Patent Description:
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. Specifically, there is strong interest in replacing metal alloy components within the combustion and turbine sections of a gas turbine engine with CMC components. CMC materials can withstand higher operating temperatures than metal alloys. Higher operating temperatures, in turn, increase the efficiency of the gas turbine engine. Moreover, CMC components require less cooling than metallic components. Additionally, CMC materials are lighter than metallic components and may reduce the structural demands on the engine.

However, gas turbine components formed from CMC materials can be quite expensive. In this respect, when a CMC gas turbine component becomes worn or damaged, it is desirable to repair, rather than replace, the component. As such, methods of repairing CMC components have been developed. For example, the worn or damaged portion(s) of a CMC component may be removed and replaced with new CMC material. While such methods work well, improvements are needed. <CIT> and <CIT> relate to a method of repairing a CMC composite article.

Accordingly, an improved method for repairing composite components would be welcomed in the technology.

In one aspect, the present subject matter is directed to a method for repairing a composite turbomachine component as set forth in claim <NUM>.

Reference now will be made in detail to exemplary embodiments of the presently disclosed subject matter, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation and should not be interpreted as limiting the present disclosure. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present disclosure without departing from the scope or spirit of the present disclosure. Thus, it is intended that the present disclosure covers such modifications and variations as come within the scope of the appended claims.

Furthermore, the terms "upstream" and "downstream" refer to the relative direction with respect to fluid flow in a fluid pathway.

Additionally, the terms "low," "high," or their respective comparative degrees (e.g., lower, higher, where applicable) each refer to relative speeds within an engine, unless otherwise specified. For example, a "low-pressure turbine" operates at a pressure generally lower than a "high-pressure turbine. " Alternatively, unless otherwise specified, the aforementioned terms may be understood in their superlative degree. For example, a "low-pressure turbine" may refer to the lowest maximum pressure turbine within a turbine section, and a "high-pressure turbine" may refer to the highest maximum pressure turbine within the turbine section.

In general, the present subject matter is directed to a method for repairing composite components. More specifically, when repairing a composite component, worn or damaged material may be removed (e.g., via machining, grinding, etc.) from a repair region of the component. Repair material (e.g., a fiber preform, a fiber tape, and/or the like) is then placed within the prepared repair region in place of the removed material. The repair material is infiltrated (e.g., via melt infiltration) to densify the repaired region of the component, thereby forming new composite material in place of the worn/damaged material. For example, the disclosed method may be used to repair various turbomachine components, such as ceramic matrix composite (CMC) gas turbine engine blades, vanes, shroud blocks, and/or the like.

The disclosed method includes filling one or more features defined by the composite component with a filler material prior to infiltration. More specifically, the composite component defines various features, such as holes, slots, and the like. Furthermore, in certain instances, during repair of the composite component, it may be desired to fill in one or more of the features with new material, effectively eliminating such features from the component. However, when the feature(s) are filled with only infiltrant, the new material present within the feature(s) is highly porous. Such porosity may weaken the component and limit its operating life. In this respect, filling the feature(s) with the filler material (e.g., in a slurry or powder form) before infiltration provides a composite precursor with which the infiltrant can bond. During infiltration, the infiltrant densifies the filler material (as opposed to simply filling the feature(s)), thereby forming new composite material having little to no porosity within the feature(s). As such, the filler material is a precursor to the composite material used to form the composite component, such as silicon or silicon carbide.

Referring now to the drawings, <FIG> is a schematic cross-sectional view of one embodiment of a gas turbine engine <NUM>. In the illustrated embodiment, the engine <NUM> is configured as a high-bypass turbofan engine. However, in alternative embodiments, the engine <NUM> 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.

As shown in <FIG>, the engine <NUM> defines a longitudinal direction L, a radial direction R, and a circumferential direction C. In general, the longitudinal direction L extends parallel to an axial centerline <NUM> of the engine <NUM>, the radial direction R extends orthogonally outward from the axial centerline <NUM>, and the circumferential direction C extends generally concentrically around the axial centerline <NUM>.

