Composite airfoil assembly and method for forming composite airfoil assembly

Abstract

A composite airfoil assembly for a gas turbine engine. The composite airfoil assembly includes a composite airfoil defined by a core and a skin. The composite airfoil assembly also includes a cladding. The core defines a core exterior, wherein the skin is applied to at least a portion of the core exterior. The cladding is prepared prior to being attached or adhered to the composite airfoil.

Description

Technical Field

[0001] This disclosure generally relates to turbine engines, and more specifically, to composite airfoil assemblies for turbine engines. Background Technology

[0002] A turbofan engine typically comprises an engine core with a compressor section, a combustor section, and a turbine section arranged in a series flow configuration. In a turbofan or bypass configuration, the fan section may be located upstream of the compressor section. The compressor section compresses air directed to the combustor section, where the air is mixed with fuel, and the mixture is then ignited to generate hot combustion gases. The combustion gases are directed to the turbine section, where energy is extracted from the combustion gases to power the compressor section and to generate useful work to propel the aircraft in flight or power loads such as generators.

[0003] With the advent of composite materials, they have been used to manufacture components for gas turbine engines. Composite materials typically consist of a fiber-reinforced matrix and exhibit a high strength-to-weight ratio. Due to their high strength-to-weight ratio and formability in relatively complex shapes, composite materials are used in a variety of applications, such as turbine engines or aircraft. For example, composite materials can be installed on or define parts of fuselages and / or wings, rudders, manifolds, airfoils, or other components of aircraft or turbine engines. Attached Figure Description

[0004] The complete and feasible disclosure of this disclosure, including its best mode, is set forth in the specification with reference to the accompanying drawings, for those skilled in the art, wherein:

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

[0006] Figure 2 It includes Figure 1 A schematic perspective view of an aircraft with a non-pipeline or open rotor turbine engine.

[0007] Figure 3 It is suitable for use in accordance with exemplary embodiments of this disclosure. Figure 1 and Figure 2 A schematic perspective view of the composite airfoil assembly and disk assembly used in the turbine engine. The composite airfoil assembly includes the composite airfoil, cladding, and dovetail joint.

[0008] Figure 4 This is based on exemplary embodiments of the present disclosure. Figure 3 A partial exploded view of the composite airfoil assembly.

[0009] Figure 5It is along the exemplary embodiments of this disclosure. Figure 3 The schematic cross-sectional view taken by line VV shows the interior of the composite airfoil and a subset of the cladding applied to the composite airfoil.

[0010] Figure 6 This is based on exemplary embodiments of the present disclosure. Figure 4 or Figure 5 A schematic diagram of a portion of the inner surface of the cladding.

[0011] Figure 7 This illustrates the formation according to exemplary embodiments of the present disclosure. Figure 3 The flowchart shows the method for assembling a composite airfoil component. Detailed Implementation

[0012] The aspects disclosed herein relate to composite components for manufacturing engine parts for turbine engines. The composite component is shown as a composite airfoil assembly having a cladding composite airfoil. The composite airfoil includes at least a core and a skin, shown as a braided core and a laminated skin, wherein the skin is applied to at least a portion of the outer surface of the core. The cladding includes at least two members, wherein at least one of the two members is coupled (e.g., adhered) to the composite airfoil adjacent to its tip or trailing edge. The cladding includes an outer cladding surface and an inner cladding surface. The inner cladding surface is coupled to the outer skin surface. Before coupling the inner cladding surface to the outer skin surface, the inner cladding surface is prepared or otherwise roughened to have an arithmetic mean roughness (Ra) of at least 100. An inner cladding surface having an Ra of at least 100, more specifically, having an Ra between 150 and 400, can improve the bonding between the cladding and the skin. The preparation of the internal surface of each cladding layer results in a morphology with at least 10 random peaks per 645 square millimeters (approximately 1 square inch). When measured from base to peak, the height of the random peaks can range from 0.007 mm to 0.025 mm (approximately 0.0003 inches to 0.0010 inches).

[0013] Extreme loads or sudden forces can be applied to composite components of aircraft or turbine engines. For example, one or more airfoils may experience extreme loads during turbine engine material intake. Cladding with improved connectivity between the inner surface of the cladding and the outer surface of the skin can reduce deflection of the composite airfoil assembly. Furthermore, improved connectivity between the inner surface of the cladding and the outer surface of the skin can reduce cladding-skin separation during extreme load events or when sudden forces are applied to the composite airfoil assembly.

[0014] It should be understood that, in the non-limiting examples, this disclosure applies to other engine components of turbine engines, not just airfoils, such as discs or burner bushings. Furthermore, although described with reference to a core used in the manufacture of the airfoil, it should be understood that this disclosure applies to any other suitable context.

[0015] Reference will now be made in detail to the present embodiments of this disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerals and letter reference numerals to denote features in the drawings. Similar or analogous reference numerals in the drawings and description have been used to denote similar or analogous portions of this disclosure.

[0016] The term "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 superior or better than other implementations. Furthermore, unless explicitly stated otherwise, all embodiments described herein should be considered exemplary.

[0017] As used herein, the terms “first,” “second,” “third,” or “fourth” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components.

[0018] In a context such as “at least one of A, B and C”, the term “at least one” means only A, only B, only C, or any combination of A, B and C.

[0019] The term “turbine” or “turbomachinery” refers to a machine that includes one or more compressors, a heating section (e.g., a combustion section), and one or more turbines that together generate torque output.

[0020] The term "gas turbine engine" refers to an engine that has a turbine as its power source, in whole or in part. Examples of gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, inline turbine engines, non-inline turbine engines, and hybrid electric versions of one or more of these engines.

[0021] The terms "front" and "rear" refer to relative positions within a turbine engine or carrier, and specifically to the normal operating posture of the turbine engine or carrier. For example, for a turbine engine, "front" refers to the position closer to the engine inlet, while "rear" refers to the position closer to the engine nozzle or exhaust port.

[0022] 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 fluid flow. The terms "forward" or "front" indicate being in front of something, and "backward" or "rear" indicate being behind something. For example, when used in relation to fluid flow, forward / frontward can indicate upstream, while backward / rearward can indicate downstream.

[0023] The term "fluid" can refer to a gas or a liquid, or a multiphase system.

[0024] As used herein, the terms "axial" and "axially" refer to a direction and orientation that extends substantially parallel to the centerline of the gas turbine engine. Furthermore, the terms "radial" and "radially" refer to a direction and orientation that extends substantially perpendicular to the centerline of the gas turbine engine. For example, in the context of the turbine engine as a whole, radial refers to the direction of a ray extending between the engine's central longitudinal axis and the engine's outer perimeter. Additionally, as used herein, the terms "circumferential" and "circumferentially" refer to a direction and orientation that extends in an arc around the centerline of the gas turbine engine.

[0025] Unless otherwise specified herein, the terms “connection,” “fixation,” “applied to,” “attached to,” etc., refer to direct connection, fixation, or attachment, as well as indirect connection, fixation, or attachment through one or more intermediate components or features.

[0026] All directional references (e.g., radial, axial, proximal, distal, up, down, upward, downward, left, right, lateral, front, rear, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, backward, etc.) are used for identification purposes only to aid the reader's understanding of this disclosure and do not impose limitations, particularly regarding the location, orientation, or use of aspects of this disclosure described herein. Unless otherwise stated, connective references (e.g., attachment, connection, joint, and engagement) are to be interpreted broadly and may include intermediate structural elements between sets of elements and relative movement between elements. Therefore, a connective reference does not necessarily imply that two elements are directly connected and have a fixed relationship with each other. Exemplary figures are for illustrative purposes only, and the dimensions, positions, order, and relative sizes reflected in the accompanying figures may vary.

[0027] Unless the context clearly specifies otherwise, the singular forms “a,” “an,” and “the” include plural references. Furthermore, as used herein, the term “group” or a “group” of elements can be any number of elements, including only one.

[0028] As used herein, the term "stiffness" can be used to define the degree to which a structure resists deformation in response to a force. Stiffness can be defined as the ratio of the force to the displacement of an object under the action of a force. Stiffness can include resistance to deformation in response to forces applied from various directions, and thus, in non-limiting examples, stiffness can represent axial stiffness, tensile stiffness, compressive stiffness, torsional stiffness, or shear stiffness.

[0029] As used herein, the term "elasticity" can be used to define the modulus of elasticity under tension or compression (Young's modulus) and can refer to the elasticity of a particular material or a structure made of such material (such as the engine components described herein). Elasticity can also be expressed as the stress per unit area relative to local strain or its proportional deformation.

[0030] As used herein, the term "arithmetic mean roughness (Ra)" is an absolute average value relative to a reference length. That is, the Ra value represents the average surface roughness over the length of the measurement being performed, such as the average difference between peaks and valleys. Ra can be defined as the average variation of the roughness profile relative to a mean line.

