Composite airfoil assembly for turbine engine

Abstract

A composite airfoil assembly for a turbine engine. The composite airfoil assembly has an airfoil outer surface defining opposing pressure and suction sides that extend between a leading edge and a trailing edge. A composite airfoil assembly includes a composite airfoil having a braided core and a skin applied to at least a portion of an exterior of the core. The composite airfoil assembly also includes a cladding coupled to the skin outer surface, where at least one edge or portion of the cladding is positioned adjacent to at least one of the trailing edge, root, or tip, where the cladding includes a tapered portion.

Description

Technical Field

[0001] The present disclosure relates generally to composite airfoil assemblies and, more particularly, to composite airfoil assemblies for turbine engines. Background Art

[0002] A turbine engine typically includes an engine core having a compressor section, a combustor section, and a turbine section in a serial flow arrangement. The fan section may be disposed upstream of the compressor section. The compressor section compresses air that is directed to the combustor section, where the air is mixed with fuel and then the mixture is ignited in the combustor section to generate hot combustion gases. The combustion gases are directed to the turbine section, which extracts energy from the combustion gases for powering the compressor section and for generating useful work to propel an aircraft in flight or to power a load such as a generator.

[0003] With the advent of composite materials, composite materials have been used to make components of gas turbine engines, particularly in lower temperature areas, such as blades in fan sections. Composite materials typically include a fiber reinforced matrix and exhibit a high strength-to-weight ratio. Due to the high strength-to-weight ratio and formability into relatively complex shapes, composite materials are used in a variety of applications, such as turbine engines or aircraft. For example, composite materials can be mounted on a fuselage and / or wing, rudder, manifold, airfoil, or other component of an aircraft or turbine engine, or define a portion thereof. BRIEF DESCRIPTION OF THE DRAWINGS

[0004] A full and enabling disclosure of the present disclosure, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings, in which:

[0005] Figure 1 is a schematic cross-sectional view of a non-ducted or open rotor turbine engine.

[0006] Figure 2 Yes Figure 1 A schematic perspective view of an aircraft having a non-ducted or open rotor turbine engine.

[0007] Figure 3 is suitable for use in an exemplary embodiment according to the present disclosure Figure 1 and Figure 2 A schematic perspective view of a composite airfoil assembly and a disk assembly for use within a turbine engine of FIG. 1 , the composite airfoil assembly comprising a composite airfoil, a cladding, and a dovetail.

[0008] Figure 4 According to an exemplary embodiment of the present disclosure Figure 3 An enlarged view of the composite airfoil assembly.

[0009] Figure 5 According to an exemplary embodiment of the present disclosure, Figure 4 Schematic cross-sectional view taken along line VV of FIG. 1 , showing the interior of the composite airfoil with the cladding removed.

[0010] Figure 6 is a cladding having a cladding according to an exemplary embodiment of the present disclosure Figure 5 Schematic cross section of the interior of a composite airfoil.

[0011] Figure 7 According to an exemplary embodiment of the present disclosure Figure 6 A variation of the schematic cross-sectional view of FIG.

[0012] Figure 8 According to an exemplary embodiment of the present disclosure Figure 7 An enlarged schematic cross-sectional view of a portion of the first cladding. DETAILED DESCRIPTION

[0013] Aspects disclosed herein relate to a composite component for manufacturing an engine component of a turbine engine. The composite component is shown as a composite airfoil assembly having a composite airfoil, the composite airfoil including at least a core and a skin shown as a braided core and a laminated skin. The composite airfoil assembly may be located in a fan section of a turbine engine. The laminated skin has a skin outer surface, wherein a first portion of the skin outer surface defines a first portion of an airfoil outer surface of the composite airfoil assembly. A second portion of the skin outer surface is recessed to receive a cladding.

[0014] The cladding received by the recess has a cladding outer surface that defines a second portion of the airfoil outer surface of the composite airfoil assembly. That is, the composite airfoil includes one or more recesses that can receive the cladding so that the outer surface of the cladding and the first portion of the skin outer surface define the airfoil outer surface of the composite airfoil assembly.

[0015] As used herein, the term "cladding" refers to a reinforcement or material added to a composite airfoil to increase the stiffness and / or strength at a certain location of the composite airfoil. The cladding can reduce bending forces or moments in a desired direction.

[0016] The benefits of the composite airfoil assembly include weight savings of the composite airfoil, strength, stiffness, or strength and stiffness benefits of the cladding, and aerodynamic benefits of the cladding being recessed into the composite airfoil. The smooth outer airfoil surface provides fewer interruptions to the airfoil passing through the composite airfoil assembly. That is, the outer airfoil surface does not include positive bumps. The improved aerodynamics results in an overall better airflow, which improves the efficiency of the engine.

[0017] It should be understood that, in a non-limiting example, the present disclosure is applicable to other engine components of a turbine engine, not just airfoils, such as disks or combustor liners. In addition, although described in terms of a core for making an airfoil, it should be understood that the present disclosure is applicable to any other suitable environment.

[0018] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter references to refer to features in the drawings. Like or similar reference numbers in the drawings and description have been used to refer to like or similar parts of the present disclosure.

[0019] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. In addition, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

[0020] As used herein, the terms “first,” “second,” “third,” or “fourth” may be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0021] The term "at least one" in a context such as "at least one of A, B, and C" means only A, only B, only C, or any combination of A, B, and C.

[0022] The terms “turbomachine” or “turbomachine” refer to a machine that includes one or more compressors, a heat generating section (eg, a combustion section), and one or more turbines that together generate a torque output.

[0023] The term "turbine engine" refers to an engine having a turbine as all or part of its power source. Example turbine engines include turbofan engines, turboprop engines, turbojets, turboshaft engines, etc., as well as hybrid electric versions of one or more of these engines.

[0024] The terms "front" and "rear" refer to relative positions within a turbine engine or vehicle and refer to the normal operating attitude of the turbine engine or vehicle. For example, for a turbine engine, front refers to a position closer to the engine inlet, while rear refers to a position closer to the engine nozzle or exhaust.

[0025] As used herein, the term "upstream" refers to the direction opposite to the direction of fluid flow, while the term "downstream" refers to the direction in the same direction as the direction of fluid flow. The terms "front" or "forward" mean in front of something, and "rear" or "rearward" mean behind something. For example, when used in relation to fluid flow, front / forward may mean upstream, while rear / rearward may mean downstream.

[0026] The term "fluid" may be gas, liquid or multiphase.

[0027] As used herein, the terms "axial" and "axially" refer to directions and orientations extending substantially parallel to the centerline of a turbine engine. In addition, the terms "radial" and "radially" refer to directions and orientations extending substantially perpendicular to the centerline of a turbine engine. For example, in the context of a turbine engine as a whole, radial refers to directions along a ray extending between the central longitudinal axis of the engine and the periphery of the engine. In addition, as used herein, the terms "circumferential" and "circumferentially" refer to directions and orientations extending in an arc around the centerline of a turbine engine.

[0028] Unless otherwise specified herein, the terms "coupled," "fixed," "applied to," "attached to," and the like refer to both direct coupling, fixing, or attachment as well as indirect coupling, fixing, or attachment through one or more intermediate components or features.

[0029] All directional references (e.g., radial, axial, up, down, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, backward, etc.) are used only for identification purposes to help the reader understand the present disclosure, and do not create limitations, particularly limitations on the position, orientation, or use of aspects of the present disclosure described herein. Unless otherwise stated, connection references (e.g., attachment, connection, connection, and engagement) will be broadly interpreted and may include intermediate structural elements between sets of elements and relative movement between elements. Therefore, connection references do not necessarily infer that two elements are directly connected and have a fixed relationship to each other. The exemplary drawings are only for illustrative purposes, and the dimensions, positions, orders, and relative sizes reflected in the attached drawings may vary.

[0030] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise. Also, as used herein, the term "set" or a "set of" elements may refer to any number of elements, including only one.

[0031] As used herein, the term "stiffness" may be used to define the degree to which a structure resists deformation in response to a force. Stiffness may be defined as the ratio of the force to the displacement of an object under the action of a force. Stiffness may include resistance to deformation in response to forces applied from various directions, so that in non-limiting examples, stiffness may refer to axial stiffness, tensile stiffness, compressive stiffness, torsional stiffness, or shear stiffness.

[0032] As used herein, the term "composite" refers to a component having two or more materials. A composite can be a combination of at least two or more metals, non-metals, or a combination of metal and non-metal elements or materials. Examples of composite materials can be, but are not limited to, polymer matrix composites (PMC), ceramic matrix composites (CMC), metal matrix composites (MMC), carbon fibers, polymer resins, thermoplastics, bismaleimide (BMI) materials, polyimide materials, epoxy resins, glass fibers, and silicon matrix materials.

[0033] As used herein, a "composite" component refers to a structure or component that includes any suitable composite material. A composite component, such as a composite airfoil, may include several layers or several layups of composite materials. The stiffness, material, and dimensions of the layers or layups may vary to achieve a desired composite component or composite portion of a component having a predetermined weight, size, stiffness, and strength.

[0034] One or more adhesive layers may be used to form or join composite components. The adhesive may include resins and phenolics, where the adhesive may require curing at elevated temperatures or other hardening techniques.

[0035] In the present disclosure, when a layer is described as being "on" or "over" another layer or substrate, it should be understood that the layers may be directly in contact with each other or have another layer or feature between the layers unless explicitly stated to the contrary. Therefore, these terms simply describe the relative position of layers with respect to each other and do not necessarily mean "on top of...", as the relative position of above or below depends on the orientation of the device with respect to the viewer.

[0036] As used herein, PMC refers to a class of materials. As an example, PMC materials are partially defined by prepregs, which are reinforcing materials pre-impregnated with polymer matrix materials (e.g., thermoplastic resins). Non-limiting examples of processes for producing thermoplastic prepregs include: hot melt prepregs, in which fiber reinforcements are pulled through a molten bath of resin; and powder prepregs, in which resin is deposited on the fiber reinforcements, as a non-limiting example, electrostatically deposited on the fiber reinforcements, and then adhered to the fibers, as a non-limiting example, in an oven or with the help of heated rollers. Prepregs can be in the form of unidirectional tapes or woven fabrics, which are then stacked on top of each other to form the desired number of stacked layers of a part.

