 positioned downstream of the HP turbine 48 along the axial centerline 12. Each turbine 48, 50 may in turn include one or more rows of stator vanes 52 interleaved with one or more rows of turbine rotor blades 54.

[0040] In addition, the LP spool 16 includes a low pressure (LP) shaft 56, and the HP spool 18 includes a high pressure (HP) shaft 58 positioned concentrically about the LP shaft 56. In such an embodiment, the HP shaft 58 rotatably couples the rotor blades 54 of the HP turbine 48 and the rotor blades 46 of the HP compressor 42, such that rotation of the HP turbine rotor blades 54 rotatably drives the HP compressor rotor blades 46. As shown, the LP shaft 56 is directly coupled to the rotor blades 54 of the LP turbine 50 and the rotor blades 46 of the LP compressor 40. In addition, the LP shaft 56 is coupled to the fan 14 via a gearbox 60. In this regard, rotation of the LP turbine rotor blades 54 rotatably drives the LP compressor rotor blades 46 and the fan blades 24.

[0041] In several embodiments, the engine 10 may generate thrust to propel an aircraft. More specifically, during operation, air (indicated by arrow 62) enters an inlet portion 64 of the engine 10. The fan 14 supplies a first portion of the air 62 (indicated by arrow 66) to the bypass airflow passage 30 and a second portion of the air 62 (indicated by arrow 68) to the compressor section 32. The second portion 68 of the air 62 first flows through the LP compressor 40, where the rotor blades 46 gradually compress the second portion 68 of the air 62 in the LP compressor 40. Next, the second portion 68 of the air 62 flows through the HP compressor 42, where the rotor blades 46 continue to gradually compress the second portion 68 of the air 62 in the HP compressor 42. The compressed second portion 68 of the air 62 is then delivered to the combustion section 34. In the combustion section 34, the second portion 68 of the air 62 is mixed with fuel and combusted to produce high temperature and high pressure combustion gases 70. Thereafter, the combustion gases 70 flow through the HP turbine 48, where the HP turbine rotor blades 54 extract a first portion of kinetic energy and / or thermal energy therefrom. This energy extraction rotates the HP shaft 58, thereby driving the HP compressor 42. The combustion gases 70 then flow through the LP turbine 50, wherein the LP turbine rotor blades 54 extract a second portion of the kinetic and / or thermal energy therefrom. This energy extraction rotates the LP shaft 56, thereby driving the LP compressor 40 and the fan 14 via the gearbox 60. The combustion gases 70 then exit the engine 10 through the exhaust section 38.

[0042] Figure 2 is a side view of an embodiment of a shroud block 72 of a gas turbine engine 10. Typically, several shroud blocks 72 are circumferentially arranged to form a shroud (not shown) that surrounds or otherwise encloses a row of rotor blades 46 in the compressor section 32 or a row of rotor blades 54 in the turbine section 36. As shown, the shroud block 72 includes an annular wall 74 extending in a radial direction R between an inner surface 76 and an outer surface 78. The inner surface 76 is in turn positioned proximate to the tip of the corresponding blade 46, 54 to minimize leakage of air / combustion gases 68 / 70 through the blade 46, 54. In addition, the shroud 72 includes a pair of mounting rails 80 (one is shown). The rails 80 are spaced apart from each other in the longitudinal direction L and extend outwardly from the outer surface 78 of the annular wall 74 in the radial direction R. In addition, each rail 80 defines a pair of mounting holes 82 for coupling the shroud block 72 to the outer casing 28 of the engine 10. However, in alternative embodiments, the shroud block 72 may have any other suitable configuration.

[0043] In addition, one or more components of the gas turbine engine 10 may be formed of a composite material, such as a ceramic matrix composite (CMC) material or a polymer matrix composite (PMC) material. For example, in several embodiments, the compressor bucket 44, the compressor blade 46, the turbine bucket 52, the turbine blade 54, and the shroud block 72 may be formed of a CMC material, while the fan blade 24 may be formed of a PMC material. However, in alternative embodiments, any other suitable component of the engine 10 may be formed of a composite material.