[0031] As a non-limiting example, Ra can be determined directly, for example, by using a sliding or non-sliding roughness measuring tool. Additionally or alternatively, Ra can be determined optically, for example, by using an interferometer. Additionally or alternatively, Ra can be determined indirectly by combining one or more indentations with direct or indirect measurements.

[0032] Ra can be calculated using the following formula:

[0033] (1) Where L is the length of the region to be inspected, Z(x) is the profile height function or roughness profile, and dx is the differential. The term x can be the profile height.

[0034] 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, nonmetals, or metal and nonmetal elements or materials. Examples of composite materials can be, but are not limited to, polymer matrix composites (PMCs), ceramic matrix composites (CMCs), metal matrix composites (MMCs), carbon fibers, polymeric resins, thermoplastics, bismaleimide (BMI) materials, polyimide materials, epoxy resins, glass fibers, and silicon matrix materials.

[0035] As used herein, a "composite" component refers to a structure or component comprising any suitable composite material. A composite component (e.g., a composite airfoil) may comprise several layers or several plies of composite material. The stiffness, material, and dimensions of the layers or plies may vary to achieve a desired composite component or composite portion of a component having a predetermined weight, size, stiffness, and strength.

[0036] One or more intervening or adhesive layers may be used to form or join composite components. The adhesive may include resins and phenolic resins, where the adhesive may require curing at elevated temperatures or other hardening techniques. As a non-limiting example, if the interior of a first object is coupled to the exterior of a second object, one or more intervening or adhesive layers may be present between the interior of the first object and the exterior of the second object.

[0037] In this disclosure, when a layer is described as being “on” or “above” another layer or substrate, it should be understood that, unless explicitly stated otherwise, these layers may be in direct contact with each other, or there may be another layer or feature between these layers. Therefore, these terms merely describe the relative position of the layers with respect to each other and do not necessarily mean “on top of”, as the relative position above or below depends on the orientation of the device relative to the observer.

[0038] 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 a polymer matrix material (e.g., a thermoplastic resin). Non-limiting examples of processes used to produce thermoplastic prepregs include: hot melt prepreg, in which the fiber reinforcement is drawn through a molten bath of resin; and powder prepreg, in which the resin is deposited onto the fiber reinforcement, as a non-limiting example, electrostatically deposited onto the fiber reinforcement, and then adhered to the fibers, as a non-limiting example, in an oven or with the aid of heated rollers. The prepregs may 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 layups for the part.

[0039] Multilayer prepregs are stacked to the appropriate thickness and orientation of the composite part, and then the resin is cured and solidified to provide fiber-reinforced composite parts. Resins used for PMC matrix materials are generally classified as thermosetting or thermoplastic resins. Thermoplastic resins are generally classified as polymers that can repeatedly soften and flow upon heating and harden upon sufficient cooling due to physical rather than chemical changes. Well-known examples of thermoplastic resins include nylon, thermoplastic polyesters, polyaryletherketones (PAEKs), and polycarbonate resins. Specific examples of high-performance thermoplastic resins envisioned for aerospace applications include polyetheretherketone (PEEK), polyetherketoneketone (PEKK), polyetherimide (PEI), polyaryletherketone (PAEK), and polyphenylene sulfide (PPS). In contrast, thermosetting resins do not undergo significant softening upon heating once fully cured into a rigid solid, but rather thermally decompose upon sufficient heating. Well-known examples of thermosetting resins include epoxy resins, bismaleimide (BMI), and polyimide resins.

[0040] Instead of using prepreg, in another non-limiting example, woven fabrics can be utilized by using thermoplastic polymers. Woven fabrics may include, but are not limited to, dry carbon fibers woven together with thermoplastic polymer fibers or filaments. Non-prepreg braided structures can be fabricated in a similar manner. With this method, the fiber volume of the part can be customized by specifying the relative concentrations of the woven or braided thermoplastic fibers and reinforcing fibers. Furthermore, different types of reinforcing fibers can be braided or woven together at different concentrations to customize the properties of the part. For example, glass fibers, carbon fibers, and thermoplastic fibers can all be woven together at different concentrations to customize the properties of the part. Carbon fibers provide the strength of the system, can be incorporated into glass fibers to enhance impact characteristics—a design feature of parts located near the engine inlet—and thermoplastic fibers provide bonding for the reinforcing fibers.

[0041] In yet another non-limiting example, resin transfer molding (RTM) can be used to form at least a portion of a composite part. Typically, RTM involves applying a dry fiber or matrix material to a mold or cavity. The dry fiber or matrix material may include prepreg, woven material, braided material, or any combination thereof.

[0042] Resin can be pumped into or otherwise supplied to a mold or cavity to impregnate dry fibers or matrix material. The impregnated fibers or matrix material, combined with the resin, is then cured and removed from the mold. Post-curing may be required when removing the composite component from the mold.

[0043] It is conceivable that RTM could be a vacuum-assisted process. That is, air in the cavity or mold can be removed and replaced with resin before heating or curing. It is further conceivable that the placement of dry fibers or matrix materials can be manual or automated.

[0044] Dry fibers or matrix materials can be molded to form composite components or guide resins. Optionally, additional layers or reinforcing layers of materials different from the dry fibers or matrix materials may be included or added prior to heating or curing.

[0045] As used herein, CMC refers to a class of materials having reinforcing fibers within 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, alumina (Al₂O₃), silicon dioxide (SiO₂), aluminosilicates (such as mullite), or mixtures thereof), or mixtures thereof.

[0046] 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, alumina (Al₂O₃), silicon dioxide (SiO₂), aluminosilicates, or mixtures thereof), or mixtures thereof. Optionally, ceramic components (e.g., oxides of Si, Al, Zr, and Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite) may also be included within the ceramic matrix.

[0047] Typically, a particular CMC can be referred to as a combination of its fiber type / matrix type. 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, and SiC / SiC-SiN is a silicon carbide fiber-reinforced silicon carbide / silicon nitride matrix mixture, etc. In other examples, a CMC may consist of a matrix comprising an oxide-based material (such as alumina (Al₂O₃), silicon dioxide (SiO₂), aluminosilicates, and mixtures thereof) and reinforcing fibers. Aluminosilicates may include crystalline materials (e.g., mullite (3Al₂O₃·2SiO₂)) as well as glassy aluminosilicates.

[0048] In some non-limiting examples, the reinforcing fibers may be bundled and / or coated before being incorporated into the ceramic matrix. For example, the fiber bundles may be formed as reinforcing tapes, such as unidirectional reinforcing tapes. Multiple tapes may be stacked together to form a preform component. The fiber bundles may be impregnated with a slurry composition before or after the formation of the preform. The preform may then undergo heat treatment and subsequent chemical treatment (such as silicon melt infiltration) to obtain a component formed from a CMC material having a desired chemical composition. For example, the preform may undergo curing or burnout to produce a high coke residue in the preform and subsequently undergo melt infiltration with silicon, or undergo curing or pyrolysis to produce a silicon carbide matrix in the preform and subsequently undergo chemical vapor infiltration with silicon carbide. Additional steps may be taken to enhance the densification of the preform, either before or after chemical vapor infiltration, by injecting the preform with a liquid resin or polymer and then performing a heat treatment step to fill the voids with silicon carbide. The CMC materials used herein can be formed using any known or later developed methods (including, but not limited to, melt infiltration, chemical vapor infiltration, polymer impregnation pyrolysis (PIP), or any combination thereof).

[0049] The reinforcing fibers can be at least some portions of individual filaments or strands. As used herein, “ceramic fiber bundle,” “fiber bundle,” or simply “bundle” refers to a bundle of multiple individual fibers, filaments, or loose strands. The filaments of a bundle can be randomly mixed or arranged in a pattern, and / or can be continuous or discontinuous. For example, a bundle can include broken filaments or filament segments. As another example, the filaments of a bundle can be substantially parallel, twisted, or otherwise arranged. A bundle can function in substantially the same manner as a single or individual filament. It will also be understood that “individual ceramic filament,” or simply “individual filament,” as used herein, refers to a single or unbundled elongated ceramic component.

[0050] These materials, along with certain monolithic ceramics (i.e., ceramic materials without reinforcement), are particularly well-suited for higher-temperature applications. Furthermore, these ceramic materials are lighter than superalloys while still providing strength and durability for components made from them. Therefore, the use of such materials in many gas turbine components used in the higher-temperature range of gas turbine engines (such as airfoils (e.g., turbine blades and impellers), combustors, shrouds, etc.) is currently under consideration; these components would benefit from the lighter weight and higher-temperature capabilities these materials can offer.