[0037] Multilayer prepreg is stacked to the appropriate thickness and orientation of composite parts, and then resin is cured and solidified to provide fiber reinforced composite parts. The resin for PMC matrix material can be generally classified as thermosetting resin or thermoplastic resin. Thermoplastic resin is generally classified as a polymer that can be repeatedly softened and flowed when heated and can be hardened due to physical changes rather than chemical changes when fully cooled. The famous example category of thermoplastic resin includes nylon, thermoplastic polyester, polyaryletherketone and polycarbonate resin. The specific example of high-performance thermoplastic resin for aerospace applications has been envisioned to include polyetheretherketone (PEEK), polyetherketoneketone (PEKK), polyetherimide (PEI), polyaryletherketone (PAEK) and polyphenylene sulfide (PPS). On the contrary, once fully cured into a hard rigid solid, thermosetting resin will not experience significant softening when heated, but thermally decomposes when fully heated. The famous example of thermosetting resin includes epoxy resin, bismaleimide (BMI) and polyimide resin.

[0038] Instead of using prepreg, in another non-limiting example, by using thermoplastic polymers, braided fabrics can be utilized. Braided fabrics may include, but are not limited to, dry carbon fibers woven together with thermoplastic polymer fibers or filaments. Non-prepreg braided architectures can be made in a similar manner. In this way, the fiber volume of a part can be customized by specifying the relative concentrations of thermoplastic fibers and reinforcing fibers that have been woven or braided together. In addition, different types of reinforcing fibers can be braided or braided together at different concentrations to customize the properties of a part. For example, glass fibers, carbon fibers, and thermoplastic fibers can all be woven together at different concentrations to customize the properties of a part. Carbon fibers provide the strength of the system, glass fibers can be incorporated to enhance impact properties, which is a design feature of a part located near the engine inlet, and thermoplastic fibers provide bonding for reinforcing fibers.

[0039] In yet another non-limiting example, resin transfer molding (RTM) can be used to form at least a portion of a composite component. Typically, RTM includes applying a dry fiber or matrix material to a mold or cavity. The dry fiber or matrix material may include prepregs, woven materials, braided materials, or any combination thereof.

[0040] The resin may be pumped or otherwise provided to the mold or cavity to impregnate the dry fibers or matrix material. The combination of the impregnated fibers or matrix material and the resin is then cured and removed from the mold. When removed from the mold, the composite part may require a post-curing process.

[0041] It is contemplated that RTM may be a vacuum assisted process. That is, the air in the cavity or mold may be removed and replaced with resin prior to heating or curing. It is further contemplated that placement of dry fiber or matrix material may be manual or automated.

[0042] The dry fiber or matrix material may be shaped to form the composite part or to guide the resin. Optionally, additional layers or reinforcements of a material different from the dry fiber or matrix material may also be included or added prior to heating or curing.

[0043] As used herein, CMC refers to a class of materials having reinforcing fibers in a ceramic matrix. Typically, the reinforcing fibers provide structural integrity to the ceramic matrix. Some examples of reinforcing fibers may 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, aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates (such as mullite), or mixtures thereof), or mixtures thereof.

[0044] Some examples of ceramic matrix materials may 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, aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates, or mixtures thereof), or mixtures thereof. Optionally, ceramic components (e.g., oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite) may also be included in the ceramic matrix.

[0045] In general, specific CMCs can be referred to as their fiber type / matrix type combination. For example, C / SiC is carbon fiber reinforced silicon carbide, SiC / SiC is silicon carbide fiber reinforced silicon carbide, SiC / SiN is silicon carbide fiber reinforced silicon nitride, SiC / SiC-SiN is silicon carbide fiber-reinforced silicon carbide / silicon nitride matrix mixture, etc. In other examples, CMCs can be composed of a matrix and reinforcing fibers comprising oxide-based materials such as aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates, and mixtures thereof. Aluminosilicates can include crystalline materials such as mullite (3Al2O3·2SiO2), as well as glassy aluminosilicates.

[0046] In certain non-limiting examples, the reinforcing fibers may be bundled and / or coated before being included in the ceramic matrix. For example, the fiber bundle may be formed into a reinforcing tape, such as a unidirectional reinforcing tape. Multiple tapes may be stacked together to form a preform component. The fiber bundle may be impregnated with a slurry composition before or after forming the preform. The preform may then be subjected to heat treatment and subsequent chemical treatment (such as silicon melt infiltration) to obtain a component formed of a CMC material having a desired chemical composition. For example, the preform may undergo solidification or burnout to produce a high coke residue in the preform, and then undergo melt infiltration with silicon, or undergo solidification or pyrolysis to produce a silicon carbide matrix in the preform, and then undergo chemical vapor infiltration with silicon carbide. Additional steps may be taken to improve the densification of the preform, by injecting the preform with a liquid resin or polymer before or after chemical vapor infiltration, and then performing a heat treatment step to fill the voids with silicon carbide. CMC materials as used herein may be formed using any known or later developed method including, but not limited to, melt infiltration, chemical vapor infiltration, polymer infiltration pyrolysis (PIP), or any combination thereof.

[0047] The reinforcing fibers may be at least some portions of individual filaments or strands. As used herein, "ceramic fiber tow," "fiber tow," or simply "tow" refers to a bundle of multiple individual fibers, filaments, or loose strands. The filaments of the tow may be randomly mixed or arranged in a pattern, and / or may be continuous or discontinuous. For example, the tow may include broken filaments or filament segments. As another example, the filaments of the tow may be substantially parallel, twisted, or otherwise arranged. The tow may function in substantially the same manner as a single or individual filament. It will also be understood that "individual ceramic filaments" or simply "individual filaments" used herein refers to a single or non-bundled elongated ceramic member.

[0048] Such materials, along with certain monolithic ceramics (i.e., ceramic materials without reinforcement materials), are particularly suitable for higher temperature applications. In addition, these ceramic materials are lightweight compared to superalloys, but still provide strength and durability to the parts made from them. As a result, such materials are currently being considered for use in many turbine components used in the higher temperature sections of turbine engines (such as airfoils (e.g., turbine blades and vanes), combustors, shrouds, etc.), which would benefit from the lighter weight and higher temperature capabilities that these materials can provide.

[0049] The term "metal" as used herein refers to materials including metals such as, but not limited to, titanium, iron, aluminum, stainless steel, brass, copper, and nickel alloys. A metal material or alloy may be a combination of at least two or more elements or materials, at least one of which is a metal. As used herein, the term "additive manufacturing" generally refers to a manufacturing process in which a feedstock of a material in the form of a particulate powder or wire is aggregated to form a three-dimensional component. The feedstock material is then melted by applying heat or other curing treatments to form an integral single component that may have various integral subcomponents. As used herein, integral refers to a single structure in which the individual layers lose their characteristics in the final single structure by virtue of the material of each layer being melted to or with the material of an adjacent layer and lacking an interface or joint.

[0050] Suitable additive manufacturing techniques according to the present disclosure include, for example, directed energy deposition (DED), fused deposition modeling (FDM), selective laser sintering (SLS), 3D printing such as by inkjet and laser jetting, stereolithography (SLA), direct selective laser sintering (DSLS), electron beam sintering (EBS), electron beam melting (EBM), laser engineered net shaping (LENS), laser net shape manufacturing (LNSM), direct metal deposition (DMD), digital light processing (DLP), direct selective laser melting (DSLM), selective laser melting (SLM), direct metal laser melting (DMLM) and other known processes.

[0051] Digital light processing (DLP) can include a 3D DLP printer having a transparent vat or tank, a building platform, and a light assembly. The vat or tank can contain, for example, a photopolymer resin.

[0052] The DLP build platform may be coupled to, for example, a motor or other mechanism that allows the build platform to move in one or more dimensions, such as raising or lowering the build platform from or toward resin in a vat or tank.

[0053] The DLP printing component can be coupled to a lower portion of the building platform facing the barrel or tank. The lighting assembly is at least partially located below the barrel or tank. The lighting assembly may include at least one light source and at least one optical reflector or refractor, such as a deflecting mirror or at least one lens.

[0054] A controller coupled to or included in a DLP printer can control one or more aspects of the DLP printer, such as the position of a DLP build platform or the intensity, duration, or orientation of a light source.

[0055] In addition to using a direct metal laser sintering (DMLS), direct metal laser melting (DMLM) process, or electron beam melting (EBM) process in which an energy source is used to selectively sinter or melt portions of a powder layer, it will be appreciated that, according to alternative aspects of the present disclosure, the additive manufacturing process may be a "binder jetting" process. In this regard, binder jetting involves the continuous deposition of layers of additive powder in a manner similar to that described above. However, instead of using an energy source to generate an energy beam to selectively melt or fuse the additive powder, binder jetting involves selectively depositing a liquid binder onto each powder layer. The liquid binder may be, for example, a photocurable polymer or another liquid binder. Other suitable additive manufacturing methods and variations are intended to fall within the scope of the present subject matter.

[0056] Figure 1 is a schematic cross-sectional view of a turbine engine, particularly an open rotor or non-ducted turbine engine 10 for an aircraft. The non-ducted turbine engine 10 has a generally longitudinally extending axis or engine centerline 12 extending from a forward end 14 to an aft end 16. The non-ducted turbine engine 10 includes, in downstream serial flow relationship, a set of circumferentially spaced blades or propellers defining: a fan section 18 including a fan 20; a compressor section 22 including a supercharger or low pressure (LP) compressor 24 and a high pressure (HP) compressor 26; a combustion section 28 including a combustor 30; a turbine section 32 including an HP turbine 34 and an LP turbine 36; and an exhaust section 38. The non-ducted turbine engine 10 as described herein is a non-limiting example, and other architectures are possible, such as, but not limited to, a steam turbine engine, a supercritical carbon dioxide turbine engine, or any other suitable turbine engine.

[0057] The exterior surface of the non-ducted turbine engine 10 defined by a casing, such as a nacelle 40, extends from a forward end 14 of the non-ducted turbine engine 10 toward a rearward end 16 of the non-ducted turbine engine 10 and covers at least a portion of the compressor section 22, the combustion section 28, the turbine section 32, and the exhaust section 38. The fan section 18 may be positioned forward of the nacelle 40 and extends radially outward from the nacelle 40 of the non-ducted turbine engine 10, specifically, the fan section 18 extends radially outward from the nacelle 40. The fan section 18 includes a set of fan blades 42 and a set of stationary fan blades 82 downstream of the set of fan blades 42, both of which are radially arranged around the engine centerline 12. The non-ducted turbine engine 10 includes any number of one or more sets of rotating blades or propellers (e.g., the set of fan blades 42) disposed upstream of the set of stationary fan blades 82. As a non-limiting example, the non-ducted turbine engine 10 may include multiple sets of fan blades 42 or the set of stationary fan blades 82. Therefore, the non-ducted turbine engine 10 is further defined as a non-ducted single fan turbine engine. The non-ducted turbine engine 10 is further defined by the position of the fan section 18 relative to the combustion section 28. The fan section 18 may be upstream, downstream, or aligned with the axial positioning of the combustion section 28.