[0044] Above and Figure 1 and 2 The configuration of the gas turbine engine 10 shown in the figure is provided only to place the present subject matter in an exemplary field of use. Therefore, the present subject matter can be easily adapted to any manner of gas turbine engine configuration, including other types of aviation-based gas turbine engines, marine-based gas turbine engines, and / or land-based / industrial-based gas turbine engines.

[0045] Figure 3 is a flow chart of one embodiment of a method 100 for repairing a composite component. Figure 3 The steps performed in a specific order are depicted, but the disclosed method is not limited to any specific order or arrangement. Therefore, the various steps of the disclosed method can be omitted, rearranged, combined and / or adapted in various ways without departing from the scope of the present disclosure.

[0046] In general, the steps of method 100 will be described below in the context of repairing a composite component 200. Figure 4-11 For example, as will be described below, composite component 200 may correspond to a composite component of gas turbine engine 10. However, in alternative embodiments, composite component 200 may correspond to any other suitable composite component.

[0047] Figure 4 200. As shown, the composite component 200 includes a repair area 202. Generally, the repair area 202 corresponds to a portion of the component 200 that will be repaired according to the methods disclosed herein. For example, in some cases, the damaged / worn portion of the component 200 may be disconnected or separated from the remainder of the component 200. In this case, at least a portion of the repair area 202 may be free of composite material prior to repair. In addition, as will be described below, it may be necessary to remove additional material from the repair area 200, such as when the repair area 202 has an irregular surface 204, in order to facilitate the repair. Therefore, the repair area 202 is a portion of the component 200 that will be replaced with new composite material. Although in Figure 1The component 200 shown in FIG. 1 includes only one repair region 202 , but in other embodiments, the component 200 may include any other suitable number of repair regions 202 .

[0048] In addition, the composite component 200 can be formed by any suitable composite material. For example, the composite material can be selected from, but not limited to, a group consisting of ceramic matrix composite (CMC), polymer matrix composite (PMC), metal matrix composite (MMC) or a combination thereof. Suitable examples of matrix materials for CMC include, but are not limited to, silicon carbide, aluminum oxide, silicon oxide and combinations thereof. Suitable examples of matrix materials for PMC include, but are not limited to, epoxy-based matrix, polyester-based matrix and combinations thereof. Suitable examples of matrix materials for MMC include, but are not limited to, aluminum, titanium and combinations thereof. For example, MMC can be formed by powdered metal (such as, but not limited to, aluminum powder or titanium powder), which can be melted into a continuous molten liquid metal that can encapsulate the fibers present in the component and then cool into a solid ingot with embedded fibers. The resulting MMC is a metal product with increased rigidity, and the metal part (matrix) is the main load-bearing element.

[0049] Reference again Figure 3 At (102), method 100 includes preparing a repair interface of a composite component for repair. Specifically, in several embodiments, at (102), worn or damaged material within the repair region 202 that has not been disconnected or separated from the remainder of the composite component 200 can be removed from the component 200 by machining, grinding, cutting, etc. For example, in some cases, such as Figure 4 As shown, the repair area 202 may include an irregular surface 204 that is not suitable for bonding to the new composite material. In this case, at (102), material from the component 200 may be removed to form a suitable repair interface 206. The repair interface 206 is in turn the surface of the component 200 to which the new composite material will be bonded. Figure 5 As shown, upon completion (102), the worn / damaged material of the repair area 202 has been removed and a repair interface 206 has been formed. In one embodiment, the repair interface 206 forms a regular surface, ie, a smooth surface or a uniform surface, such as a planar surface.

[0050] In addition, if Figure 3 As shown, after preparing the repair interface, at (104), method 100 includes positioning one or more repair fiber layers relative to the repair interface of the composite component so that a butt joint is formed at the repair interface between the one or more repair fiber layers and the composite component. Figure 6 and 7As shown, the repair fiber layer 208 is positioned relative to the component 200 so that a butt joint 210 is formed at the repair interface 206 between the repair fiber layer 208 and the component 200. In this regard, the repair fiber layer 208 contacts the repair interface 206 and occupies the space where the worn / damaged material originally existed. As described below, a matrix precursor material is provided (e.g., via infiltration or injection) so that a new composite material is formed in the repair area 202, thereby repairing the component 200. In this regard, the repair fiber layer 208 corresponds to the fiber material that forms the composite material of the component 200.