[0051] As used herein, the term "metal" refers to materials that include metals (such as, but not limited to, titanium, iron, aluminum, stainless steel, brass, copper, and nickel alloys). Metallic materials or alloys can be combinations of at least two or more elements or materials (at least one of which is a metal).

[0052] As used herein, the term "metal" refers to materials including, but not limited to, metals such as, but not limited to, titanium, iron, aluminum, stainless steel, brass, copper, and nickel alloys. Metallic materials or alloys can be combinations 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 raw materials in the form of particulate powder or wire are aggregated to form a three-dimensional part. The raw material is then melted by applying heat or other curing processes to form a monolithic single part that can have various integral sub-parts. As used herein, "monolithic" refers to a single structure in which the materials of each layer are melted into or with the materials of adjacent layers without an interface or joint, such that the individual layers lose their properties in the final monolithic structure.

[0053] Suitable additive manufacturing techniques according to this 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 forming (LENS), laser net forming 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.

[0054] Digital light processing (DLP) can include a 3D DLP printer with a transparent tank or canister, a build platform, and a lamp assembly. The transparent tank or canister may contain, for example, a photopolymer resin.

[0055] The DLP build platform can be connected to, for example, an electric 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 the resin in a barrel or tank).

[0056] The DLP printing component can be attached to the lower portion of a build platform facing a barrel or tank. An illumination assembly is located at least partially below the barrel or tank. The illumination assembly may include at least one light source and at least one optical reflector or refractor, such as a deflector or at least one lens.

[0057] A controller connected to or included in a DLP printer can control one or more aspects of the DLP printer, such as the position of the DLP build platform or the intensity, duration, or orientation of the light source.

[0058] In addition to processes such as direct metal laser sintering (DMLS), direct metal laser melting (DMLM), or electron beam melting (EBM) (where an energy source is used to selectively sinter or melt portions of the powder layers), it should be appreciated that, according to an alternative aspect of this disclosure, the additive manufacturing process can be a "binder jetting" process. In this respect, binder jetting involves the continuous deposition of additive powder layers 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 can 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 this subject matter.

[0059] Figure 1This is a schematic cross-sectional view of a gas turbine engine, particularly an open rotor or inline turbine engine 10 for use in aircraft. The inline turbine engine 10 has a generally longitudinally extending axis or engine centerline 12 extending from a front end 14 to a rear end 16. The inline turbine engine 10 includes a set of circumferentially spaced blades or propellers in a downstream series flow relationship, 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 inline turbine engine 10 described herein is a non-limiting example, and other turbine engine architectures are possible, such as, but not limited to, steam turbine engines, supercritical carbon dioxide turbine engines, or any other suitable turbine engine. These other turbine engine architectures can be inline or inline and can have or not have a fan section.

[0060] The outer surface of the non-ducted turbine engine 10, defined by a casing (such as nacelle 40), extends from the front end 14 of the non-ducted turbine engine 10 toward the rear end 16 of the non-ducted turbine engine 10, and covers at least a portion of the compressor section 22, combustion section 28, turbine section 32, and exhaust section 38. A fan section 18 may be located at the front 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 radially arranged around the engine centerline 12. The non-ducted turbine engine 10 includes any number of 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 axially aligned with the combustion section 28.

[0061] The compressor section 22, combustion section 28, and turbine section 32 are collectively referred to as the engine core 44, which generates combustion gases. The engine core 44 is surrounded by an engine housing 46, which is operatively connected to a portion of the nacelle 40 of the non-ducted turbine engine 10.

[0062] An HP shaft or spool 48, coaxially arranged about the engine centerline 12 of the non-ducted turbine engine 10, drives the HP turbine 34 to the HP compressor 26. An LP shaft or spool 50, coaxially arranged within a larger diameter annular HP spool 48 about the engine centerline 12 of the non-ducted turbine engine 10, drives the LP turbine 36 to the LP compressor 24 and the fan 20. The spools 48 and 50 are rotatable about the engine centerline 12 and are connected to a set of rotatable elements that collectively define a rotor 51.

[0063] It should be understood that the non-pipeline turbine engine 10 is a direct drive or integral drive engine that utilizes a reduction gearbox that connects the LP shaft or spool 50 to the fan 20.

[0064] LP compressor 24 and HP compressor 26 each include a set of compressor stages 52 and 54, respectively, in which a set of compressor blades 56 and 58 rotate relative to a corresponding set of static compressor impeller blades 60 and 62 (also referred to as nozzles) to compress or pressurize the fluid flow passing through the stage. In a single compressor stage 52 or 54, multiple compressor blades 56 and 58 are arranged in a ring and extend radially outward from the blade platform relative to the engine centerline 12 to the blade tips, while the corresponding static compressor impeller blades 60 and 62 are positioned upstream of and adjacent to the compressor blades 56 and 58. It is worth noting that... Figure 1 The number of blades, impellers, and compressor stages shown is selected for illustrative purposes only, and other numbers are also possible.

[0065] The compressor blades 56, 58 for the first stage of the compressor are mounted to a disc 61, which is mounted to a corresponding one of the HP spool 48 and the LP spool 50, with each stage having its own disc 61. The static compressor blades 60, 62 for the first stage of the compressor are mounted to the engine housing 46 in a circumferential arrangement.

[0066] HP turbine 34 and LP turbine 36 each comprise a set of turbine stages 64 and 66, respectively, in which a set of turbine blades 68 and 70 rotate relative to a corresponding set of static turbine blades 72 and 74 (also referred to as nozzles) to extract energy from the fluid flow passing through the stage. In a single turbine stage 64 and 66, multiple turbine blades 68 and 70 are arranged in a ring and extend radially outward from the blade platform relative to the engine centerline 12 to the blade tips, while the corresponding static turbine blades 72 and 74 are positioned upstream of and adjacent to the turbine blades 68 and 70. It is worth noting that... Figure 1 The number of blades, impellers, and turbine stages shown is selected for illustrative purposes only; other numbers are also possible.

[0067] Turbine blades 68, 70 for the first stage of turbine section 32 are mounted to disc 71, which is mounted to a corresponding one of HP spool 48 and LP spool 50, with each stage having a dedicated disc 71. Static turbine blades 72, 74 for the first stage of turbine section 32 are mounted to engine housing 46 in a circumferential arrangement.

[0068] The rotating parts of the non-piped turbine engine 10 (e.g., blades 56, 58, 68, 70 in the compressor section 22 and turbine section 32) are also referred to individually or collectively as rotor 51. Therefore, rotor refers to the combination of rotating elements throughout the non-piped turbine engine 10.

[0069] Complementing the rotor section, the stationary parts of the non-ducted turbine engine 10 (e.g., the static blades 60, 62, 72, 74 in the compressor section 22 and turbine section 32) are also referred to individually or collectively as the stator 63. Therefore, stator 63 refers to the combination of non-rotating elements throughout the non-ducted turbine engine 10.

[0070] The nacelle 40 is operatively coupled to the inline turbine engine 10 and covers at least a portion of the engine core 44, engine casing 46, or exhaust section 38. At least a portion of the nacelle 40 extends axially forward or upstream at the indicated location. For example, the nacelle 40 extends axially forward such that a portion of the nacelle 40 covers or conceals a portion of the fan section 18 or turbocharger section (not shown) of the inline turbine engine 10.

[0071] During operation of the non-ducted turbine engine 10, a free-flowing airflow 80 flows against the front of the non-ducted turbine engine 10. A portion of the free-flowing airflow 80 enters an annular region 25 defined by the swept area between the outer surface of the nacelle 40 and the tips of the blades, where this airflow is the inlet airflow 78. A portion of the inlet airflow 78 enters the engine core 44 and is described as the working airflow 76, which is used for combustion within the engine core 44.

[0072] More specifically, the working airflow 76 flows into the LP compressor 24, which then pressurizes the working airflow 76, thereby defining a pressurized airflow supplied to the HP compressor 26, which further pressurizes the air. The working airflow 76 or pressurized airflow from the HP compressor 26 mixes with and ignites fuel in the combustor 30, 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 finally discharged from the inline 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, comprising the pressurized airflow and combustion gases, defines the working airflow flowing through the compressor section 22, the combustion section 28, and the turbine section 32 of the inline turbine engine 10.

[0073] The inlet airflow 78 flows over the set of fan blades 42 and over the nacelle 40 of the non-ducted turbine engine 10. Subsequently, the inlet airflow 78 flows over at least a portion of the set of stationary fan blades 82, which directs the inlet airflow 78 laterally toward the engine centerline 12. Then, the inlet airflow 78 follows the curvature of the nacelle 40 and flows toward the exhaust section 38 over the set of stationary fan blades 82. A pylon 84 mounts the non-ducted turbine engine 10 to an external structure (e.g., the fuselage, wings, tail, etc. of the aircraft).