[0058] The compressor section 22 , the combustion section 28 , and the turbine section 32 are collectively referred to as an engine core 44 , which generates combustion gases. The engine core 44 is surrounded by an engine casing 46 operably coupled to a portion of a nacelle 40 of the non-ducted turbine engine 10 .

[0059] An HP shaft or spool 48 disposed coaxially about the engine centerline 12 of the non-ducted turbine engine 10 drivingly connects the HP turbine 34 to the HP compressor 26. An LP shaft or spool 50 disposed coaxially about the engine centerline 12 of the non-ducted turbine engine 10 within the larger diameter annular HP spool 48 drivingly connects the LP turbine 36 to the LP compressor 24 and the fan 20. The spools 48, 50 are rotatable about the engine centerline 12 and are coupled to a set of rotatable elements that collectively define a rotor 51.

[0060] It should be appreciated that the non-ducted turbine engine 10 is a direct drive or integral drive engine utilizing a reduction gearbox coupling the LP shaft or spool 50 to the fan 20 .

[0061] The LP compressor 24 and the HP compressor 26 each include a set of compressor stages 52, 54, wherein a set of compressor blades 56, 58 rotate relative to a corresponding set of static compressor vanes 60, 62 (also referred to as nozzles) to compress or pressurize the fluid flow passing through the stage. In a single compressor stage 52, 54, a plurality of compressor blades 56, 58 are arranged in a ring and extend radially outward from a blade platform to a blade tip relative to the engine centerline 12, while the corresponding static compressor vanes 60, 62 are positioned upstream of and adjacent to the compressor blades 56, 58. Notably, Figure 1 The numbers of blades, buckets, and compressor stages shown in FIG. 5 are selected for illustration purposes only, and other numbers are possible.

[0062] Compressor blades 56, 58 for one stage of the compressor are mounted to disks 61 mounted to respective ones of the HP spool 48 and the LP spool 50, with each stage having its own disk 61. Static compressor vanes 60, 62 for one stage of the compressor are mounted to the engine case 46 in a circumferential arrangement.

[0063] The HP turbine 34 and the LP turbine 36 each include a set of turbine stages 64, 66 in which a set of turbine blades 68, 70 rotate relative to a corresponding set of static turbine vanes 72, 74 (also referred to as nozzles) to extract energy from the fluid flow passing through the stage. In a single turbine stage 64, 66, a plurality of turbine blades 68, 70 are arranged in a ring and extend radially outward from the blade platform to the blade tip relative to the engine centerline 12, while the corresponding static turbine vanes 72, 74 are positioned upstream of and adjacent to the turbine blades 68, 70. Notably, Figure 1 The number of blades, buckets, and turbine stages shown in FIG. 5 are selected for illustrative purposes only, and other numbers are possible.

[0064] Turbine blades 68, 70 for one stage of turbine section 32 are mounted to disks 71 mounted to a respective one of HP spool 48 and LP spool 50, with each stage having a dedicated disk 71. Static turbine buckets 72, 74 for one stage of turbine section 32 are mounted to engine casing 46 in a circumferential arrangement.

[0065] The rotating portions of the non-ducted turbine engine 10, such as the blades 56, 58, 68, 70 in the compressor section 22 and the turbine section 32, are also individually or collectively referred to as rotors 51. Thus, a rotor refers to the combination of rotating elements throughout the non-ducted turbine engine 10.

[0066] Complementing the rotor portion, the stationary portion of the non-ducted turbine engine 10 (e.g., the stationary blades 60, 62, 72, 74 in the compressor section 22 and the turbine section 32) are also individually or collectively referred to as stators 63. Therefore, the stator 63 refers to the combination of non-rotating elements throughout the non-ducted turbine engine 10.

[0067] The nacelle 40 is operably coupled to the non-ducted turbine engine 10 and covers at least a portion of the engine core 44, the engine case 46, or the exhaust section 38. At least a portion of the nacelle 40 extends axially forward or upstream of the position shown. For example, the nacelle 40 extends axially forward so that a portion of the nacelle 40 covers or covers a portion of the fan section 18 or the supercharger section (not shown) of the non-ducted turbine engine 10.

[0068] During operation of non-ducted turbine engine 10, free stream airflow 80 flows against the front of non-ducted turbine engine 10. A portion of free stream airflow 80 enters annular region 25 defined by the outer surface of nacelle 40 and the swept area between the tips of the blades, where the airflow is inlet airflow 78. A portion of inlet airflow 78 enters engine core 44 and is described as working airflow 76, which is used for combustion within engine core 44.

[0069] More specifically, the working airflow 76 flows into the LP compressor 24, which then pressurizes the working airflow 76, thereby defining a pressurized airflow that is supplied to the HP compressor 26, which further pressurizes the air. The working airflow 76 or pressurized airflow from the HP compressor 26 is mixed with the fuel in the combustor 30 and ignited, thereby generating combustion gases. The HP turbine 34 extracts some work from these gases, which drives the HP compressor 26. The combustion gases are discharged into the LP turbine 36, which extracts additional work to drive the LP compressor 24, and the working airflow 76 or exhaust gas is ultimately discharged from the non-ducted turbine engine 10 via the exhaust section 38. The drive of the LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP compressor 24. The working airflow 76, including the pressurized airflow and the combustion gases, defines the working airflow that flows through the compressor section 22, the combustion section 28, and the turbine section 32 of the non-ducted turbine engine 10.

[0070] Inlet airflow 78 flows through the set of fan blades 42 and over nacelle 40 of non-ducted turbine engine 10. Inlet airflow 78 then flows over at least a portion of the set of stationary fan blades 82, which directs inlet airflow 78 laterally toward engine centerline 12. Inlet airflow 78 then follows the curvature of nacelle 40 and flows through the set of stationary fan blades 82 toward exhaust section 38. Pylon 84 mounts non-ducted turbine engine 10 to an external structure (e.g., a fuselage, wing, tail, etc. of an aircraft).

[0071] Working airflow 76 and at least some inlet airflow 78 merge downstream of exhaust section 38 of non-ducted turbine engine 10. Working airflow 76 and inlet airflow 78 together create the overall thrust of non-ducted turbine engine 10.

[0072] It is envisioned that a portion of the working airflow 76 is extracted as bleed air 77 (e.g., from the compressor section 22). The bleed air 77 provides airflow to engine components that require cooling. The temperature of the working airflow 76 leaving the combustor 30 is significantly increased relative to the working airflow 76 within the compressor section 22. Therefore, the cooling provided by the bleed air 77 is necessary for operating these engine components in an elevated temperature environment or in the hot section of the non-ducted turbine engine 10. In the case of a turbine engine, the hot section of the engine is typically downstream of the combustor 30, particularly the turbine section 32, with the HP turbine 34 being the hottest section because it is located directly downstream of the combustion section 28. Other sources of cooling fluid are, but are not limited to, fluid discharged from the LP compressor 24 or the HP compressor 26.

[0073] Figure 2 is suitable for use as Figure 1 Schematic perspective view of an aircraft 86 of a generic non-ducted turbine engine 88 of a non-ducted turbine engine 10. Aircraft 86 includes a fuselage 90 having an exterior surface. At least one wing 92 and an empennage 94 extend from fuselage 90. Emitter 94 is operably coupled to and spaced apart from fuselage 90 via an empennage pylon 96. Non-ducted turbine engine 88 is operably coupled to the exterior surface of fuselage 90 via pylon 98. Non-ducted turbine engine 88 includes a set of circumferentially spaced fan blades 100. A set of stationary fan blades 102 are disposed downstream of the set of circumferentially spaced fan blades 100. Fuselage 90 extends between a nose 104 and a tail 106 and includes a fuselage centerline 108 extending therebetween.

[0074] Additionally, while the empennage 94 is a T-wing empennage (e.g., as shown in the figure), other conventional empennages are contemplated, such as a cruciform empennage, an H-tail, a triple tail, a V-tail, an inverted tail, a Y-tail, a double tail, a boom tail, or a ring tail, all of which are referred to herein as the empennage 94.

[0075] Figure 3 Is suitable for Figure 1 A non-ducted turbine engine 10 or Figure 2 A schematic perspective view of a composite airfoil assembly 110 and a disk assembly 112 for use within a non-ducted turbine engine 88. The disk assembly 112 is suitable for use as a disk 61, 71 ( Figure 1) or any other disk, such as, but not limited to, a disk within the fan section 18, the compressor section 22, or the turbine section 32 of the non-ducted turbine engine 10. The composite airfoil assembly 110 may be rotating or non-rotating, such that the composite airfoil assembly 110 may include static compressor blades 60, 62 ( Figure 1 ), the set of compressor blades 56, 58 ( Figure 1 ), static turbine blades 72, 74 ( Figure 1 ), the group of turbine blades 68, 70 ( Figure 1 ) or the set of fan blades 42 ( Figure 1 ). As a non-limiting example, the composite airfoil assembly 110 may be a composite fan blade assembly.

[0076] The disk assembly 112 may be rotatable or stationary about an axis of rotation 114. The axis of rotation 114 may be aligned with an engine centerline (e.g., Figure 1 The disk assembly 112 includes a plurality of slots 116 extending axially through a radially outer portion of the disk assembly 112 and spaced circumferentially about the disk assembly 112 relative to the rotational axis 114.

[0077] The composite airfoil assembly 110 includes a composite airfoil 118 and a cladding 120. The composite airfoil 118 extends between a leading edge 124 and a trailing edge 126 opposite the leading edge 124 to define a chordwise direction. The composite airfoil 118 extends between a root 128 and a tip 130 to define a spanwise direction. The composite airfoil 118 includes a pressure side 132 and a suction side 134 opposite the pressure side 132.