In general, the engine <NUM> includes a fan <NUM>, a low-pressure (LP) spool <NUM>, and a high pressure (HP) spool <NUM> at least partially encased by an annular nacelle <NUM>. More specifically, the fan <NUM> may include a fan rotor <NUM> and a plurality of fan blades <NUM> (one is shown) coupled to the fan rotor <NUM>. In this respect, the fan blades <NUM> are spaced apart from each other along the circumferential direction C and extend outward from the fan rotor <NUM> along the radial direction R. Moreover, the LP and HP spools <NUM>, <NUM> are positioned downstream from the fan <NUM> along the axial centerline <NUM> (i.e., in the longitudinal direction L). As shown, the LP spool <NUM> is rotatably coupled to the fan rotor <NUM>, thereby permitting the LP spool <NUM> to rotate the fan <NUM>. Additionally, a plurality of outlet guide vanes or struts <NUM> spaced apart from each other in the circumferential direction C extend between an outer casing <NUM> surrounding the LP and HP spools <NUM>, <NUM> and the nacelle <NUM> along the radial direction R. As such, the struts <NUM> support the nacelle <NUM> relative to the outer casing <NUM> such that the outer casing <NUM> and the nacelle <NUM> define a bypass airflow passage <NUM> positioned therebetween.

The outer casing <NUM> generally surrounds or encases, in serial flow order, a compressor section <NUM>, a combustion section <NUM>, a turbine section <NUM>, and an exhaust section <NUM>. For example, in some embodiments, the compressor section <NUM> may include a low-pressure (LP) compressor <NUM> of the LP spool <NUM> and a high-pressure (HP) compressor <NUM> of the HP spool <NUM> positioned downstream from the LP compressor <NUM> along the axial centerline <NUM>. Each compressor <NUM>, <NUM> may, in turn, include one or more rows of stator vanes <NUM> interdigitated with one or more rows of compressor rotor blades <NUM>. Moreover, in some embodiments, the turbine section <NUM> includes a high-pressure (HP) turbine <NUM> of the HP spool <NUM> and a low-pressure (LP) turbine <NUM> of the LP spool <NUM> positioned downstream from the HP turbine <NUM> along the axial centerline <NUM>. Each turbine <NUM>, <NUM> may, in turn, include one or more rows of stator vanes <NUM> interdigitated with one or more rows of turbine rotor blades <NUM>.

Additionally, the LP spool <NUM> includes the low-pressure (LP) shaft <NUM> and the HP spool <NUM> includes a high pressure (HP) shaft <NUM> positioned concentrically around the LP shaft <NUM>. In such embodiments, the HP shaft <NUM> rotatably couples the rotor blades <NUM> of the HP turbine <NUM> and the rotor blades <NUM> of the HP compressor <NUM> such that rotation of the HP turbine rotor blades <NUM> rotatably drives HP compressor rotor blades <NUM>. As shown, the LP shaft <NUM> is directly coupled to the rotor blades <NUM> of the LP turbine <NUM> and the rotor blades <NUM> of the LP compressor <NUM>. Furthermore, the LP shaft <NUM> is coupled to the fan <NUM> via a gearbox <NUM>. In this respect, the rotation of the LP turbine rotor blades <NUM> rotatably drives the LP compressor rotor blades <NUM> and the fan blades <NUM>.

In several embodiments, the engine <NUM> may generate thrust to propel an aircraft. More specifically, during operation, air (indicated by arrow <NUM>) enters an inlet portion <NUM> of the engine <NUM>. The fan <NUM> supplies a first portion (indicated by arrow <NUM>) of the air <NUM> to the bypass airflow passage <NUM> and a second portion (indicated by arrow <NUM>) of the air <NUM> to the compressor section <NUM>. The second portion <NUM> of the air <NUM> first flows through the LP compressor <NUM> in which the rotor blades <NUM> therein progressively compress the second portion <NUM> of the air <NUM>. Next, the second portion <NUM> of the air <NUM> flows through the HP compressor <NUM> in which the rotor blades <NUM> therein continue progressively compressing the second portion <NUM> of the air <NUM>. The compressed second portion <NUM> of the air <NUM> is subsequently delivered to the combustion section <NUM>. In the combustion section <NUM>, the second portion <NUM> of the air <NUM> mixes with fuel and burns to generate high-temperature and high-pressure combustion gases <NUM>. Thereafter, the combustion gases <NUM> flow through the HP turbine <NUM> which the HP turbine rotor blades <NUM> extract a first portion of kinetic and/or thermal energy therefrom. This energy extraction rotates the HP shaft <NUM>, thereby driving the HP compressor <NUM>. The combustion gases <NUM> then flow through the LP turbine <NUM> in which the LP turbine rotor blades <NUM> extract a second portion of kinetic and/or thermal energy therefrom. This energy extraction rotates the LP shaft <NUM>, thereby driving the LP compressor <NUM> and the fan <NUM> via the gearbox <NUM>. The combustion gases <NUM> then exit the engine <NUM> through the exhaust section <NUM>.