[0074] The working airflow 76 and at least some of the inlet airflow 78 converge downstream of the exhaust section 38 of the non-ducted turbine engine 10. Together, the working airflow 76 and the inlet airflow 78 form the total thrust of the non-ducted turbine engine 10.

[0075] Imagine a portion of the working airflow 76 is drawn as bleed air 77 (e.g., from compressor section 22). Bleed air 77 supplies airflow to engine components requiring cooling. The temperature of the working airflow 76 exiting the combustor 30 is significantly higher than that of the working airflow 76 within compressor section 22. Therefore, cooling provided by bleed air 77 is necessary for operating these engine components in elevated temperature environments or in the hot sections of the non-ducted turbine engine 10. In the case of a turbine engine, the hot sections of the engine are typically downstream of the combustor 30, particularly the turbine section 32, where the HP turbine 34 is the hottest section because it is directly downstream of the combustion section 28. Other sources of cooling fluid are, but are not limited to, fluids discharged from the LP compressor 24 or the HP compressor 26.

[0076] Figure 2 It includes those suitable for use as Figure 1A schematic perspective view of an aircraft 86 comprising a general-purpose non-ducted turbine engine 88 and a non-ducted turbine engine 10. The aircraft 86 includes a fuselage 90 having an external surface. At least one wing 92 and a tail 94 extend from the fuselage 90. The tail 94 is operatively coupled to and spaced apart from the fuselage 90 via a tail pylon 96. The non-ducted turbine engine 88 is operatively coupled to the external surface of the fuselage 90 via a pylon 98. The non-ducted turbine engine 88 includes a set of circumferentially spaced fan blades 100. A set of stationary fan blades 102 is disposed downstream of the set of circumferentially spaced fan blades 100. The fuselage 90 extends between a nose 104 and a tail 106 and includes a fuselage centerline 108 extending therebetween.

[0077] In addition, although tail fin 94 is a T-wing tail fin (e.g., tail fin 94 as shown in the figure), other conventional tail fins are envisioned, such as cruciform tail fins, H-shaped tails, triple tails, V-shaped tails, inverted tails, Y-shaped tails, twin tails, boom tails, or ring tails, all of which are referred to herein as tail fin 94.

[0078] Figure 3 Is it suitable for Figure 1 non-pipeline turbine engine 10 or Figure 2 A schematic perspective view of the composite airfoil assembly 110 and disk assembly 112 used in the non-ducted turbine engine 88. The disk assembly 112 is suitable for use as disks 61 and 71 (…). Figure 1 Or any other disk, such as, but not limited to, a disk within the fan section 18, compressor section 22, or turbine section 32 of a non-ducted turbine engine 10. The compound airfoil assembly 110 may be rotating or non-rotating, such that the compound airfoil assembly 110 may include static compressor blades 60, 62 ( Figure 1 The compressor blades in this group are 56 and 58. Figure 1 ), static turbine blades 72, 74 ( Figure 1 The turbine blades in this group are 68 and 70. Figure 1 ) or multiple fan blades 42 ( Figure 1 At least one of the following. As a non-limiting example, the composite airfoil assembly 110 may be a composite fan blade assembly.

[0079] The disc assembly 112 can rotate or remain stationary about a rotation axis 114. The rotation axis 114 can be aligned with the engine centerline (e.g., ...). Figure 1 The engine centerline 12) coincides with or deviates from the centerline. The disc assembly 112 includes a plurality of grooves 116 that extend axially through the radially outer portion of the disc assembly 112 and are circumferentially spaced relative to the axis of rotation 114 around the disc assembly 112.

[0080] 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 to the leading edge 124 to define a chordal 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 to the pressure side 132.

[0081] Leading edge 124 and trailing edge 126 extend radially from root 128 to tip 130. Pressure side 132 and suction side 134 opposite to pressure side 132 extend axially between leading edge 124 and trailing edge 126. Dovetail portion 138 may extend from composite airfoil 118.

[0082] 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 one of 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 (e.g., but not limited to welding, adhesion, fastening, etc.). Although only a single composite airfoil assembly 110 is shown, it should be understood that any number of composite airfoil assemblies 110 may be coupled to the disk assembly 112. As a non-limiting example, a plurality of composite airfoil assemblies 110 may be present corresponding to the total number of slots in the plurality of slots 116.

[0083] For ease of reference, a set of relative reference directions and coordinate systems can be applied to the composite airfoil assembly 110. The axial direction (Ad) may extend from front to back and is shown to extend at least partially into the page. The axial direction (Ad) may be arranged parallel to the axis of rotation 114. The radial direction (Rd) extends perpendicular to the axial direction (Ad) and may extend perpendicular to the engine centerline 12. The circumferential direction (Cd) may be defined perpendicular to the radial direction (Rd) and may be defined relative to the engine centerline 12. Figure 1 Along the non-pipeline turbine engine 10 ( Figure 1 The circumference of ).

[0084] Figure 4 This is a partial exploded view of the composite airfoil assembly 110, where the cladding 120 is exploded from the composite airfoil 118. 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 the average length, minimum length, or maximum length measured in the radial direction (Rd) from the root 128 to the tip 130 of the composite airfoil 118.

[0085] Cladding 120 is attached to at least a portion of the outer surface of the composite airfoil 118. As an example, cladding 120 includes a first cladding 150, a second cladding 152, and a third cladding 153. Alternatively, cladding 120 may include any one or more of a fourth cladding 154, a leading-edge sheath 156, or a root cladding 158. Although shown as formed separately, any one or more of the first cladding 150, second cladding 152, third cladding 153, fourth cladding 154, leading-edge sheath 156, or root cladding 158 may be integrally formed or combined.

[0086] Cladding 120 can cover or conceal 0.6% to 70% of the total outer surface area of ​​the composite airfoil 118. More specifically, cladding 120 can cover 2% to 60% of the total outer surface area of ​​the composite airfoil 118, wherein the total outer surface area is defined as the sum of the pressure-side surface area and the suction-side surface area of ​​the composite airfoil 118. More specifically, cladding 120 can cover 5% to 30% of the total outer surface area of ​​the composite airfoil 118. The range of values ​​for cladding 120 maintains the weight efficiency of the composite airfoil 118 and the improved stiffness provided by cladding 120.

[0087] As an example, the first cladding 150 is shown positioned adjacent to the trailing edge 126 of the composite airfoil 118 when mounted, bonded, or otherwise attached to the outer skin surface 142. As used herein, "adjacent to the trailing edge 126" means that the maximum distance measured from the trailing edge 126 to any part of a neighboring object is less than 30% of the airfoil length 140. Furthermore, as an example, the first cladding 150 is shown located at the trailing edge 126 of the composite airfoil 118. That is, when the first cladding 150 is mounted, bonded, or otherwise attached to the composite airfoil 118, at least a portion of the first cladding 150 contacts the trailing edge 126.

[0088] Although shown as being mounted to the pressure side 132 of the composite airfoil 118, it is contemplated that the first cladding 150 may be mounted to the suction side 134 adjacent to the trailing edge 126 or adjacent to the tip 130. As used herein, "adjacent to tip 130" means that the maximum distance measured from tip 130 to any part of an adjacent object is less than 30% of the airfoil length 140.

[0089] Although shown as a general rectangle, the first cladding 150 may have the shape of any combination of one or more regular polygons, irregular polygons, or circles.

[0090] The first cladding 150 includes a first cladding outer surface 160 and a first cladding inner surface 162. The first cladding inner surface 162 may have a Ra value greater than 100. More specifically, the first cladding inner surface 162 may have a Ra value in the range of 150 to 400. Still more specifically, the first cladding inner surface 162 may have a Ra value in the range of 150 to 300. When coupled to the composite airfoil 118, the first cladding inner surface 162 of the first cladding 150 is coupled to a portion of the outer skin surface 142 of the composite airfoil 118.

[0091] As shown by the dashed line, the first coverage area 164 is defined as the amount of area of ​​the outer skin surface 142 of the composite airfoil 118 covered by the first cladding 150 when the inner surface 162 of the first cladding is adhered, bonded, or otherwise attached to the outer skin surface 142. In other words, the first coverage area 164 is the portion of the outer skin surface 142 attached to the inner surface 162 of the first cladding. The first coverage area 164 can range from 0.3% to 70% of the total outer surface area of ​​the composite airfoil 118. More specifically, the first coverage area 164 can range from 1% to 20% of the total outer surface area of ​​the composite airfoil 118. The range of values ​​for the first coverage area 164 of the first cladding 150 maintains the weight efficiency of the composite airfoil 118 and the improved stiffness provided by the first cladding 150.