[0078] Leading edge 124 and trailing edge 126 extend radially from root 128 to tip 130. A pressure side 132 and a suction side 134 opposite pressure side 132 extend axially between leading edge 124 and trailing edge 126. A dovetail portion 138 may extend from composite airfoil 118.

[0079] The composite airfoil assembly 110 is coupled to the disk assembly 112 by inserting at least a portion of the dovetail portion 138 into a corresponding slot in the plurality of slots 116. The composite airfoil assembly 110 is held in place by frictional contact with the slots 116, or may be coupled to the slots 116 via any suitable coupling method, such as, but not limited to, welding, adhering, fastening, etc. Although only a single composite airfoil assembly 110 is shown, it should be understood that there may be any number of composite airfoil assemblies 110 coupled to the disk assembly 112. As a non-limiting example, there may be a number of composite airfoil assemblies 110 corresponding to the total number of slots in the plurality of slots 116.

[0080] For ease of reference, a set of relative reference directions and coordinate systems may be applied to the composite airfoil assembly 110. An axial direction (Ad) may extend from front to back and is shown as extending at least partially into the page. The axial direction (Ad) may be arranged parallel to the axis of rotation 114. A radial direction (Rd) extends perpendicular to the axial direction (Ad) and may extend perpendicular to the engine centerline 12. A circumferential direction (Cd) may be defined as being perpendicular to the radial direction (Rd) and may be defined relative to the engine centerline 12 ( Figure 1 ) along the non-ducted turbine engine 10 ( Figure 1 )’s circumference.

[0081] Figure 4 1 is an enlarged view of the composite airfoil assembly 110, further illustrating the cladding 120. The airfoil length 140 may be measured from the root 128 to the tip 130 of the composite airfoil 118. Although shown as variable, the airfoil length 140 may be an average length, a weighted average, a minimum length, or a maximum length measured from the root 128 to the tip 130 of the composite airfoil 118 in a radial direction (Rd) or a spanwise direction defined from the root 128 to the tip 130.

[0082] The airfoil outer surface 142 of the composite airfoil assembly 110 is defined by a pressure side 132 and a suction side 134, which is opposite the pressure side 132. The pressure side 132 and the suction side 134 extend radially from the root 128 to the tip 130 in the span direction. The suction side 134 may be a mirror image of the pressure side 132 and have the same surface area. However, it is contemplated that the surface area of ​​the suction side may be less than or greater than the surface area of ​​the pressure side 132.

[0083] The axial airfoil length 144 may be measured between the leading edge 124 and the trailing edge 126 of the composite airfoil 118. Although shown as variable, the axial airfoil length 144 may be measured from the leading edge 124 to the trailing edge 126 of the composite airfoil 118 in the axial direction (Ad), the chord-wise direction, or an average length, a weighted average, a minimum length, or a maximum length measured from the leading edge 124 to the trailing edge 126 along a mean camber line equidistant from the pressure side 132 and the suction side 134.

[0084] The cladding 120 includes two or more cladding elements (or cladding layers). The cladding 120 refers to the combination of all cladding elements on the composite airfoil assembly 110. Figure 4 and Figure 6As shown in , for example, the cladding 120 includes a first cladding 150 and a third cladding 154 at or adjacent the trailing edge 126 , a leading edge jacket 156 at the leading edge 124 , a second cladding 152 at or adjacent the tip 130 , and a root cladding assembly 158 at the root 128 .

[0085] The cladding 120 is coupled to or applied to the composite airfoil 118 at an outer surface 142 of the composite airfoil 118 , which is shown as a skin outer surface 190 .

[0086] The first cladding 150 has a first perimeter defined by edges. As shown, as an example, the perimeter of the first cladding 150 may include a root edge 151a located closest to the root 128, a tip edge 151b located closest to the tip 130, an axial forward edge 151c located at the pressure side 132 between the leading edge 124 and the trailing edge 126, and an axial rearward edge 151d adjacent to or at the trailing edge 126.

[0087] The second cladding 152 has a second periphery, which is shown as a root edge 153a positioned closest to the root 128, a tip edge 153b located at the tip 130, an axial forward edge 153c located at the pressure side 132 between the leading edge 124 and the trailing edge 126, and an axial rearward edge 153d adjacent to the trailing edge 126 or at the trailing edge 126.

[0088] The cladding 120 may further include a third cladding 154 ( Figure 6 ). The third cladding layer 156 ( Figure 6 ) has a similar geometry to the first cladding 150, but is located on the suction side 134 at the trailing edge 126. As used herein, when an object has a "similar geometry" to a second object, the surface area of ​​the second object is within 10% of the surface area of ​​the first object. That is, the cladding 120 on the suction side 134 may mirror portions of the cladding on the pressure side 132. However, in different and non-limiting examples, it is contemplated that the cladding 120 on the suction side 134 may be different in size and location when compared to the cladding 120 on the pressure side 132.

[0089] Alternatively, in a different and non-limiting example, the first cladding 150 may be integrally formed on at least a portion of the trailing edge 126. That is, the first cladding 150 may have portions on both the suction side 134 and the pressure side 132, such that a portion of the first cladding 150 covers a portion of the trailing edge 126. In other words, the first cladding 150 may wrap around the trailing edge 126 of the composite airfoil 118, with a portion of the first cladding 150 on the pressure side 132 and another portion extending to the suction side 134. In this non-limiting example where the first cladding 150 wraps around the trailing edge 126, the first cladding 150 may have a V-shaped or U-shaped cross-section.

[0090] The cladding 120 may also include a fourth cladding (not shown). The fourth cladding may be located on the suction side 134 at the tip 130, opposite the second cladding 152. While it is contemplated that the fourth cladding may have a surface area having a similar geometry to the second cladding 152, it is also contemplated that the surface area of ​​the fourth cladding is less than or greater than 10% of the surface area of ​​the second cladding 152.

[0091] Alternatively, in a different and non-limiting example, the second cladding 152 may be integrally formed over at least a portion of the tip 130. That is, the second cladding 152 may have portions on both the suction side 134 and the pressure side 132 such that a portion of the second cladding 152 covers a portion of the tip 130. In other words, the second cladding 152 may surround the tip 130 of the composite airfoil 118 with a portion of the second cladding 152 on the pressure side 132 and another portion extending to the suction side 134.

[0092] Optionally, the cladding 120 may also include a root cladding assembly 158 at the root 128 of the composite airfoil 118. The root cladding assembly 158 may include any number of cladding layers in contact with the root portion on the pressure side 132, the suction side 134, or the trailing edge 126.

[0093] As a non-limiting example, adhesive 217 ( Figure 6 ) may be used to couple one or more portions of the cladding 120 to the skin outer surface 190 of the composite airfoil 118. Examples of adhesives may include, but are not limited to, epoxies, phenolics, adhesive films, adhesive tapes, cyanoacrylates, anaerobic adhesives, thermoplastic adhesives, polymer resins, or other thermosetting adhesives.

[0094] Additionally or alternatively, mechanical fasteners may be used to couple one or more portions of the cladding 120 to the skin outer surface 190 of the composite airfoil 118 .

[0095] Alternatively, in various non-limiting examples, one or more portions of the cladding 120 may be applied to the skin outer surface 190 of the composite airfoil 118 using additive manufacturing. That is, the first cladding 150, the second cladding 152, the third cladding 154 ( Figure 6 ), the fourth cladding, the leading edge sheath 156 or the root cladding assembly 158 may be one or more layers additively manufactured on the skin outer surface 190 of the composite airfoil 118 .

[0096] The total surface area of ​​the composite airfoil 118 is defined as the sum of the surface area of ​​the pressure side 132 and the surface area of ​​the suction side 134. The cladding 120 may overlap or cover 0.6% to 70% of the total surface area of ​​the composite airfoil 118. More specifically, the cladding 120 may cover a range of 2% to 60% of the total surface area of ​​the composite airfoil 118. The cladding 120 provides strength or stiffness to the composite airfoil 118, and the range of total surface area covered by the cladding 120 improves strength or stiffness while maintaining the weight benefits provided by the composite airfoil 118.

[0097] Assuming the first cladding 150, the second cladding 152 and the third cladding 154 ( Figure 6 ) may cover a range of 2% to 60% of the pressure side surface area, the suction side surface area, or the total surface area defined by the sum of the pressure side surface area and the suction side surface area. The first cladding 150, the second cladding 152, and the third cladding 154 ( Figure 6 ) improves the strength or stiffness while maintaining the weight benefits provided by the composite airfoil 118.

[0098] As an example, the first cladding 150 is shown as being positioned adjacent to the trailing edge 126 of the composite airfoil 118 when mounted, bonded, or otherwise coupled to the skin outer surface 190 of the composite airfoil 118. As used herein, “adjacent to the trailing edge 126” means that the maximum distance measured from the trailing edge 126 to any portion of the adjacent object is less than 10% of the airfoil length 140. Further, as an example, the first cladding 150 is shown as being located at the trailing edge 126 of the composite airfoil 118. That is, when the first cladding 150 is mounted, bonded, or otherwise coupled to the composite airfoil 118, at least a portion of the first cladding 150 is in contact with the trailing edge 126.

[0099] As an example, the second cladding 152 is shown as being positioned adjacent to the tip 130 of the composite airfoil 118 when mounted, bonded, or otherwise coupled to the skin outer surface 190 of the composite airfoil 118. As used herein, “adjacent to the tip 130” means that the maximum distance measured from the tip 130 to any portion of the adjacent object is less than 10% of the airfoil length 140. Additionally, as an example, the second cladding 152 is shown as being located at the tip 130 of the composite airfoil 118. That is, when the second cladding 152 is mounted, bonded, or otherwise coupled to the composite airfoil 118, at least a portion of the second cladding 152 is in contact with the tip 130.

[0100] Although illustrated as trapezoids and irregular pentagons, portions of the cladding 120 (such as the first cladding 150 or the second cladding 152) may have one or more shapes that are any combination of regular polygons, irregular polygons, or circles with any number of linear or curved portions.