<FIG> is a side view of one embodiment of a shroud block <NUM> of the gas turbine engine <NUM>. In general, several shroud blocks <NUM> are circumferentially arranged to form a shroud (not shown) enclosing or otherwise surrounding one of the rows of rotor blades <NUM> in the compressor section <NUM> or one of the rows of rotor blades <NUM> in the turbine section <NUM>. As shown, the shroud block <NUM> includes an annular wall <NUM> extending between an inner surface <NUM> and an outer surface <NUM> in the radial direction R. The inner surface <NUM> is, in turn, positioned in close proximity to the tips of the corresponding blades <NUM>, <NUM> to minimize the leakage of the air/combustion gases <NUM>/<NUM> past the blades <NUM>, <NUM>. Furthermore, the shroud <NUM> includes a pair of mounting rails <NUM> (one is shown). The rails <NUM> are spaced apart from each other in the longitudinal direction L and extend outward from the outer surface <NUM> of the annular wall <NUM> in the radial direction R. Moreover, each rail <NUM> defines a pair of mounting holes <NUM> for coupling the shroud block <NUM> to the outer casing <NUM> of the engine <NUM>. However, in alternative embodiments, the shroud block <NUM> may have any other suitable configuration.

Additionally, one or more the components of the gas turbine engine <NUM> may be formed of a composite material, such as ceramic matrix composite (CMC) material. For example, in several embodiments, the compressor vanes <NUM>, the compressor blades <NUM>, the turbine vanes <NUM>, the turbine blades <NUM>, and shroud blocks <NUM> may be formed from CMC materials. However, in alternative embodiments, any other suitable components of the engine <NUM> may be formed by composite materials.

The configuration of the gas turbine engine <NUM> described above and shown in <FIG> and <FIG> is provided only to place the present subject matter in an exemplary field of use. Thus, the present subject matter may be readily adaptable 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 gas turbine engines.

<FIG> is a flow diagram of one embodiment of a method <NUM> for repairing composite components. Although <FIG> depicts steps performed in a particular order, the disclosed methods are not limited to any particular order or arrangement. As such, the various steps of the disclosed methods can be omitted, rearranged, combined, and/or adapted in various ways without deviating from the scope of the present disclosure.

In general, the various steps of the method <NUM> will be described below in the context of repairing a composite component <NUM>. For example, as will be described below, the composite component <NUM> may correspond to a composite component of the gas turbine engine <NUM>. However, in alternative embodiments, the composite component <NUM> may correspond to any other suitable composite component.

<FIG> is a perspective view of one embodiment of the composite component <NUM>. In the general, the component <NUM> defines various features therein. More specifically, as shown, in the illustrated embodiment, the component <NUM> defines a hole <NUM> and a slot <NUM>. For example, in one embodiment, the hole <NUM> may be a mounting hole (e.g., the hole <NUM> of the shroud block <NUM>) configured to receive a fastener for use in mounting the component <NUM>. Moreover, in one embodiment, the slot <NUM> may be configured to receive another component, such as a seal (not shown). As such, the hole <NUM> and the slot <NUM> may be blind features (i.e., not a through hole or through slot) that do not extend through the component <NUM>. Thus, the hole <NUM> and the slot <NUM> may extend into the component <NUM> from one surface thereof and terminate before extending through another surface of the component <NUM>. However, in alternative embodiments, the component <NUM> may define any other suitable type or number of features therein, such as additional holes <NUM>, additional slots <NUM>, a channel(s) (not shown), a passage(s) (not shown), and/or the like.

Furthermore, as shown in <FIG>, the composite component <NUM> includes a repair region <NUM>. In general, the repair region <NUM> corresponds to a portion of the component <NUM> that will be repaired in accordance with the method <NUM>. More specifically, the repair region <NUM> may be a worn or damaged portion of the component <NUM>. For example, in the illustrated embodiment, the repair region <NUM> includes several cracks <NUM>. Although the component <NUM> shown in <FIG> only includes one repair region <NUM>, the component <NUM> may, in other embodiments, include any other suitable number of repair regions <NUM>.

Moreover, the composite component <NUM> may be formed from any suitable composite material. For example, the composite material may be selected from the group consisting of, but not limited to, a ceramic matrix composite (CMC), a polymer matrix composite (PMC), a metal matrix composite (MMC), or a combination thereof. Suitable examples of matrix material for a CMC matrix is ceramic powder, including but not limited to, silicon carbide, aluminum-oxide, silicon oxide, and combinations thereof. Suitable examples of matrix material for a PMC include, but are not limited to, epoxy-based matrices, polyester-based matrices, and combinations thereof. Suitable examples of a MMC matrix material include, but are not limited to powder metals such as, but not limited to, aluminum or titanium capable of being melted into a continuous molten liquid metal which can encapsulate fibers present in the assembly, before being cooled into a solid ingot with incased fibers. The resulting MMC is a metal article with increased stiffness, and the metal portion (matrix) is the primary load caring element. For example, in one embodiment, the composite component <NUM> may be formed from a silicon carbide-silicon carbide (SiC-SiC) matrix composite.