[0092] As an example, the second cladding 152 is shown positioned adjacent to the trailing edge 126 of the composite airfoil 118 when mounted, coupled, or otherwise attached to the outer skin surface 142 of the composite airfoil 118. Furthermore, as an example, the second cladding 152 is shown located at the trailing edge 126 of the composite airfoil 118.

[0093] Although shown as a general rectangle, the second cladding 152 may have any combination of shapes, including one or more regular polygons, irregular polygons, or circles.

[0094] Similar to the first cladding 150, the second cladding 152 includes a second cladding outer surface 166 and a second cladding inner surface 168. The second cladding inner surface 168 may have an Ra value greater than 100. More specifically, the second cladding inner surface 168 may have an Ra value in the range of 150 to 400. Even more specifically, the second cladding inner surface 168 may have an Ra value in the range of 150 to 300.

[0095] It is envisioned that the Ra value of the inner surface 168 of the second cladding layer may be within 5% of the Ra value of the inner surface 162 of the first cladding layer. Alternatively, in another different and non-limiting example, it is envisioned that the Ra value of the inner surface 168 of the second cladding layer may differ from the Ra value of the inner surface 162 of the first cladding layer. That is, while falling within the same range, the Ra value of the inner surface 168 of the second cladding layer may differ from the Ra value of the inner surface 162 of the first cladding layer. It is further envisioned that the percentage difference between the Ra value of the inner surface 168 of the second cladding layer and the Ra value of the inner surface 162 of the first cladding layer may be greater than 5%. The difference between the Ra value of the inner surface 168 of the second cladding layer and the Ra value of the inner surface 162 of the first cladding layer may be caused by one or more of the following: the second cladding layer 152 and the first cladding layer 150 are made of different materials; the second cladding layer 152 and the first cladding layer 150 are made using the same materials but in different proportions; or the second cladding layer 152 and the first cladding layer 150 have different thicknesses 216, 222 (…). Figure 5 ).

[0096] When attached to the composite airfoil 118, the inner surface 168 of the second cladding 152 is attached to a portion of the outer skin surface 142. The second coverage area (not shown) is defined as the amount of area of ​​the outer skin surface 142 of the composite airfoil 118 covered by the second cladding 152 when the second cladding 152 is bonded to or otherwise attached to the outer skin surface 142 on the suction side 134 of the composite airfoil 118. The second coverage area can range from 0.3% to 70% of the total surface area of ​​the composite airfoil 118. More specifically, the second coverage area can range from 1% to 20% of the total outer surface area of ​​the composite airfoil 118. The range of values ​​for the second coverage area of ​​the second cladding 152 maintains the weight efficiency of the composite airfoil 118 and the improved stiffness provided by the second cladding 152.

[0097] The second covering area of ​​the second cladding 152 may have a similar shape, mirror shape, or area to the first covering area 164 of the first cladding 150. That is, the area of ​​the first covering area 164 may be equal to the area of ​​the second covering area, or within 20% of it.

[0098] Imagine that when attached to the outer skin surface 142 of the composite airfoil 118, at least a portion of the first cladding 150 and the second cladding 152 can come into contact. That is, the first cladding 150 and the second cladding 152 may have edges that will come into contact, for example, when mounted to the composite airfoil 118.

[0099] Alternatively, in various and non-limiting examples, the first cladding 150 and the second cladding 152 may be integrally formed, having a V-shaped or U-shaped cross-section. That is, the first cladding 150 may be partially present on both the suction side 134 and the pressure side 132, such that a portion of the integrally formed first and second claddings 150, 152 covers a portion of the trailing edge 126. In other words, the first cladding 150 or the second cladding 152 may define a cladding surrounding the trailing edge 126 of the composite airfoil 118, wherein a portion of the cladding is on the pressure side 132, while another portion extends to the suction side 134. Further envisioned, a third cladding 153 may surround the top, partially present on both the suction side 134 and the pressure side 132.

[0100] As an example, the third cladding 153 is shown positioned adjacent to the tip 130 of the composite airfoil 118 when mounted, bonded, or otherwise attached to the outer skin surface 142 of the composite airfoil 118. Furthermore, as an example, the third cladding 153 is shown located at the tip 130 of the composite airfoil 118. That is, when the third cladding 153 is mounted, bonded, or otherwise attached to the composite airfoil 118, at least a portion of the third cladding 153 contacts the tip 130.

[0101] Although shown as having a generally L-shaped body, the third cladding 153 may have a shape of any combination of one or more polygons, irregular polygons, or circles.

[0102] The third cladding 153 includes an outer surface 170 and an inner surface 172. The inner surface 172 may have a Ra value greater than 100. More specifically, the inner surface 172 may have a Ra value in the range of 150 to 400. Even more specifically, the inner surface 172 may have a Ra value in the range of 150 to 300.

[0103] It is envisioned that the Ra value of the outer surface 170 of the third cladding layer may be within 5% of the Ra value of the inner surface 162 of the first cladding layer or the inner surface 168 of the second cladding layer. Alternatively, in another different and non-limiting example, it is envisioned that the Ra value of the outer surface 170 of the third cladding layer may differ from the Ra value of the inner surface 162 of the first cladding layer or the inner surface 168 of the second cladding layer. That is, while falling within the same range, the Ra value of the inner surface 170 of the third cladding layer may differ from the Ra value of the inner surface 162 of the first cladding layer or the inner surface 168 of the second cladding layer. It is further envisioned that the percentage difference between the Ra value of the inner surface 170 of the third cladding layer and the Ra value of the inner surface 162 of the first cladding layer or the inner surface 168 of the second cladding layer may be greater than 5%. The difference between the Ra value of the inner surface 170 of the third cladding layer and the Ra value of the inner surface 162 of the first cladding layer or the inner surface 168 of the second cladding layer can be caused by one or more of the following: the third cladding layer 153 and the first or second cladding layers 150, 152 are made of different materials; the third cladding layer 153 and the first or second cladding layers 150, 152 are made of the same materials but in different proportions; or the third cladding layer 153 and the first or second cladding layers 150, 152 have different thicknesses.

[0104] When coupled to the composite airfoil 118, the inner surface 172 of the third cladding 153 is coupled to a portion of the outer skin surface 142 of the composite airfoil 118. The third coverage area 174, indicated by the dashed line, can be defined as the surface area covered by the third cladding 153 when it is coupled to the pressure side 132 of the composite airfoil 118. The third coverage area 174 can range from 0.3% to 70% of the total surface area of ​​the composite airfoil 118. More specifically, the third coverage area 174 can range from 2% to 20% of the total outer surface area of ​​the composite airfoil 118. The range of values ​​for the third coverage area 174 provides a balance between the weight benefits of the composite airfoil 118 and the increased stiffness or strength from the third cladding 153.

[0105] Imagine that when coupled to the outer skin surface 142 of the composite airfoil 118, at least a portion (shown as edge portion 176) of the first cladding 150 and the third cladding 153 can contact. Additionally or alternatively, the leading portion 178 of the third cladding 153 can contact one or more portions of the leading edge sheath 156.

[0106] Optionally, cladding 120 may include a fourth cladding 154. The fourth cladding 154 may be attached to the side of the composite airfoil 118 opposite to the third cladding 153. As shown, the fourth cladding 154 may be attached to the suction side 134. It is envisioned that the fourth cladding 154 may have a surface area with a similar geometry to the third cladding 153. That is, the surface area of ​​the fourth cladding 154 may be within 10% of that of the third cladding 153. Alternatively, in another non-limiting example, it is envisioned that the surface area of ​​the fourth cladding 154 may have a percentage difference greater than 10% compared to the third cladding 153.

[0107] The fourth cladding 154 includes an outer surface 180 and an inner surface 182. The inner surface 182 may have an Ra value greater than 100. More specifically, the inner surface 182 may have an Ra value in the range of 150 to 400. Even more specifically, the inner surface 182 may have an Ra value in the range of 150 to 300.

[0108] It is envisioned that the Ra value of the outer surface 180 of the fourth cladding layer may be within 5% of the Ra value of one or more of the inner surfaces 162, 168, or 170 of the first cladding layer. Alternatively, in another different and non-limiting example, it is envisioned that the Ra value of the outer surface 180 of the fourth cladding layer may differ from the Ra values ​​of the inner surfaces 162, 168, and 170 of the first cladding layer. That is, while falling within the same range, the Ra value of the inner surface 180 of the fourth cladding layer may differ from the Ra value of the inner surfaces 162, 168, or 170 of the first cladding layer. It is further envisioned that the percentage difference between the Ra value of the inner surface 180 of the fourth cladding layer and the Ra value of one or more of the inner surfaces 162, 168, or 170 of the third cladding layer may be greater than 5%. The difference between the Ra value of the inner surface 180 of the fourth cladding layer and the Ra value of one or more of the inner surface 162 of the first cladding layer, the inner surface 168 of the second cladding layer, or the outer surface 170 of the third cladding layer can be caused by one or more of the following: the fourth cladding layer 154 and the first, second, or third cladding layers 150, 152, 153 are made of different materials; the fourth cladding layer 154 and the first, second, or third cladding layers 150, 152, 153 are made of the same materials but in different proportions; or the fourth cladding layer 154 and the first, second, or third cladding layers 150, 152, 153 have different thicknesses.