[0101] The cladding 120 may include any number of cladding layers or cladding assemblies on any one or more of the suction side 134, the pressure side 132, the trailing edge 126, the leading edge 124, the root 128, or the tip 130. That is, the cladding 120 may include any material applied to the composite airfoil 118 adjacent the trailing edge 126, the leading edge 124, the root 128, or the tip 130. The term "adjacent the leading edge 124" means that the maximum distance measured from the leading edge 124 to any portion of the adjacent object is less than 10% of the airfoil length 140. Additionally, the term "adjacent the root 128" means that the maximum distance measured from the root 128 to any portion of the adjacent object is less than 10% of the airfoil length 140. Positioning the cladding 120 adjacent the trailing edge 126, the leading edge 124, the root 128, or the tip 130 (by a distance less than 10%) provides the benefit of improved resistance to torsion. The location of the cladding 120 may also provide the benefit of increased strength or resistance to forces caused by impact or rotation while maintaining weight benefits.

[0102] Figure 5 Shown along Figure 4 The line VV intercepts Figure 4 1 is a schematic cross-sectional view of a trailing edge portion of a composite airfoil assembly 110, showing a portion of an interior 180 of the composite airfoil assembly 110. For ease of understanding, the cladding 120 ( Figure 3 ).

[0103] The composite airfoil 118 includes a core, shown as a braided core 182, and a skin, shown as a laminated skin 184, disposed on the braided core 182. The braided core 182 includes a composite structure having a core bulk modulus. The core bulk modulus is less than the first cladding 150 ( Figure 4 ) of the first bulk modulus and the second cladding 152 ( Figure 4 )’s second bulk modulus.

[0104] In one non-limiting example, the braided core 182 can be dry, without additional materials, or alternatively, impregnated with resin and cured. The braided core 182 can be made of a braided structure. Such a braided structure can be a three-dimensional braided structure. More specifically, the braided structure can be axially (Ad), radially (Rd), and circumferentially (Cd) ( Figure 3 ) combination, it should be understood that the braid pattern can be combined with the non-ducted turbine engine 10 ( Figure 1 ) are separately formed and defined so that the weaving pattern is woven in any three mutually orthogonal planes to define a three-dimensional object relative to the planes. In a non-limiting example, the woven structure may include a three-dimensional weaving that includes a plurality of warp fibers 186 and a plurality of weft fibers 188 that can be woven in three directions to form a three-dimensional structure of the woven core 182. The three directions of the warp fibers 186 and the weft fibers 188 can be along the axial direction (Ad), the radial direction (Rd), and the circumferential direction (Cd) ( Figure 3 ) is defined by, or angled relative to. In one non-limiting example, a jacquard loom or 3D loom can be used to create a complex three-dimensional woven structure, which can include interweaving one or more composite materials to form a braided core 182. The braided core 182 can be composed of a composite material such as carbon or carbon fiber, glass or fiberglass, nylon, rayon or other aramid fibers, while other materials such as nickel, titanium or ceramic composites are contemplated in non-limiting examples.

[0105] It is further contemplated that the braided core 182 can be formed into a three-dimensional braided structure having a woven or braided geometry or pattern. For example, the woven or braided geometry or pattern can include a braided pattern that includes three or more interwoven fibers woven in a repeating pattern. In another non-limiting example, the woven geometry can include a group of fibers or strands that are sequentially stacked on top of each other to define the woven geometry. The woven or braided geometry or pattern can be repeated for the entire braided core 182, or only for a portion thereof. Such additional woven geometries can be similar, where the arrangement of the fibers is the same but the orientation is different, or where the arrangement of the fibers is different and the orientation can be similar or dissimilar. The woven geometry or pattern can be formed with a composite material by a jacquard loom or a 3D loom. The three-dimensional braided structure can include a plurality of fibers in three dimensions (such as the axial direction (Ad), the radial direction (Rd), and the circumferential direction (Cd) ( Figure 3 )'s combination).

[0106] The laminate skin 184 can be formed as a set of laminate layers disposed around or around the braided core 182. In a non-limiting example, the laminate skin 184 can be a pre-impregnated, fiber-placed, or dry fiber laminate. In a non-limiting example, such laminate layers forming the laminate skin 184 can be formed by resin transfer molding (RTM), partial RTM, same qualified resin transfer molding (SQRTM), or out of autoclave. The braided core 182 can be made of a different material than the laminate skin 184.

[0107] The laminate skin 184 can include a skin bulk modulus. The skin bulk modulus can be different from the core bulk modulus. In one example, the skin bulk modulus can be greater than the core bulk modulus.

[0108] The laminate skin 184 may include a skin outer surface 190 and a skin inner surface 192. The skin inner surface 192 may at least partially contact at least a portion of an exterior 194 of the braided core 182. In other words, the laminate skin 184 may be applied to at least a portion of an exterior 194 of the braided core 182.

[0109] The skin thickness 196 may be measured from the skin outer surface 190 to the skin inner surface 192. Although shown as uniform, the skin thickness 196 may be in the axial direction (Ad), the radial direction (Rd) ( Figure 3 ) or circumferential direction (Cd)( Figure 3 ) changes.

[0110] The first portion 190a of the skin outer surface 190 defines a first portion 142a of the airfoil outer surface 142 of the composite airfoil assembly 110. The second portion 190b of the skin outer surface 190 is located below the second portion 142b of the airfoil outer surface 142.

[0111] The intersection 198 is defined at a location where the first portion 190a of the epidermal outer surface 190 changes, deflects, or otherwise intersects with the second portion 190b of the epidermal outer surface 190. As shown as an example, the first portion 190a of the epidermal outer surface 190 and the second portion 190b of the epidermal outer surface 190 can be different portions of the same epidermal outer surface 190. That is, the intersection 198 separates the first portion 190a and the second portion 190b of the epidermal outer surface 190. Alternatively, in different and non-limiting examples, the first portion 190a of the epidermal outer surface 190 and the second portion 190b of the epidermal outer surface 190 can be different epidermis that can be adjacent or partially overlapped at the intersection 198.

[0112] The notch or recess 200 is defined by the second portion 190b of the skin outer surface 190 located below the airfoil outer surface 142. The recess depth or recess distance 202 is measured between the second portion 142b of the airfoil outer surface 142 and the second portion 190b of the skin outer surface 190. As shown, the recess distance 202 can be perpendicular to the radial direction (Rd) ( Figure 4 ) is measured in an axial plane of the axis of the axis. The recess distance 202 may be in a range of 0.3 mm to 1.0 mm. More specifically, the recess distance 202 may be in a range of 0.5 mm to 0.7 mm.

[0113] Additionally or alternatively, the recess distance 202 may be within the axial airfoil length 144 ( Figure 4 ). More specifically, the recess distance 202 may be within a range of 0.001% to 5% of the axial airfoil length 144 ( Figure 4 ) is in the range of 0.01% to 1.5%.

[0114] The boundary region 206 can be defined as a location where the depression distance 202 varies along the axial direction (Ad). The boundary region 206 has a boundary length 208 measured axially from the intersection point 198 to a point 210 along the second portion 190b of the outer surface 190 of the epidermis. The point 210 can be a location beyond which the depression distance 202 varies less than 5% per centimeter along the axial direction (Ad). As shown, as an example, the boundary length 208 can be different from the suction side 134 to the pressure side 132. Alternatively, when compared to the depression distance 202 or the boundary region 206 on the pressure side 132, the depression distance 202 on the suction side 134, the boundary region 206, or both the depression distance 202 and the boundary region 206 can be within 5%.

[0115] Figure 6 yes Figure 5 1 is a schematic cross-sectional view of a composite airfoil assembly 110 having cladding 120 , and more specifically, a first cladding 150 and a third cladding 154 .

[0116] The first cladding 150 is coupled to the second portion 190b of the skin outer surface 190. The cladding thickness 212 may be measured from the inner surface 214 to the outer surface 216 of the first cladding 150. The cladding axial length 218 may be measured from the axial forward edge 151c to the axial rearward edge 151d.

[0117] The inner surface 214 of the first cladding layer 150 may cover a portion of the second portion 190b of the skin outer surface 190. An adhesive 217 may be located on the inner surface 214, the second portion 190b, or both the second portion 190b and the inner surface 214 to couple the first cladding layer 150 to the composite airfoil 118. The adhesive thickness 219 may be in the range of 0.01% to 50% of the recess distance 202. It is contemplated that the adhesive thickness 219 may be in the axial direction (Ad) or the radial direction (Rd) ( Figure 4 ) changes.

[0118] The outer surface 216 of the first cladding layer 150 may define a portion of the second portion 142b of the composite airfoil assembly 110. It is contemplated that the cladding thickness 212 of the first cladding layer 150 is in a range of 40% to 100% of the recessed distance 202. More specifically, the cladding thickness 212 of the first cladding layer 150 is in a range of 50% to 90% of the recessed distance 202. Still more specifically, the cladding thickness 212 of the first cladding layer 150 is in a range of 60% to 85% of the recessed distance 202. This improves the aerodynamics of the composite airfoil assembly 110 because the first cladding layer 150 does not extend or protrude into the airflow as indicated by arrows 204.

[0119] It is contemplated that the sum of adhesive thickness 219 and cladding thickness 212 is within 5% of recess distance 202. More specifically, the sum of adhesive thickness 219 and cladding thickness is within 1% of recess distance 202. Still more specifically, the sum of adhesive thickness 219 and cladding thickness 212 differs from recess distance 202 by less than 0.5%.

[0120] Alternatively, in a different and non-limiting example, the difference between the cladding thickness 212 and the recess distance 202 of the first cladding 150 may be greater than 10%.

[0121] Similar to the first cladding 150, the third cladding 154 has a cladding thickness 220 and a cladding axial length 222. The cladding thickness 220 may be measured from an inner surface 224 to an outer surface 226 of the third cladding 154. The cladding axial length 222 of the third cladding 154 may be measured from an axial forward edge 155a to an axial backward edge 155b.

[0122] Although shown as having similar cladding thicknesses 212 , 220 , it is contemplated that the cladding thickness 212 of the first cladding 150 may be different than the cladding thickness 220 of the third cladding 154 .

[0123] The cladding thickness 212, 220 of the first cladding 150 or the third cladding 154 may be in the range of 0.076 mm to 1.016 mm. More specifically, the first cladding thickness 212 or the second cladding thickness 220 may be in the range of 0.15 mm to 0.51 mm.

[0124] Additionally or alternatively, the cladding thicknesses 212, 220 may be greater than or equal to the axial airfoil length 144 ( Figure 4 ). More specifically, the cladding thickness 212, 220 may be within a range of 0.001% to 5% of the axial airfoil length 144 ( Figure 4 ) is in the range of 0.01% to 1.5% of the cladding thickness. The range of cladding thickness 212, 220 improves strength or stiffness while maintaining weight benefits.