[0051] like Figure 7 Specifically shown, at (104), in several embodiments, the repair fiber layer 208 is transversely oriented relative to the fiber layer 212 of the composite component 200. This transverse orientation strengthens the butt joint 206, thereby increasing the operational loads that the repaired component 200 can withstand. However, in alternative embodiments, the repair fiber layer 208 can be oriented relative to the fiber layer 212 of the composite component 200 in any other suitable manner.

[0052] In addition, if Figure 3 As shown, at (106), after positioning the repair layer, method 100 includes forming a cavity in the composite component such that the cavity extends through the repair interface. Figure 8 As shown, in one embodiment, after positioning the repair fiber layer 208, a pair of holes 214 are formed (e.g., via drilling) in the repair fiber layer 208 and the composite component 200 such that the holes 214 extend through the repair interface 206. In this regard, a first portion of each hole 214 is positioned within the composite component 200, and a second portion of each hole 214 is positioned within the repair area 202. In another embodiment, as shown in FIG. Fig. 9 As shown, after positioning the repair fiber layer 208, a groove 215 is formed (e.g., via milling) in the repair fiber layer 208 and the composite component 200 such that the groove 215 extends through the repair interface 206. In this regard, a first portion of the groove 215 is positioned within the composite component 200 and a second portion of the groove 215 is positioned within the repair area 202. However, in alternative embodiments, any other suitable number and / or type of cavities may be formed at (106).

[0053] In addition, if Figure 3 As shown, at (108), method 100 includes installing a support member in the cavity. Fig.10 As shown, in one embodiment, a pin 216 is installed in each hole 214 such that a first portion 218 of each pin 216 is positioned within the composite component 200 and a second portion 220 of each pin 216 is positioned within the repair fiber layer 208. Fig.11As shown, in another embodiment, a key or plate 217 is installed in the slot 215 so that a first portion 219 of the plate 217 is positioned within the composite component 200 and a second portion 221 of the plate 217 is positioned within the repair fiber layer 208. As will be described below, a support member (e.g., a pin 216 or a plate 217) supports the repair fiber layer 208 relative to the component 200, thereby strengthening the butt joint 210. The support member can be formed of any suitable material, such as a metallic material or a composite material. However, in alternative embodiments, the support member can be configured as any other suitable device or component capable of supporting the repair fiber layer 208 and the component 200 relative to each other in a butt joint connection.

[0054] In addition, if Figure 3 As shown, after the support member is installed, at (108), the method 100 includes providing a matrix precursor to one or more repair fiber layers. More specifically, after the repair fiber layer 208 has been positioned and the pin 216 has been installed, the matrix precursor is supplied or otherwise provided to the repair fiber layer 208. For example, in one embodiment, the composite part 200 is formed of a ceramic matrix composite (CMC) material. In such an embodiment, at (108), the repair fiber layer 208 is infiltrated (e.g., by melt infiltration) using a suitable infiltrant (e.g., silicon). In another embodiment, the composite part 200 is formed of a polymer matrix composite (PMC) material. In such an embodiment, (at 108), the repair fiber layer 208 is injected (e.g., via vacuum infiltration) using a suitable resin. However, in alternative embodiments, the matrix precursor can be provided to the repair fiber layer 208 in any other suitable manner.

[0055] In addition, at (112), after providing the matrix material, the method 100 includes heat treating one or more repair fiber layers. For example, after infiltrating or injecting the repair fiber layer 208, the layer 208 is heat treated (e.g., in a suitable furnace). After the heat treatment at (112), the support member (e.g., pin 216 or plate 217) remains in the composite component 200. In this regard, the support member supports the butt joint 210 between the fiber layer 212 and the repair fiber layer 208 of the component 200. In this way, the support member strengthens the butt joint 210 so that the butt joint 210 is strong enough to support the operating loads of the component 200.