[0109] Alternatively, in various and non-limiting examples, the third cladding 153 and the fourth cladding 154 may be integrally formed, having a V-shaped or U-shaped cross-section. That is, the third cladding 153 may be partially present on both the suction side 134 and the pressure side 132, such that a portion of the integrally formed third and fourth claddings 153, 154 covers a portion of the tip 130. In other words, the third cladding 153 or the third cladding 153 and the fourth cladding 154 may define a cladding surrounding the tip 130 of the composite airfoil 118, wherein a portion of the cladding is on the pressure side 132, while another portion extends to the suction side 134.

[0110] The leading edge sheath may surround the leading edge 124. That is, the leading edge sheath 156 may have a V-shaped or U-shaped cross-section.

[0111] The leading edge sheath 156 includes an outer sheath surface 184 and an inner sheath surface 186. The inner sheath surface 186 may have an Ra value greater than 100. More specifically, the inner sheath surface 186 may have an Ra value in the range of 150 to 400. Even more specifically, the inner sheath surface 186 may have an Ra value in the range of 150 to 300.

[0112] The leading edge sheath 156 may be mounted, coupled, or otherwise attached to the composite airfoil 118 at the leading edge 124. The sheath coverage area 187 of the leading edge sheath 156 may have a larger surface area than the first coverage area 164, the second coverage area, or the third coverage area 174.

[0113] The root cladding 158 may include a pressure-side root portion 188, a trailing-edge root portion 190, and a suction-side root portion 192. Each portion of the root cladding 158 may have an outer root surface 194 and an inner root surface 196, wherein one or more portions of the inner root surface 196 have a Ra value greater than 100. More specifically, one or more portions of the inner root surface 196 may have a Ra value in the range of 150 to 400. Even more specifically, one or more portions of the inner root surface 196 may have a Ra value in the range of 150 to 300.

[0114] The root cladding 158 may be mounted, joined, or otherwise attached to the composite airfoil 118 adjacent to the dovetail portion 138 or at the root 128 of the composite airfoil 118. As used herein, "adjacent to the dovetail portion 138" means that the maximum distance measured from the dovetail portion 138 to any part of a neighboring object is less than 30% of the airfoil length 140.

[0115] The root coverage area 198 of the root cladding 158 may have a surface area smaller than that of the first coverage area 164, the second coverage area, the third coverage area 174, or the sheath coverage area 187. Alternatively, in another non-limiting example, the root coverage area 198 may be equal to or greater than that of the first coverage area 164, the second coverage area, the third coverage area 174, or the sheath coverage area 187.

[0116] It is envisioned that the Ra value of the inner surface 186 of the sheath or the inner surface 196 of the root can be within 5% of the Ra value of one or more of the inner surfaces 162, 168, 170, and 180 of the first, second, third, or fourth cladding layers. Alternatively, in another different and non-limiting example, it is envisioned that the Ra value of the inner surface 186 of the sheath or the inner surface 196 of the root can be different from the Ra value of one or more of the inner surfaces 162, 168, 170, and 180 of the first, second, third, or fourth cladding layers. That is, while falling within the same range, the Ra value of the inner surface 186 of the sheath or the inner surface 196 of the root can be different from the Ra value of one or more of the inner surfaces 162, 168, 170, and 180 of the first, second, third, or fourth cladding layers.

[0117] It is further envisioned that the percentage difference between the Ra value of the inner surface 186 of the sheath or the inner surface 196 of the root and the Ra value of one or more of the inner surfaces 162, 168, 170, and 180 of the first, second, third, or fourth cladding layers can be greater than 5%. The difference between the Ra value of the inner surface 186 of the sheath or the inner surface 196 of the root and the Ra value of one or more of the inner surfaces 162, 168, 170, and 180 of the first, second, third, or fourth cladding layers can be caused by one or more of the following: being made of different materials, being made using the same materials but in different proportions, or having different cladding thicknesses.

[0118] Imagine using adhesive 232 ( Figure 5 Cladding 120 to the outer skin surface 142 of the composite airfoil 118 may be attached by means of a mechanical fastener or other means. As a non-limiting example, adhesives used to attach the first cladding 150, second cladding 152, third cladding 153, fourth cladding 154, leading edge sheath 156, or root cladding 158 to the outer skin surface 142 of the composite airfoil 118 may include epoxy resins, phenolic resins, adhesive films, adhesive tapes, cyanoacrylates, anaerobic adhesives, thermoplastic adhesives, polymeric resins, or other thermosetting adhesives.

[0119] The first cladding layer 150, the second cladding layer 152, the third cladding layer 153, or the fourth cladding layer 154 may be formed of one or more of a metallic material, a thermoplastic material, or a composite material, and have a second bulk modulus. As used herein, "bulk modulus" refers to a numerical constant describing the elastic properties of a solid or fluid under pressure on all surfaces. The first cladding layer 150, the second cladding layer 152, the third cladding layer 153, and the fourth cladding layer 154 may be made of similar materials. That is, the first bulk modulus of the first cladding layer 150 may be equal to, or within 10% of, the second bulk modulus of the second cladding layer 152, the third bulk modulus of the third cladding layer 153, or the fourth bulk modulus of the fourth cladding layer 154. Additionally or alternatively, the second bulk modulus of the second cladding layer 152 may be equal to, or within 10% of, the third bulk modulus of the third cladding layer 153 or the fourth bulk modulus of the fourth cladding layer 154. Additionally or alternatively, the third bulk modulus of the third cladding 153 may be equal to, or within 10% of, the fourth bulk modulus of the fourth cladding 154.

[0120] The leading edge sheath 156 or the root cladding 158 may comprise one or more of a metallic material, a thermoplastic material, or a composite material, and has a fifth bulk modulus. The bulk modulus of the sheath or the root is equal to, or within 20% of, the value of the first, second, third, or fourth bulk modulus.

[0121] Alternatively, in various and non-limiting examples, the bulk modulus of the sheath or the bulk modulus of the root may be greater than 20% of the value of the first, second, third, or fourth bulk modulus.

[0122] Although shown by way of example as having five different cladding layers, cladding 120 may include any number of cladding layers positioned adjacent to tip 130, adjacent to trailing edge 126, adjacent to dovetail portion 138, or positioned at leading edge 124 of composite airfoil 118.

[0123] Further, although shown as a single piece, the first cladding 150, the second cladding 152, the third cladding 153, the leading edge sheath 156, the root cladding 158, or the fourth cladding 154 can be a combination of multiple cladding pieces.

[0124] Figure 5 The figure shown is taken along section VV. Figure 3A schematic cross-sectional view of the composite airfoil assembly 110 shows portions of the interior 200 and cladding 120. The composite airfoil 118 includes a core, shown as a braided core 202, and a skin, shown as a laminated skin 204 disposed on the braided core 202. The braided core 202 includes a composite structure having a core bulk modulus. The core bulk modulus is less than the first bulk modulus of the first cladding 150 and the second bulk modulus of the second cladding 152. It is also contemplated that the core bulk modulus is less than the third bulk modulus of the third cladding 153 and the fourth bulk modulus of the fourth cladding 154. Further contemplated, the core bulk modulus may be less than the sheath bulk modulus or the root bulk modulus.

[0125] In a non-limiting example, the braided core 202 can be dry, requiring no additional material, or alternatively, impregnated and cured with resin. The braided core 202 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 in the axial direction Ad, the radial direction Rd, and the circumferential direction Cd. Figure 3 The combination of ) is woven, and it should be understood that the woven pattern can be used with non-pipeline turbine engines 10 ( Figure 1 The braided pattern is formed and defined separately, such that it is woven in any three mutually orthogonal planes to define a three-dimensional object relative to said planes. In a non-limiting example, the braided structure may include a three-dimensional braid comprising a plurality of warp fibers 206 and a plurality of weft fibers 208 that can be woven in three directions to form a three-dimensional braided pattern structure of the braided core 202. The three directions of the plurality of warp fibers 206 and the plurality of weft fibers 208 may be along the axial direction Ad, the radial direction Rd, and the circumferential direction Cd. Figure 3 ( ) is defined, or is at an angle relative to it. In a non-limiting example, a jacquard loom or 3D loom can be used to create complex three-dimensional woven patterns or structures, which may include interlacing one or more composite materials to form a woven core 202. The woven core 202 may be composed of composite materials such as carbon or carbon fibers, glass or glass fibers, nylon, rayon, or aramid fibers, while other materials such as nickel, titanium, or ceramic composite materials are contemplated in the non-limiting examples.