Referring again to <FIG>, at (<NUM>), the method <NUM> may include preparing a repair region of a composite component for repair. Specifically, in several embodiments, at (<NUM>), the worn or damaged material of the repair region <NUM> (e.g., the portion of the component <NUM> containing the cracks <NUM>) may be removed from the composite component <NUM> via machining, grinding, cutting, and/or the like. As shown in <FIG>, upon completion of (<NUM>), the repair region <NUM> is a void where the worn/damaged material was originally present.

Additionally, as shown in <FIG>, at (<NUM>), the method <NUM> includes placing a repair material within the prepared repair region. For example, as shown in <FIG>, repair material <NUM> may be placed with the void at prepared repair region <NUM> such that the repair material <NUM> occupies the space where the worn/damaged material was originally present. As will be described below, the repair material <NUM> will be infiltrated such that new composite material is formed in the repair region <NUM>, thereby repairing the component <NUM>. In this respect, the repair material <NUM> corresponds to a precursor material for the composite material from which the component <NUM> is formed. As such, the repair material <NUM> may include a plurality of fibers defining voids that receive the infiltrant. For example, in embodiments in which the component <NUM> is formed from a SiC-SiC matrix composite, the repair material <NUM> may correspond to a silicon carbide (SiC) fiber preform having the same shape and size as the void left in the repair region. However, in alternative embodiments, the repair material <NUM> may correspond to any other suitable composite precursor material, such as a fiber preform formed of another suitable material, fiber tapes, fiber mats, and the like.

Furthermore, as shown in <FIG>, at (<NUM>), the method <NUM> includes filling a feature defined by the composite component with a filler material. For example, as shown in <FIG>, in the illustrated embodiment, the hole <NUM> and the slot <NUM> defined by the component <NUM> are filled with a filler material <NUM>. In general, the filler material <NUM> provides a composite precursor with which infiltrant can bond. As such, when the component <NUM> is infiltrated as will be described below, the infiltrant densifies the filler material (as opposed to simply filling the hole <NUM> and the slot <NUM>), thereby forming new composite material having little to no porosity within the hole <NUM> and the slot <NUM>. Thus, at (<NUM>), one or more features defined by the composite component <NUM> are effectively eliminated.

At (<NUM>), any suitable features of the composite component <NUM> may be filled with filler material <NUM>. More specifically, as described above, in certain instances, it may be desired to fill in one or more features of the component <NUM> with new composite material, thereby eliminating these features. For example, this may be done when it is desired to form a new feature(s) at a different location(s) within the component <NUM>. For example, in the embodiment shown in <FIG>, every feature defined by the component <NUM> (i.e., both the hole <NUM> and the slot <NUM>) is filled with the filler material <NUM> such that every feature will be filled in with new composite material. However, in other instances, it may be desired to only fill in a portion of the features of the component <NUM> with new composite material.

The filler material <NUM> is any suitable precursor to the composite material from which the component <NUM> is formed. For example, in embodiments in which the composite component <NUM> is formed from a SiC-SiC matrix composite, the filler material <NUM> may be silicon carbide. In another embodiment, the composite material <NUM> may be silicon.

Moreover, the filler material <NUM> may be in any suitable form. For example, in one embodiment, the filler material <NUM> may be a slurry. In such an embodiment, the hole <NUM> and the slot <NUM> defined by the component <NUM> are filled with the slurry containing the filler material <NUM>. Furthermore, in another embodiment, the filler material <NUM> may be a powder. In such an embodiment, the hole <NUM> and the slot <NUM> defined by the component <NUM> are filled with a powder form of the filler material <NUM>.

In addition, as shown in <FIG>, after filling the feature with the filler material, at (<NUM>), the method <NUM> includes infiltrating the composite component with an infiltrant to densify the repair region and the filler material such that the feature is filled with new material. Specifically, one or more features of the composite component <NUM> have been filled with the filler material <NUM> at (<NUM>), the composite component <NUM> is infiltrated with a suitable infiltrant. During infiltration, the infiltrant densifies the repair material <NUM> present within the repair region <NUM>, thereby forming new composite material within the repair region <NUM>. Additionally, during infiltration, the infiltrant densifies the filler material <NUM> present within the feature(s) of the component <NUM>, thereby forming new composite material within the feature(s). As such, the feature(s) that have been filled with the filler material <NUM> are effectively eliminated during infiltration.