[0056] Fig.12 FIG. 3 is a flow chart of another embodiment of a method 300 for repairing a composite component. As will be described below, Fig.12 The steps of the embodiment of the method 300 shown are similar to Figure 3 The steps of the embodiment of the method 100 are shown. However, Fig.12 Some steps of the embodiment of the method 300 are shown in Figure 3The embodiments of the method 100 shown are performed in different orders. Generally speaking, the steps of the method 300 will be performed in the following order: Figure 13-18 The following description will now be made of a repair composite component 200 schematically depicted in FIG. 1 , wherein like reference numerals are used to refer to reference numerals in the drawings. Figure 4-11 Similar features are depicted in component 200 .

[0057] like Fig.12 As shown, at ( 302 ), method 300 may include preparing a repair interface in repair region 202 of composite component 200 for repair, as described above with reference to ( 102 ).

[0058] In addition, if Fig.12 As shown, after preparing the repair interface, at (304), method 300 includes forming a cavity in the composite component such that the cavity extends through the repair interface. Fig.13 As shown, in one embodiment, a pair of holes 214 are formed (e.g., via drilling) in the composite component 200 such that the holes 214 extend through the repair interface 206. Fig.14 As shown, in one embodiment, grooves 215 are formed (eg, via milling) in composite component 200 such that grooves 215 extend through repair interface 206. However, in alternative embodiments, any other suitable number and / or type of cavities may be formed at (304).

[0059] In addition, if Fig.12 As shown, at (306), method 300 includes installing a support member in the cavity. Fig.15 As shown, in one embodiment, the pins 216 are installed in the holes 214 such that the first portion 218 of each pin 216 is positioned within the composite component 200 and the second portion 220 of each pin 216 extends outwardly from the repair interface 206. Fig.16 As shown, in another embodiment, a key or plate 217 is installed in the slot 215 so that a first portion 219 of the plate 217 is positioned within the composite component 200 and a second portion 221 of the plate 217 extends outwardly from the repair interface 206. However, in alternative embodiments, the support member can be configured as any other suitable device or member capable of supporting the repair fiber layer 208 and the component 200 relative to each other in a butt joint connection.

[0060] In addition, if Fig.12 As shown, after installing the pin, at (308), method 300 includes positioning one or more repair fiber layers relative to the repair interface of the composite component so that a butt joint is formed at the repair interface between the one or more repair fiber layers and the composite component. Fig.17As shown, in one embodiment, the repair fiber layer 208 is pressed onto the pin 216 so that the pin 216 pierces the repair fiber layer 208 and forms a butt joint 210 at the repair interface 206 of the component 200. Fig.18 As shown, in one embodiment, the repair fiber layer 208 is pressed against the plate / key 217 so that the plate / key 217 pierces the repair fiber layer 208 and forms a butt joint 210 at the repair interface 206 of the component 200. In this regard, the repair fiber layer 208 contacts the repair interface 206 and occupies the space where the worn / damaged material originally existed. Figure 7 As shown, in several embodiments, the repair fiber layer 208 is oriented transversely relative to the fiber layer 212 of the composite component 200 .

[0061] In addition, if Fig.12 As shown, at (310), after positioning the repair fiber layers, method 300 includes providing a matrix precursor to the one or more repair fiber layers as described above.

[0062] Additionally, at (312), after providing the matrix material, method 100 includes heat treating the one or more repair fiber layers as described above.

[0063] In various embodiments described herein, the support member (eg, pin 216 or plate 217) may extend completely through one or more repair fiber layers 208 to an exterior surface of the repair component, such as Fig.10 and 11 Optionally, the support member (eg, pin 216 or plate 217) may not extend completely through one or more repair fiber layers 208 so as to be encapsulated within the repaired portion of the composite component, such as Fig.17 and 18 For example, when the plurality of repair fiber layers 208 are pressed against the pin 216 or the plate 217, the pin 216 or the plate 217 may only pierce one or more of the plurality of repair fiber layers 208, but not all.

[0064] Figure 3 Method 100 and / or Fig.12 The method 300 may be used to repair a composite turbine component. Thus, in the above-described embodiments, the composite component 200 may be a composite component of the engine 10. For example, in some embodiments, the composite component 200 corresponds to the fan blade 24, the compressor blade 44, the compressor blade 46, the turbine blade 52, the turbine blade 54, and / or the shroud block 72 of the engine 10. However, in alternative embodiments, the composite component may correspond to any suitable component, such as other components of a turbine or any other turbine component.