[0125] The core thickness 228 may be measured across the braided core 182 from the pressure side 132 to the suction side 134. That is, the core thickness 228 may be measured across the braided core 182 from the pressure side 132 to the suction side 134 in a direction perpendicular to the chordwise direction and the spanwise direction, wherein the chordwise direction is from the leading edge 124 ( Figure 3 ) to the trailing edge 126, and the spanwise direction is from the root 128 ( Figure 3 ) to the tip 130( Figure 3 ) is defined. The core thickness 228 can be in the range of 2.54 mm to 63.5 mm. More specifically, the core thickness 228 can be in the range of 3.17 mm to 12.7 mm.

[0126] Additionally or alternatively, the core thickness 228 may be greater than the axial airfoil length 144 ( Figure 4 ). More specifically, the core thickness 228 may be within a range of 0.05% to 50% of the axial airfoil length 144 ( Figure 4 ) is in the range of 0.2% to 20% of the core thickness. The range of core thickness 228 provides a desired aerodynamic profile while maintaining weight benefits.

[0127] The cladding thickness 212, 220 may be in a range of 0.6% to 25% of the core thickness 228. More specifically, the cladding thickness 212, 220 may be in a range of 4% to 16% of the core thickness 228. The range provided for the cladding thickness 212, 220 improves the strength or stiffness of the composite airfoil assembly 110 when compared to the core thickness 228 while maintaining weight benefits.

[0128] Although shown as having a generally uniform cladding thickness 212 , 220 , the cladding thickness 212 , 220 may vary in any direction (eg, in an axial direction (Ad), a radial direction (Rd), a spanwise direction, or a chordwise direction).

[0129] The cladding axial lengths 218, 222 may be within the axial airfoil length 144 ( Figure 4 ). More specifically, the cladding axial lengths 218, 222 may be within a range of 1% to 10% of the axial airfoil length 144 ( Figure 4). Providing a range for the cladding axial lengths 218, 222 improves strength or stiffness while maintaining weight benefits.

[0130] Although shown as having different cladding axial lengths 218 , 222 , it is contemplated that the cladding axial length 218 of the first cladding 150 may be within 5% or less of the cladding axial length 222 of the third cladding 154 .

[0131] When the first and third envelopes 150, 154 are adjacent the trailing edge 126 of the composite airfoil 118, as an example, the axially aft edge 151 d of the first and third envelopes 150, 155 b of the third envelope 154 are shown as being axially forward of the trailing edge 126. However, it is contemplated that the axially aft edges 151 d, 155 b may extend past or be aligned with at least a portion of the trailing edge 126.

[0132] Optionally, the trailing edge cladding 230 may be coupled to or defined by a portion of the first cladding 150 or the third cladding 154. The trailing edge cladding 230 covers at least a portion of the trailing edge 126 of the composite airfoil 118. The trailing edge cladding 230 may include one or more portions or sections.

[0133] Although shown as being aligned with intersection point 198, point 210, or boundary region 206 ( Figure 5 ), but the first portion 142a of the airfoil outer surface 142 and the second portion 142b of the airfoil outer surface 142 are considered flush, so that the second portion 142b of the airfoil outer surface 142 does not protrude, extend outward, or include a positive bulge. In other words, the second portion 142b of the airfoil outer surface 142 follows the aerodynamic line or is exactly flush with the first portion 142a of the airfoil outer surface 142. Since the first portion 142a and the second portion 142b are connected, the first portion 142a and the second portion 142b can be considered continuous.

[0134] In various and non-limiting examples, it is contemplated that the first cladding 150 or the third cladding 154 may contact the intersection 198 or the point 210, or extend to the boundary region 206 ( Figure 5 That is, the intersection point 198, point 210 or boundary region 206 and the axial forward edges 151c, 153c ( Figure 4 ) can be minimized or eliminated by extending the axial forward edges 151c, 153c into the boundary region 206.

[0135] In another different and non-limiting example, one or more of the coating, filler material, or portion of the laminate skin 184 may be geometrically shaped or applied to fill a gap or a portion of a gap between the intersection 198, point 210, or boundary region 206 and the axial forward edges 151c, 153c of the first and third cladding layers 150, 154. That is, the first portion 142a and the second portion 142b may be in contact or continuous. As a non-limiting example, the filler material or coating may define a portion of the first portion 142a or the second portion 142b of the airfoil outer surface 142 such that there is no gap or space between the first portion 142a and the second portion 142b.

[0136] The non-limiting example shows that the first portion 142a and the second portion 142b are flush and continuous.

[0137] Figure 7 yes Figure 6 A variation of the schematic cross-sectional view of Figure 7 Composite airfoil assembly 310 is shown having a first cladding layer 350 and a third cladding layer 354. Composite airfoil assembly 310 is similar to composite airfoil assembly 110 and, therefore, like parts will be identified with like numerals increased by two hundred (200), with the understanding that the description of like parts of composite airfoil assembly 110 applies to composite airfoil assembly 310 unless otherwise noted.

[0138] The skin 384 may be coupled to the exterior of the braided core 382. The skin 384 may include a first layer 384a and a second layer 384b. The first layer 384a may be a laminated skin. The first layer 384a may be applied before or after curing the braided core 382.

[0139] The second layer 384b may be a coating applied directly onto the first layer (laminate skin) 384a. The second layer 384b may be applied before or after curing the braided core 382 or the first layer 384a.

[0140] However, it is contemplated that an adhesive 217 is disposed between the first layer 384a and the second layer 384b, and more specifically, an intermediate adhesive layer ( Figure 6 ). Although shown as covering a portion of the first layer 384a, the second layer 384b can cover the entire first layer 384a. It is further contemplated that the second layer 384b can be disposed on the laminate skin 384, wherein portions of the laminate skin 384 are not covered by the cladding 320. It is also further contemplated that the second layer 384b can be applied to one or more portions of the cladding 320. The second layer 384b can include a second layer bulk modulus that is different from or the same as one or more of the core bulk modulus or the first layer bulk modulus. It is contemplated that the second layer bulk modulus is less than the bulk modulus of the cladding 320.

[0141] In addition, the second layer 384b can be an environmental barrier coating, for example, which can be used to resist oxidation or corrosion. In another example, the second layer 384b can be a thermal barrier coating that at least partially thermally isolates the braided core 382 and the first layer 384a. Additional non-limiting examples of coatings can include anti-icing coatings, such as polyurethane, ice-repelling materials, ultraviolet radiation coatings, or oil barrier coatings, such as polyethylene or polypropylene. In an additional non-limiting example, the second layer 384b can be formed as a polypropylene base layer and a polyurethane layer disposed on the polypropylene base layer. It is further contemplated that an external coating layer (not shown) can be provided on the outside of the second layer 384b, wherein such a coating layer can provide radiation protection, such as ultraviolet radiation.

[0142] If a portion of the first layer 384a is not covered by the second layer 384b, the epidermis outer surface 390 may be the outer surface of the second layer 384b or the outer surface of the first layer 384a. The epidermis inner surface 392 may be the inner surface of the first layer 384a.

[0143] The skin thickness 396 can be measured from the skin outer surface 390 to the skin inner surface 392. As shown as an example, the skin thickness 396 can vary in the axial direction (Ad). It is also contemplated that the skin thickness 396 can vary in the radial direction (Rd) ( Figure 3 ), axial direction (Ad) or circumferential direction (Cd) ( Figure 3 ) in one or more of the above changes.

[0144] The first portion 390a of the skin outer surface 390 defines a first portion 342a of the airfoil outer surface 342 of the composite airfoil assembly 310. The second portion 390b of the skin outer surface 390 is located below the second portion 342b of the airfoil outer surface 342.

[0145] The intersection point 398 is defined at a location where the first portion 390a of the epidermal outer surface 390 changes, deflects, or otherwise intersects the second portion 390b of the epidermal outer surface 390 .

[0146] The boundary region 406 may be defined as a location along the axial direction (Ad) where the recess distance 402 varies. The recess distance 402 may be measured from the second portion 390b of the skin outer surface 390 to the second portion 342b of the airfoil outer surface 342 .

[0147] The boundary length 408 of the boundary region 406 is measured axially from the intersection point 398 to a point 410 along the second portion 390b of the epidermal outer surface 390. The point 410 may be a location beyond which the depression distance 402 varies less than 5% per centimeter along the axial direction (Ad).

[0148] The boundary length 408 may be in a range of 1% to 300% of the core thickness 228. More specifically, the boundary length 408 may be in a range of 30% to 120% of the core thickness 228.

[0149] The boundary length 408 may be within the axial airfoil length 144 ( Figure 4 ). More specifically, the boundary length 408 may be in the range of 0.5% to 20% of the axial airfoil length 144 .

[0150] The first cladding 350 and the third cladding 354 are coupled to the second portion 390b of the skin outer surface 390. The cladding thickness 412 can be measured from the inner surface 415 to the outer surface 417 of the first cladding 350 or the third cladding 354. The cladding length 418 can be the axial length of the cladding measured from the first end or axial forward edge 436 of the first cladding 350 or the third cladding 354 to the respective second end or axial backward edge 438 of the first cladding 350 or the third cladding 354. Although shown as being equal, the cladding length 418 of the first cladding 350 can be different from the cladding thickness 418 of the third cladding 354. In various non-limiting examples, it is further contemplated that the cladding length can be the radial length measured from the radial inside to the radial outside of the cladding 320 in the radial direction (Rd).

[0151] Adhesive 217( Figure 6 ) may be located on the inner surface 415 , the second portion 390 b , or both the second portion 390 b and the inner surface 415 to couple the first cladding 350 or the third cladding 354 to the composite airfoil 318 .

[0152] The outer surface 417 of the first cladding 350 or the third cladding 354 may define a portion of the second portion 342b of the composite airfoil assembly 310. It is contemplated that the cladding thickness 412 of the first cladding 350 or the third cladding 354 is in a range of 40% to 100% of the recessed distance 402. More specifically, the cladding thickness 412 of the first cladding 350 or the third cladding 354 is in a range of 50% to 90% of the recessed distance 402. Still more specifically, the cladding thickness 412 of the first cladding 350 or the third cladding 354 is in a range of 60% to 85% of the recessed distance 402.

[0153] The cladding thickness 412 of the first cladding 350 or the third cladding 354 improves the aerodynamics of the composite airfoil assembly 310 because the first cladding 350 and / or the third cladding 354 do not extend or protrude into the airflow as indicated by arrows 404 .