In several embodiments, at (<NUM>), the method <NUM> may include melt infiltrating the composite component <NUM>. More specifically, as mentioned above, the component <NUM> may be formed from a SiC-SiC matrix composite. In such an embodiment, the repair material <NUM> corresponds to a silicon carbide preform, the filler material <NUM> corresponds to a silicon carbide slurry or powder, and the infiltrant corresponds to silicon. Thus, at (<NUM>), molten silicon may be poured onto the component <NUM>. The molten silicon then infiltrates the repair material <NUM> and the filler material <NUM> by capillary pressure. A first portion of the silicon reacts with the carbon within the repair material <NUM> and the filler material <NUM>. Moreover, a second portion of the carbon fills the voids within the repair material <NUM> and the filler material <NUM> (e.g., the voids between the powder particles in the filler material <NUM>), thereby densifying the repair material <NUM> and the filler material <NUM>. However, in alternative embodiments, any suitable type of infiltration may be used at (<NUM>).

In addition, after infiltrating the composite component, at (<NUM>), the method <NUM> may include forming a second feature within the composite component. More specifically, after the component <NUM> has been infiltrated at (<NUM>), the hole <NUM> and the slot <NUM> are filled in with new composite material. Thus, the hole <NUM> and the slot <NUM> are effectively eliminated from the component <NUM>. As such, in certain instances, it may be desirable to form new features within the component <NUM>. For example, as shown in <FIG>, a new hole <NUM> and a new slot <NUM> are formed within the component <NUM>. In the illustrated embodiment, the new hole <NUM> is spaced apart from the location of the hole <NUM> that has been filled in with new composite material. Furthermore, the new slot <NUM> is spaced apart from the location of the slot <NUM> that has been filled in with new composite material. The new hole <NUM> and the new slot <NUM> may be formed using any suitable method, such as machining, drilling, and/or the like. However, in alternative embodiments, the new hole <NUM> and the new slot <NUM> may be positioned at the same locations as the original hole <NUM> and the original slot <NUM>, such as when it is desired to make such features smaller.

<FIG> is a flow diagram of one embodiment of a method <NUM> for repairing composite turbomachine components. Although <FIG> depicts steps performed in a particular order, the disclosed methods are not limited to any particular order or arrangement. As such, the various steps of the disclosed methods can be omitted, rearranged, combined, and/or adapted in various ways without deviating from the scope of the present disclosure.

In several embodiments, the method <NUM> may be used to repair a composite component(s) of the engine <NUM>. For example, in some embodiments, composite component(s) correspond to a compressor vane(s) <NUM>, a compressor blade(s) <NUM>, a turbine vane(s) <NUM>, a turbine blade(s) <NUM>, and/or a shroud block(s) <NUM> of the engine <NUM>. However, in alternative embodiments, the composite component(s) may correspond to any suitable component(s), such as other component(s) of a turbomachine or component(s) of any other turbomachine.

As shown in <FIG>, at (<NUM>), the method <NUM> may include preparing a repair region of a composite turbomachine component for repair. Additionally, at (<NUM>), the method <NUM> includes placing a repair material within the prepared repair region. Furthermore, at (<NUM>), the method <NUM> includes filling a feature defined by the composite turbomachine component with a filler material. Moreover, after filling the feature with the filler material, at (<NUM>), the method <NUM> includes infiltrating the composite turbomachine component with an infiltrant to densify the repair region and the filler material such that the feature is filled with new material. In addition, after infiltrating the composite turbomachine component, at (<NUM>), the method <NUM> may include forming a second feature within the composite turbomachine component.

Claim 1:
A method (<NUM>) for repairing a composite turbomachine component (<NUM>), the method comprising:
positioning repair material (<NUM>) within a repair region (<NUM>) of the composite component (<NUM>) formed of a composite material; characterised in that the method further comprises the steps of:
filling a feature defined by the composite component (<NUM>) with a filler material (<NUM>), the filler material being a precursor to the composite material, wherein the feature defined by the composite component comprises a hole (<NUM>) or a slot (<NUM>); and
after filling the feature with the filler material (<NUM>), infiltrating the composite component (<NUM>) with an infiltrant to densify the repair region (<NUM>) and the filler material (<NUM>) such that the feature is filled with new material.