[0065] This written description uses examples to disclose the invention, including the best mode, and also enables those skilled in the art to practice the invention, including making and using any device or system and performing any combined method. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. These 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 that do not differ substantially from the literal language of the claims.

[0066] Further aspects of the invention are provided by the subject matter of the following clauses:

[0067] A method for repairing a composite component, the method comprising: positioning one or more repair fiber layers relative to a repair interface of the composite component so that a butt joint is formed at the repair interface between the one or more repair fiber layers and the composite component; forming a cavity in the composite component so that the cavity extends through the repair interface; and installing a support member in the cavity, wherein the support member extends across the butt joint so that a first portion of the support member is positioned in the composite component and a second portion of the support member is positioned in the one or more repair fiber layers.

[0068] A method as recited in one or more of these clauses, wherein the support member comprises a pin and the cavity comprises a hole.

[0069] A method as recited in one or more of these clauses, wherein the support member comprises a plate and the cavity comprises a slot.

[0070] The method of one or more of these clauses, wherein positioning the one or more repair fiber layers comprises orienting the one or more repair fiber layers transversely relative to the plurality of original fiber layers forming the composite component.

[0071] The method according to one or more of these clauses, wherein forming a cavity in the composite component and installing a support member in the cavity is performed before positioning the one or more repair fiber layers relative to the repair interface.

[0072] The method of one or more of these clauses, wherein positioning the one or more repair fiber layers comprises pressing the one or more repair fiber layers against the support member such that the support member pierces the one or more repair fiber layers.

[0073] The method according to one or more of these clauses, wherein positioning one or more repair fiber layers relative to the repair interface is performed prior to forming the cavity.

[0074] The method of one or more of these clauses, wherein forming the cavity comprises forming the cavity within one or more of the repair fiber layers and the composite component such that the cavity extends across the repair interface.

[0075] The method according to one or more of these clauses, further comprising: providing a matrix precursor to the one or more repair fiber layers.

[0076] The method of one or more of these clauses, wherein: the composite component is formed from a ceramic matrix composite material; and providing the matrix precursor includes infiltrating the one or more layers of repair fibers with an infiltrant.

[0077] The method of one or more of these clauses, wherein: the composite component is formed from a polymer-based composite material; and providing the matrix precursor includes infusing one or more layers of repair fibers with a resin.

[0078] The method according to one or more of these clauses, further comprising: after providing the matrix precursor, heat treating the one or more repair fiber layers.

[0079] A method according to one or more of these clauses, wherein the support member remains in the composite component after the heat treatment.

[0080] A method as recited in one or more of these clauses, wherein the support member is formed of a metallic material or a composite material.

[0081] The method according to one or more of these clauses, wherein the composite component is a turbine component.

[0082] A method for repairing a composite turbine component, the method comprising: preparing a repair interface of the composite turbine component by removing worn or damaged composite material; positioning one or more repair fiber layers at the repair interface in a transverse orientation relative to original fiber layers forming the composite turbine component, so that a butt joint is formed at the repair interface between the one or more repair fiber layers and the composite turbine component; forming a cavity in the composite turbine component so that the cavity extends through the repair interface; and installing a support member in the cavity, wherein the support member extends across the butt joint so that a first portion of the support member is positioned in the composite turbine component and a second portion of the support member is positioned in the one or more repair fiber layers.

[0083] The method of one or more of these clauses, wherein forming a cavity within the composite turbine component and installing a support member within the cavity are performed prior to positioning the one or more repair fiber layers at the repair interface.

[0084] The method of one or more of these clauses, wherein positioning the one or more repair fiber layers comprises pressing the one or more repair fiber layers against the support member such that the support member pierces the one or more repair fiber layers.

[0085] A method according to one or more of these clauses, wherein the one or more repair fiber layers include a plurality of repair fiber layers, and wherein the support member pierces less than all of the plurality of repair fiber layers such that the support member is encapsulated within the repaired portion of the composite turbine component.

[0086] A method according to one or more of these clauses, wherein the composite turbine component is formed of a ceramic matrix composite material, the method further comprising: infiltrating one or more repair fiber layers with an infiltrant; and heat treating the one or more repair fiber layers to form a repaired composite turbine component.