[0154] The first cladding 350, the third cladding 354, or the first cladding 350 and the third cladding 354 include a tapered portion 440. As shown, by way of example, the tapered portion 440 is defined by varying the cladding thickness 412 in at least the axial direction (Ad). In various non-limiting examples, it is contemplated that the tapered portion 440 may be defined by varying the cladding thickness 412 in at least the radial direction (Rd) ( Figure 4 ) is limited by changing the cladding thickness.

[0155] The tapered portion 440 may extend into or be located at the boundary region 406. That is, the boundary region 406 receives at least a portion of the tapered portion 440 of the cladding 320.

[0156] The tapered length 442 of the tapered portion 440 is measured between the first boundary 444 and the second boundary 446 of the tapered portion 440. The tapered length 442 is in the range of 1% to 50% of the cladding length 418. More specifically, the tapered length 442 is in the range of 15% to 30% of the cladding length 418.

[0157] As shown, as an example, the taper length 442 can be within 5% or less of the boundary length 408. More specifically, the taper length 442 can be within 1% or less of the boundary length 408. It is assumed that the first portion 342a of the airfoil outer surface 342 and the second portion 342b of the airfoil outer surface are flush. That is, the first portion 390a of the skin outer surface 390 is adjacent to the cladding 320 so that the first portion 342a and the second portion 342b form a continuous airfoil outer surface 342. As used herein, the term "continuous" means without breaks, gaps, or gaps. In other words, the second portion 390b of the skin outer surface 390 defines a recess having a recessed distance 402, wherein the recess has a geometry that is complementary to the cladding 320.

[0158] In a different and non-limiting example, there may be a gap between the first portion 342a and the second portion 342b that, when filled with one or more materials, forms a continuous airfoil outer surface 342 .

[0159] Figure 8 yes Figure 7 FIG. 4 is an enlarged schematic cross-sectional view of a first cladding 350 of FIG. The first cladding 350 extends from an axial forward edge 436 to an axial backward edge 438 and includes a tapered portion 440. The tapered portion 440 includes a tapered length 442 measured from a first boundary 444 to a second boundary 446 of the tapered portion 440. As a non-limiting example, the first boundary 444 is located at the axial forward edge 436.

[0160] The first cladding thickness 412a is measured from the inner surface 415 to the outer surface 417 at the first boundary 444. The second cladding thickness 412b is measured from the inner surface 415 to the outer surface 417 at the second boundary 446. The third cladding thickness 412c may be measured from the inner surface 415 to the outer surface 417 between the second boundary 446 and the axial rearward edge 438.

[0161] The first cladding thickness 412a may be in a range of 0.05% to 4% of the tapered length 442. More specifically, the first cladding thickness 412a may be in a range of 0.1% to 2% of the tapered length 442.

[0162] The first cladding thickness 412a may be in the range of 5% to 60% of the second cladding thickness 412b. More specifically, the first cladding thickness 412a may be in the range of 5% to 40% of the second cladding thickness 412b.

[0163] The second cladding thickness 412b may be in a range of 0.5% to 10% of the tapered length 442. More specifically, the second cladding thickness 412b may be in a range of 1% to 8% of the tapered length 442.

[0164] The third cladding thickness 412c may be within 5% or less of the second cladding thickness 412b. Alternatively, in a different and non-limiting example, the first cladding 350 may include more than one tapered portion where the third cladding thickness 412c differs from the second cladding thickness 412b by more than 5%.

[0165] The ranges of first cladding thickness 412a, second cladding thickness 412b, and third cladding thickness 412c provide weight benefits while strengthening or stiffening composite airfoil 318 ( Figure 7 ).

[0166] Although shown at the axial forward edge 436 of the first cladding 350 and the third cladding 354, it is contemplated that other portions of the cladding 320 may include tapered portions. As a non-limiting example, the second cladding 152 ( Figure 4 ) or one or more portions of the axial forward edge 153c of the fourth cladding may include a tapered portion.

[0167] In operation, the cladding 120, 320 provides improved resistance to torsion, impact, or other forces caused by impact or rotation, while maintaining weight benefits. The recess 200 provides improved aerodynamics and efficiency, receiving at least some of the cladding 120, 320, which reduces or eliminates positive protrusions or bulges that would impede airflow along the pressure side 132 or suction side 134 (as shown by arrows 204). The tapered portion 440 transfers loads at the bonding interface or boundary region 206, 406.

[0168] The tapered portion 440 prevents or reduces large load discontinuities that may occur due to sudden changes caused by impact or rotation. That is, the smooth angled transition of the cladding between the first boundary 444 and the second boundary 446 helps distribute the force and prevents or reduces the force from being localized in a small surface area, which reduces the material stress at the boundary regions 206, 406.

[0169] Benefits of aspects of the present invention include a recess in a composite airfoil such that when a cladding is received by the recess, aerodynamics are improved. Aerodynamics can be improved by removing the positive bulge caused by coupling the cladding to the composite airfoil without the recess. In addition, a composite airfoil having a recess and a cladding as a reinforcement provides weight benefits and reinforcement benefits.

[0170] The tapered section of the cladding improves load transfer at the bond interface. The taper avoids large load discontinuities that can occur due to sudden changes caused by impact or rotation.

[0171] There are weight benefits to using a composite airfoil with a cladding having a greater bulk modulus than a braided core and laminate skin. When a cladding is added to a composite airfoil, resistance to torsion, impact, or other forces caused by impact or rotation may be improved while maintaining weight benefits.

[0172] To the extent not yet described, the different features and structures of the various embodiments may be used in combination as desired, or may be interchanged with one another. Not showing a feature in all embodiments does not mean that it is interpreted as not being able to be shown in this way, but rather is done for the sake of brevity of description. Therefore, the various features of the different embodiments may be mixed and matched as desired to form new embodiments, whether or not the new embodiments are explicitly described. All combinations or permutations of the features described herein are covered by the present disclosure.

[0173] This written description uses examples to disclose the present disclosure, including the best mode, and also to enable any person 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. These other examples are intended to fall 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.

[0174] Further aspects are provided by the subject matter of the following clauses:

[0175] A composite airfoil assembly for a turbine engine, the composite airfoil assembly having an airfoil outer surface, the airfoil outer surface defining opposing pressure and suction sides, the pressure and suction sides extending between a leading edge and a trailing edge, the composite airfoil assembly comprising: a braided core, the braided core comprising a composite structure defining a core exterior; a skin applied to at least a portion of the core exterior, the skin having a skin outer surface, wherein a first portion of the skin outer surface defines a first portion of the airfoil outer surface, and a second portion of the skin outer surface is located below the airfoil outer surface to define a recess; and a cladding positioned adjacent the trailing edge, the leading edge, the root, or the tip, the cladding having at least a portion located within the recess and having a cladding outer surface defining a second portion of the airfoil outer surface, wherein the first portion of the airfoil outer surface and the second portion of the airfoil outer surface are flush.

[0176] A composite airfoil assembly for a gas turbine engine, the composite airfoil assembly having an airfoil outer surface, the airfoil outer surface defining opposite pressure and suction sides, the pressure and suction sides extending in an axial direction between a leading edge and a trailing edge, and extending in a radial direction from a root to a tip, the composite airfoil assembly comprising: a composite airfoil, the composite airfoil comprising: a braided core, the braided core comprising a composite structure defining a core exterior; and a skin, the skin being applied to at least a portion of the core exterior and having a skin outer surface; and a cladding, the cladding being coupled to the skin outer surface, wherein at least one edge or portion of the cladding is positioned adjacent to at least one of the trailing edge, the root, or the tip, the cladding having a tapered portion on the pressure side or the suction side.

[0177] A composite airfoil assembly according to any preceding clause, further comprising a recessed distance measured between the second portion of the airfoil exterior surface and the second portion of the skin exterior surface, wherein the recessed distance is in the range of 0.001% to 5% of an axial airfoil length measured from the leading edge to the trailing edge.

[0178] A composite airfoil assembly according to any preceding clause, wherein the cladding comprises a cladding thickness measured from the cladding outer surface to the cladding inner surface.

[0179] A composite airfoil assembly as claimed in any preceding clause, wherein the cladding thickness is in the range of 40% to 100% of the recess distance.

[0180] A composite airfoil assembly as recited in any preceding clause, further comprising an adhesive located between the cladding inner surface and the second portion of the skin outer surface, wherein the adhesive couples the cladding to the second portion of the skin outer surface.

[0181] A composite airfoil assembly as claimed in any preceding clause, wherein the adhesive comprises an adhesive thickness in the range of 0.01% to 50% of the recessed distance.

[0182] A composite airfoil assembly as claimed in any preceding clause, wherein the sum of the adhesive thickness and the cladding thickness is within 5% of the recess distance.

[0183] A composite airfoil assembly as claimed in any preceding clause wherein the sum of the adhesive thickness and the cladding thickness is within 1% of the recess distance.

[0184] A composite airfoil assembly according to any preceding clause, further comprising a recessed distance measured between the second portion of the airfoil exterior surface and the second portion of the skin exterior surface, wherein the recessed distance is in the range of 0.01% to 1.5% of an axial airfoil length measured from the leading edge to the trailing edge.

[0185] A composite airfoil assembly according to any preceding clause, wherein the first portion of the skin exterior surface and the second portion of the skin exterior surface intersect to form a boundary region having a varying recess distance measured between the second portion of the airfoil exterior surface and the second portion of the skin exterior surface.

[0186] A composite airfoil assembly according to any preceding clause, wherein the border region has a border length in the range of 0.001% to 40% of an axial airfoil length measured from the leading edge to the trailing edge.

[0187] A composite airfoil assembly according to any preceding clause, wherein the border region has a border length in the range of 0.5% to 20% of an axial airfoil length measured from the leading edge to the trailing edge.

[0188] A composite airfoil assembly according to any preceding clause, wherein the braided core comprises a core thickness measured across the braided core from the pressure side to the suction side in a direction perpendicular to the chord-wise direction and the span-wise direction, and wherein the core thickness is in the range of 0.05% to 50% of an axial airfoil length measured from the leading edge to the trailing edge.

[0189] A composite airfoil assembly as described in any preceding clause, wherein the cladding comprises a cladding thickness, wherein the cladding thickness is in the range of 0.6% to 25% of the core thickness.