[0126] Further envisioning, the braided core 202 can be formed as a three-dimensional braided structure with an arranged or knitted geometry or pattern. For example, the arranged or knitted geometry or pattern can include a braided pattern comprising three or more interlaced fibers woven in a repeating pattern. In another non-limiting example, the arranged geometry can include a set of fibers or strands sequentially stacked on top of each other to define the arranged geometry. The braided or arranged geometry or pattern can be repeated over the entire braided core 202 or only over a portion thereof. Such additional arranged geometry can be similar, where the fiber arrangement is the same but the orientation is different, or where the fiber arrangement is different and the orientation can be similar or dissimilar. The arranged geometry or pattern can be formed using composite materials using a jacquard loom or a 3D loom. The three-dimensional arrangement structure can include a braided pattern extending in three dimensions (such as a combination of the axial direction Ad, the radial direction Rd, and the circumferential direction Cd).

[0127] The laminated skin 204 can be formed as a set of laminated layers disposed around or around the braided core 202. In a non-limiting example, the laminated skin 204 can be a pre-impregnated, fiber-placed, or dry fiber laminate. In a non-limiting example, such a laminated layer forming the laminated skin 204 can be formed by resin transfer molding (RTM), partial RTM, identical qualified resin transfer molding (SQRTM), or autoclave treatment. The laminated skin 204 can include a skin bulk modulus. The skin bulk modulus can differ from the core bulk modulus. In one example, the surface bulk modulus can be greater than the core bulk modulus.

[0128] The laminated skin 204 may include an inner skin surface 210 and an outer skin surface 142, the outer skin surface 142 defining at least a portion of the outer surface of the composite airfoil 118. The inner skin surface 210 may at least partially contact at least a portion of the core exterior 212 of the braided core 202. In other words, the laminated skin 204 may be applied to at least a portion of the core exterior 212 of the braided core 202.

[0129] The axial airfoil length 214 can be measured between the leading edge 124 and the trailing edge 126 of the composite airfoil 118.

[0130] The first cladding 150 is attached to the outer surface or skin outer surface 142 of the composite airfoil 118, such that the inner surface 162 of the first cladding faces the skin outer surface 142 of the composite airfoil 118. The first cladding thickness 216, extending from the inner surface 162 of the first cladding 150 to the outer surface 160 of the first cladding, can range from 0.0001% to 10% of the axial airfoil length 214. More specifically, the first cladding thickness 216 can range from 0.001% to 2% of the axial airfoil length 214.

[0131] Similar to the first cladding 150, the second cladding 152 has a second cladding thickness 222 extending from the outer surface 166 of the second cladding to the inner surface 168 of the second cladding. The second cladding thickness 222 can range from 0.0001% to 10% of the axial airfoil length 214. More specifically, the second cladding thickness 222 can range from 0.001% to 2% of the axial airfoil length 214. Further, it is envisioned that the second cladding thickness 222 can range from 0.07% to 0.25% of the axial airfoil length 214. Although shown as substantially uniform, it is envisioned that the second cladding thickness 222 can vary, for example, in the axial or radial direction.

[0132] Coating 230 can be applied directly to the laminated skin 204. However, it is contemplated that an intermediate adhesive layer be provided between the laminated skin 204 and coating 230. Although shown as covering a portion of the laminated skin 204, coating 230 can cover the entire laminated skin 204. It is further contemplated that coating 230 can be provided on the laminated skin 204, wherein a portion of the laminated skin 204 is not covered by cladding 120. It is also further contemplated that coating 230 can be applied to one or more portions of cladding 120. Coating 230 can include a coating bulk modulus that is different from or the same as one or more of the core bulk modulus or the skin bulk modulus. It is contemplated that the coating bulk modulus is less than a first bulk modulus, a second bulk modulus, a third bulk modulus, or a fourth bulk modulus.

[0133] Additionally, coating 230 can be an environmental barrier coating, for example, it can be used to resist oxidation or corrosion. In another example, coating 230 can be a thermal barrier coating that at least partially covers the geothermal insulation braided core 202 and the laminated skin 204. Additional non-limiting examples of the coating may include anti-icing coatings, such as polyurethane, de-icing materials, ultraviolet radiation coatings, or oil barrier coatings, such as polyethylene or polypropylene. In one additional non-limiting example, coating 230 may be formed as a polypropylene base layer and a polyurethane layer disposed on the polypropylene base layer. It is further envisioned that an external coating layer (not shown) may be provided on the exterior of coating 230, wherein such a coating layer can provide radiation protection, such as ultraviolet radiation.

[0134] Optionally, adhesive 232 may be applied to one or more portions of the outer surface 142 of the skin. Adhesive 232 may be any number of adhesive layers, including one layer. Adhesive 232 may include, but is not limited to, epoxy resin, phenolic resin, film, tape, cyanoacrylate, anaerobic adhesive, thermoplastic adhesive, polymer resin or other thermosetting adhesive.

[0135] Alternatively, in various and non-limiting examples, adhesive 232 may be applied to the cladding 120. More specifically, adhesive 232 may be applied to the inner surface 162 of the first cladding and the inner surface 168 of the second cladding. It is contemplated that adhesive 232 may be applied to the cladding 120 or to both the cladding 120 and the outer surface 142 of the skin.

[0136] Further envisioning, one or more adhesive layers or one or more types of adhesives may be located between coating 230 or skin 204 and leading edge sheath 156, third cladding 153 ( Figure 4 ), fourth cladding 154 ( Figure 4 ), root cladding 158 ( Figure 4 Between ) or any combination thereof.

[0137] Figure 6 A schematic portion of the first inner surface 162 is shown. For reference, a three-dimensional coordinate system may be applied to this portion of the first inner surface 162. The three-dimensional coordinate system may include directions indicating the length 240, width 242, and height 244, wherein the length 240 is perpendicular to the width 242, and the height 244 is perpendicular to both the length 240 and the width 242.

[0138] As a non-limiting example, this portion of the first inner surface 162 is shown as having a sample length 246 of 645 mm (approximately 1 inch) and a sample width 248. That is, this portion of the first inner surface 162 can be 645 mm by 645 mm or approximately 1 square inch.

[0139] The fabrication of the first internal surface 162 can produce a morphology with random peaks 250. It is envisioned that this portion of the fabricated first internal surface 162 comprises at least ten random peaks 250 per 645 square millimeters (approximately 1 square inch). As used herein, the term "random peak" is defined as a protrusion having a height measured from a base point 252 or reference plane to a maximum distance 254, which is furthest from the base point 252 or reference plane along a height axis 244, where the height is greater than a predetermined minimum 256, shown as 0.007 millimeters (approximately 0.0003 inches). More specifically, a protrusion is considered a random peak if it appears randomly and has a height ranging from 0.007 millimeters to a predetermined maximum 258 (shown as 0.025 millimeters) (approximately 0.0003 inches and 0.0010 inches).

[0140] As an example, although this portion of the first inner surface 162 is shown as having ten random peaks 250, more random peaks can be envisioned. While the random peaks 250 are shown as generally shaped like irregular truncated cones, any shape protruding from a reference plane or base point can be envisioned. As shown, the random peaks 250 can have varying or similar lengths, widths, or heights, wherein the height of all random peaks 250 is greater than a predetermined minimum value 256.

[0141] Further, it is envisioned that the height, maximum height, or highest height of the random peak 250 could indicate the amount of material removed during the preparation process. That is, the preparation of the first internal surface 162 can form the random peak 250. This portion of the first internal surface 162 is prepared to have an Ra value greater than 100.

[0142] Although shown as part of the first inner surface 162, it is conceivable that the schematic portion may originate from the second cladding inner surface 168, the third cladding inner surface 170, the fourth cladding inner surface 180, the root inner surface 196, or the sheath inner surface 186.

[0143] Furthermore, the figures are not scaled to scale, and elements may be exaggerated for ease of description and interpretation.

[0144] Figure 7 A method 300 for forming a composite airfoil assembly 110 is shown. See also: Figure 4 and Figure 5 At 302, the braided core 202 may be formed using a composite material. The braided core 202 may be formed to define a specific braided structure. The specific braided structure may be a preform specified to have a predetermined geometry, or its size and shape may be determined after the fabrication of the braided structure, such as by cutting or grinding the braided core 202.

[0145] At 304, the skin shown as a laminated skin 204 is applied to at least a portion of the core exterior 212 of the braided core 202. Optionally, an adhesive material layer or other layer is provided between the braided core 202 and the laminated skin 204. The laminated skin 204 and the braided core 202 define a composite airfoil 118.