[0087] The method of one or more of these clauses, wherein positioning the one or more repair fiber layers comprises pressing the one or more repair fiber layers against the support member such that the support member pierces the one or more repair fiber layers.

Claims

1. A method for repairing a composite component, characterized in that The method comprises: removing worn or damaged composite material from a composite component to form a repair interface of the composite component; positioning one or more repair fiber layers relative to the repair interface such that a butt joint is formed at the repair interface between the one or more repair fiber layers and the composite component; After removing the worn or damaged composite material, forming a blind hole or a blind slot in the composite component such that the blind hole or the blind slot extends through the repair interface and into the composite component; and Installing a supporting member in the blind hole or the blind groove, wherein the support member extends across the butt joint such that a first portion of the support member is positioned within the composite component and a second portion of the support member is positioned within the one or more repair fiber layers.

2. The method according to claim 1, characterized in that: in, The support member comprises a pin, and the blind hole or the blind slot comprises the blind hole.

3. The method according to claim 1, characterized in that in, The supporting member includes a plate, and the blind hole or the blind groove includes the blind groove.

4. The method according to claim 1, characterized in that in, Positioning the one or more repair fiber layers includes orienting the one or more repair fiber layers transversely relative to the plurality of original fiber layers forming the composite component.

5. The method according to claim 1, characterized in that in, Forming the blind hole or the blind slot in the composite component and installing the support member in the blind hole or the blind slot are performed before positioning the one or more repair fiber layers relative to the repair interface.

6. The method according to claim 5, characterized in that in, Positioning the one or more repair fiber layers includes pressing the one or more repair fiber layers onto the support member such that the support member penetrates the one or more repair fiber layers.

7. The method according to claim 1, characterized in that Further including: A matrix precursor is provided to the one or more repair fiber layers.

8. The method according to claim 7, characterized in that in: The composite component is formed of a ceramic matrix composite material; and Providing the matrix precursor includes infiltrating the one or more repair fiber layers with a penetrant.

9. The method according to claim 7, characterized in that: in: The composite component is formed of a polymer-based composite material; and Providing the matrix precursor includes infusing the one or more repair fiber layers with a resin.

10. The method according to claim 7, characterized in that Further including: After providing the matrix precursor, the one or more repair fiber layers are heat treated.

11. The method according to claim 10, characterized in that in, The support member remains in the composite component after heat treatment.

12. The method according to claim 1, characterized in that in, The supporting member is formed of a metal material or a composite material.

13. The method according to claim 1, characterized in that in, The composite component is a turbine component.

14. A method for repairing a composite turbine component, characterized in that: The method comprises: removing worn or damaged composite material from a composite turbine component to form a repaired interface of the composite turbine component; positioning one or more repair fiber layers at the repair interface in a transverse orientation relative to original fiber layers forming the composite turbine component such that a butt joint is formed at the repair interface between the one or more repair fiber layers and the composite turbine component; forming a blind hole or a blind slot in the composite turbine component after the worn or damaged composite material is removed such that the blind hole or the blind slot extends through the repair interface and into the composite turbine component; and Installing a supporting member in the blind hole or the blind groove, Wherein the support member extends across the butt joint such that a first portion of the support member is positioned within the composite turbine component and a second portion of the support member is positioned within the one or more repair fiber layers.

15. The method according to claim 14, characterized in that in, Forming the blind hole or the blind slot in the composite turbine component and installing the support member in the blind hole or the blind slot are performed prior to positioning the one or more repair fiber layers at the repair interface.

16. The method according to claim 15, characterized in that in, Positioning the one or more repair fiber layers includes pressing the one or more repair fiber layers onto the support member such that the support member penetrates the one or more repair fiber layers.

17. The method according to claim 16, characterized in that in, The one or more repair fiber layers include a plurality of repair fiber layers, and wherein the support member pierces less than all of the plurality of repair fiber layers such that the support member is encapsulated within the repaired portion of the composite turbine component.

18. The method according to claim 14, characterized in that in, The composite turbine component is formed of a ceramic matrix composite material, the method further comprising: infiltrating the one or more repair fiber layers with a penetrant; and The one or more repair fiber layers are heat treated to form a repaired composite turbine component.

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