[0190] A composite airfoil assembly as claimed in any preceding clause, wherein the boundary region has a boundary length in the range of 1% to 300% of the core thickness.

[0191] A composite airfoil assembly as described in any preceding clause, wherein the recess comprises a geometry complementary to the cladding.

[0192] A composite airfoil assembly as described in any preceding clause, wherein the braided core comprises a three-dimensional braid pattern.

[0193] A composite airfoil assembly according to any preceding clause, wherein the cladding comprises a cladding thickness measured from a cladding inner surface to a cladding outer surface, wherein the tapered portion is defined by varying the cladding thickness in at least one of the axial direction or the radial direction.

[0194] A composite airfoil assembly according to any preceding clause, wherein the cladding includes a cladding length measured from a first end to a second end of the cladding, and the tapered portion extends from a first boundary to a second boundary, defining a tapered length, wherein the tapered length is in the range of 1% to 50% of the cladding length.

[0195] A composite airfoil assembly as described in any preceding clause, wherein the first boundary is at the first end and the taper length is in the range of 15% to 30% of the cladding length.

[0196] A composite airfoil assembly as in any preceding clause, wherein the tapered portion extends from a first boundary to a second boundary, defining a tapered length, wherein the cladding thickness at the first boundary is in the range of 0.05% to 2% of the tapered length.

[0197] A composite airfoil assembly as in any preceding clause, wherein the tapered portion extends from a first boundary to a second boundary, defining a tapered length, wherein the cladding thickness at the second boundary is in the range of 0.5% to 10% of the tapered length.

[0198] A composite airfoil assembly as recited in any preceding clause, wherein the tapered portion extends from a first boundary to a second boundary, wherein the cladding thickness at the first boundary is in the range of 5% to 60% of the cladding thickness at the second boundary.

[0199] A composite airfoil assembly as claimed in any preceding clause, the skin having a first portion of the skin outer surface defining a first portion of the airfoil outer surface, and a second portion of the skin outer surface located below the airfoil outer surface to define a recess.

[0200] A composite airfoil assembly as described in any preceding clause, the cladding having at least a portion located within the recess and having a cladding outer surface defining a second portion of the airfoil outer surface, wherein the first portion of the airfoil outer surface and the second portion of the airfoil outer surface are flush.

[0201] A composite airfoil assembly as described in any preceding clause, wherein the recess comprises a geometry complementary to the cladding.

[0202] A composite airfoil assembly according to any preceding clause, wherein the first portion of the skin outer surface and the second portion of the skin outer surface intersect to form a boundary region, wherein the boundary region receives the tapered portion of the cladding.

[0203] A composite airfoil assembly as described in any preceding clause, wherein the three-dimensional weave pattern comprises a plaited pattern.

[0204] A composite airfoil assembly as described in any preceding clause wherein the skin is a laminate skin.

[0205] A composite airfoil assembly as described in any preceding clause wherein the braided core is made of a different material than the skin.

[0206] A composite airfoil assembly as described in any preceding clause, wherein the tapered portion is located axially between the leading edge and the trailing edge.

[0207] A composite airfoil assembly as described in any preceding clause, wherein the tapered portion is located at an axially inner edge.

[0208] A composite airfoil assembly as claimed in any preceding clause, wherein the boundary length is in the range of 30% to 120% of the core thickness.

[0209] A composite airfoil assembly according to any preceding clause, wherein the cladding comprises a first cladding and a second cladding, wherein the first cladding is located at the trailing edge and the second cladding is located at the tip end.

[0210] A composite airfoil assembly according to any preceding clause, wherein the cladding further comprises a third cladding located at the trailing edge opposite the first cladding.

[0211] A composite airfoil assembly according to any preceding clause, wherein the cladding further comprises a fourth cladding located at the tip end opposite the second cladding.

[0212] A composite airfoil assembly according to any preceding clause, wherein the boundary length of the boundary region is measured axially from a point of intersection to a point along the second portion of the skin outer surface, wherein the point is an axial location beyond which the depression distance varies along the axial direction by less than 5% per centimeter.

[0213] A turbine engine comprising: a fan section, a compressor section, a combustor section, and a turbine section, the fan section, the compressor section, the combustor section, and the turbine section being in a serial flow arrangement and defining an engine centerline; and a composite airfoil assembly having an airfoil outer surface, the airfoil outer surface defining opposing pressure and suction sides, the pressure and suction sides extending between a leading edge and a trailing edge to define a chordwise direction and extending from a root to a tip to define a spanwise direction, the composite airfoil assembly comprising: a core comprising a composite structure defining a core exterior; and a skin , the skin being applied to at least a portion of the exterior of the core, the skin having a skin outer surface, wherein a first portion of the skin outer surface defines a first portion of the airfoil outer surface, and a second portion of the skin outer surface is located below the airfoil outer surface to define a recess; and a cladding positioned adjacent the trailing edge, the leading edge, the root, or the tip, the cladding having at least a portion located within the recess and having a cladding outer surface defining a second portion of the airfoil outer surface, wherein the first portion of the airfoil outer surface and the second portion of the airfoil outer surface are flush.

[0214] A turbine engine comprises: a fan section, a compressor section, a combustor section and a turbine section, wherein the fan section, the compressor section, the combustor section and the turbine section are in a serial flow arrangement and define an engine centerline; and a composite airfoil assembly, wherein the composite airfoil assembly is used for a gas turbine engine, wherein the composite airfoil assembly has an airfoil outer surface, wherein the airfoil outer surface defines opposite pressure and suction sides, wherein the pressure and suction sides extend between a leading edge and a trailing edge to define a chordwise direction and extend from a root to a tip to define a spanwise direction, wherein the composite airfoil assembly comprises: a core, wherein the core comprises a composite structure defining a core exterior; a skin, wherein the skin is applied to at least a portion of the core exterior; and a cladding, wherein the cladding is positioned adjacent to the trailing edge, the root or the tip, wherein the cladding has a tapered portion.

[0215] A turbine engine according to any preceding clause, further comprising a recessed distance measured between the second portion of the airfoil outer surface and the second portion of the skin outer surface, wherein the recessed distance is in the range of 0.001% to 5% of the axial airfoil length measured from the leading edge to the trailing edge.

[0216] A turbine engine according to any preceding clause, further comprising a recessed distance measured between the second portion of the airfoil outer surface and the second portion of the skin outer surface, wherein the recessed distance is in the range of 0.01% to 1.5% of the axial airfoil length measured from the leading edge to the trailing edge.

[0217] The turbine engine of any preceding clause, further comprising an adhesive located between an inner surface of the cladding and the second portion of the outer surface of the skin, wherein the adhesive couples the cladding to the second portion of the outer surface of the skin.

[0218] A turbine engine according to any preceding clause, wherein the adhesive comprises an adhesive thickness and the cladding comprises a cladding thickness measured from the cladding outer surface to the cladding inner surface, wherein the sum of the adhesive thickness and the cladding thickness is within 5% of the recess distance.

[0219] A turbine engine as described in any preceding clause, wherein the cladding includes a cladding thickness measured from a cladding inner surface to a cladding outer surface, wherein the tapered portion is defined by varying the cladding thickness in at least one of an axial direction or a radial direction.

[0220] A turbine engine according to any preceding clause, wherein the cladding includes a cladding length measured from a first end to a second end of the cladding, and the tapered portion extends from a first boundary to a second boundary, defining a tapered length, wherein the tapered length is in the range of 1% to 50% of the cladding length.

[0221] A turbine engine as claimed in any preceding clause, wherein the first boundary is at the first end and the taper length is in the range of 15% to 30% of the cladding length.

[0222] A gas turbine engine as in any preceding clause, wherein the tapered portion extends from a first boundary to a second boundary, defining a tapered length, wherein the cladding thickness at the first boundary is in the range of 0.05% to 2% of the tapered length.

Claims

1. A composite airfoil assembly for a turbine engine, the composite airfoil assembly having an airfoil outer surface defining opposing pressure and suction sides, the pressure and suction sides extending in an axial direction between a leading edge and a trailing edge and in a radial direction from a root to a tip, characterized in that The composite airfoil assembly comprises: A composite airfoil, the composite airfoil comprising: a braided core comprising a composite structure defining a core exterior; and a skin applied to at least a portion of the core exterior and having a skin outer surface; and a cladding coupled to the skin outer surface, wherein at least one edge or portion of the cladding is positioned adjacent to at least one of the trailing edge, the root, or the tip, the cladding having a tapered portion on the pressure side or the suction side.

2. The composite airfoil assembly according to claim 1, wherein: in, The cladding includes a cladding thickness measured from an inner surface of the cladding to an outer surface of the cladding, wherein the tapered portion is defined by varying the cladding thickness in at least one of the axial direction or the radial direction.

3. The composite airfoil assembly according to claim 2, wherein: in, The cladding includes a cladding length measured from a first end to a second end of the cladding, and the tapered portion extends from a first boundary to a second boundary defining a tapered length, wherein the tapered length is in a range of 1% to 50% of the cladding length.

4. The composite airfoil assembly according to claim 3, wherein: in, The first boundary is at the first end, and the taper length is in the range of 15% to 30% of the cladding length.

5. The composite airfoil assembly according to claim 2, wherein: in, The tapered portion extends from a first boundary to a second boundary, defining a tapered length, wherein a thickness of the cladding at the first boundary is in a range of 0.05% to 2% of the tapered length.

6. The composite airfoil assembly of claim 2, wherein: in, The tapered portion extends from a first boundary to a second boundary, defining a tapered length, wherein a thickness of the cladding at the second boundary is in a range of 0.5% to 10% of the tapered length.

7. The composite airfoil assembly of claim 2, wherein: in, The tapered portion extends from a first boundary to a second boundary, wherein a thickness of the cladding at the first boundary is in a range of 5% to 60% of a thickness of the cladding at the second boundary.

8. The composite airfoil assembly of claim 1, wherein: The skin has a first portion of the skin outer surface defining a first portion of the airfoil outer surface, and a second portion of the skin outer surface located below the airfoil outer surface to define a recess.

9. The composite airfoil assembly of claim 8, wherein: The cladding has at least a portion located within the recess and has a cladding outer surface defining a second portion of the airfoil outer surface, wherein the first portion of the airfoil outer surface and the second portion of the airfoil outer surface are flush.

10. The composite airfoil assembly of claim 8, wherein: in, The recess comprises a geometry complementary to the cladding.

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