[0146] Optionally, it is envisioned that coating 230 can be applied before or after curing. That is, coating 230 can be applied to one or more portions of the outer surface 142 of the skin. It is envisioned that coating 230 can be applied to portions of the outer surface 142 of the skin that do not receive cladding 120. Alternatively, in another non-limiting example, it is also envisioned that coating 230 can be applied to one or more portions of the first coverage area 164, the second coverage area, the third coverage area 174, the sheath coverage area 187, or the root coverage area 198. It is further envisioned that any portion of the outer surface 142 of the skin or the outer surface of cladding 120 can receive coating 230.

[0147] Compared to the adhesion between the laminated skin 204 and the non-woven core, the braided structure of the braided core 202 provides greater adhesion to the laminated skin 204 and can provide an improved stiffness transition, elastic transition, or bulk modulus transition between the braided core 202 and the coating 230, so that the difference in stiffness, elasticity, or bulk modulus between adjacent materials is smaller than the difference compared with non-adjacent materials or airfoils with non-woven cores.

[0148] Optionally, additional composite layers, resins, adhesives, or other materials may be added to or applied to the composite airfoil 118 prior to curing. As a non-limiting example, the braided core may be impregnated with resin prior to curing.

[0149] At 306, the composite airfoil 118 is cured. The curing process depends on one or more of time, temperature, or pressure. Curing the composite airfoil 118 may include separate curing stages or steps. It is contemplated that curing, additional curing, or partial curing of one or more portions of the composite airfoil assembly 110 may occur during or between any one or more steps of method 300.

[0150] At 308, the cladding 120, shown as a first cladding 150, a second cladding 152, a third cladding 153, and an optional fourth cladding 154, is formed of a composite material, a thermoplastic material, a metallic material, or any combination thereof. It is contemplated that the metallic material included in the first cladding 150, the second cladding 152, the third cladding 153, or the fourth cladding 154 may be titanium, nickel, steel, tin, aluminum, brass, copper, or any combination thereof.

[0151] The formation of cladding 120 may include additive manufacturing, machining, casting of metal materials to form cladding 120, or any combination thereof.

[0152] Once formed, the cladding 120, which is shown by way of example as a first cladding 150, a second cladding 152, a third cladding 153, and an optional fourth cladding 154, has a cladding bulk modulus greater than the core bulk modulus or multiple cladding bulk moduli.

[0153] Optionally, at 310, at least a portion of the cladding 120 may be etched. Etching at 310 may remove surface contaminants, such as, but not limited to, alpha case or oxides.

[0154] As a non-limiting example, etching can be chemical etching, wherein an etching solvent is applied to at least a portion of the first cladding 150, the second cladding 152, the third cladding 153, and optionally the fourth cladding 154. It is contemplated that the leading edge sheath 156 or the root cladding 158 may also be etched. Additionally or alternatively, portions of the first cladding 150, the second cladding 152, the third cladding 153, or the fourth cladding 154 may be immersed in a chemical bath. The chemical substance used to etch the cladding 120 may include hydrofluoric acid. Additionally or alternatively, etching may include the mechanical removal of material from the surface.

[0155] At 312, at least a portion of the cladding 120 is prepared. That is, at least a portion of the cladding 120 is roughened to have an Ra of at least 100. It is contemplated that the prepared at least a portion of the cladding 120 is one or more portions of the inner surface 162 of the first cladding, the inner surface 168 of the second cladding, the inner surface 172 of the third cladding, the inner surface 182 of the fourth cladding, the inner surface 186 of the sheath, the inner surface 196 of the root, or any combination thereof.

[0156] At least a portion of the prepared cladding 120 has an Ra value of at least 100. Further envisioning, one or more portions of the inner surfaces 162 of the first cladding, 168 of the second cladding, 172 of the third cladding, 182 of the fourth cladding, 186 of the sheath, or 196 of the root have an Ra value in the range of 150 to 400. More specifically, one or more of the prepared surfaces of the inner surfaces 162 of the first cladding, 168 of the second cladding, 172 of the third cladding, 182 of the fourth cladding, 186 of the sheath, or 196 of the root have an Ra value in the range of 150 to 300.

[0157] Preparing at least a portion of the cladding 120 may include removing material from one or more portions of the first cladding inner surface 162, the second cladding inner surface 168, the third cladding inner surface 172, the fourth cladding inner surface 182, the sheath inner surface 186, or the root inner surface 196.

[0158] The preparation of the internal surfaces of each cladding results in the morphology of the first cladding internal surface 162, the second cladding internal surface 168, the third cladding internal surface 172, the fourth cladding internal surface 182, the sheath internal surface 186, or the root internal surface 196 having at least 10 random peaks 250 per 645 square millimeters (approximately 1 square inch). When measured from base to peak, the random peaks can have heights ranging from 0.007 mm to 0.025 mm (approximately 0.0003 inches and 0.0010 inches).

[0159] The material removal depth can range from 0.007 mm to 0.130 mm (approximately 0.001 inch to 0.005 inch). More specifically, the material removal depth can range from 0.025 mm to 0.127 mm (approximately 0.001 inch to 0.005 inch). Additionally or alternatively, the material removal depth can range from 0.001% to 10% of the first cladding thickness 216 or the second cladding thickness 222. Removing more than 10% may reduce the stiffness or strength of the cladding 120, while removing less than 0.001% will not provide the desired increase in bonding between the cladding 120 and the composite airfoil 118.

[0160] Optionally, at 314, additional bonding preparation may be performed on one or more portions of the cladding 120. Additional bonding preparation may include additional etching cycles, application of a conversion coating, or application of an undercoat to one or more portions of the cladding 120.

[0161] As a non-limiting example, additional etching cycles may include, but are not limited to, chemical etching. As a non-limiting example, the conversion coating may include water-based solution gelation (sol-gel). It is envisioned that one or more coatings of the same or different types of conversion coatings may be applied to one or more portions of the cladding 120.

[0162] An undercoat can promote adhesion between metallic materials and other materials. As a non-limiting example, the undercoat can be an epoxy resin. As an example, it is envisioned that the undercoat may include strontium, chromium, or zinc.

[0163] Imagine that one or more of the following could further enhance the bonding or adhesion of one or more portions of the cladding 120: additional etching cycles, application of a conversion coating, or application of an undercoat.

[0164] At 316, after preparing the inner surface of each cladding layer, at least a portion of the cladding layer 120 is adhered, applied, or otherwise attached to the outer skin surface 142 of the composite airfoil 118. That is, the first cladding layer 150, the second cladding layer 152, the third cladding layer 153, the fourth cladding layer 154, the leading edge sheath 156, the root cladding layer 158, or any combination thereof may be adhered at the trailing edge 126, leading edge 124, tip 130, or root 128 of the composite airfoil 118 to the pressure side 132 or the suction side 134. More specifically, one or more portions of the inner surface 162 of the first cladding layer, the inner surface 168 of the second cladding layer, the inner surface 172 of the third cladding layer, the inner surface 182 of the fourth cladding layer, the inner surface 186 of the sheath, or the inner surface 196 of the root face the outer skin surface 142.

[0165] The adhesive coating 120 may include applying adhesive 232 to the inner surface of the coating, which is shown as a first inner surface 162, a second inner surface 168, a third inner surface 172, a fourth inner surface 182, a sheath inner surface 186, or a root inner surface 196, an outer surface 142, or an inner surface of the coating and an outer surface of the sheath.

[0166] As a non-limiting example, adhesion may include adhering a first cladding 150 to the pressure side 132 at or near the trailing edge 126, and adhering a second cladding 152 to the suction side 134 at or near the trailing edge 126.

[0167] As a further non-limiting example, adhesion may include adhering a third sheath 153 to the pressure side 132 at or near the tip 130, and adhering a fourth sheath 154 to the suction side 134 at or near the tip 130.

[0168] If coating 230 is applied after curing, it can be added before or after the first cladding 150, second cladding 152, third cladding 153, fourth cladding 154, leading edge sheath 156, root cladding 158, or any combination thereof. Coating 230 can be applied to portions of the outer skin surface 142 not covered by cladding 120. Alternatively, in another, non-limiting example, coating 230 can be applied to a portion of cladding 120 or between the outer skin surface 142 and cladding 120.

[0169] The benefits of this disclosure include the use of a braided core, a laminated skin, and a cladding that provide weight. The cladding can be a material bonded to the outside of the laminated skin, which covers the braided core. Cladding at the tip or trailing edge can be a composite material, a thermoplastic material, or a metallic material, wherein the cladding provides improved stiffness to the airfoil to enhance robustness against deflection under extreme events such as ingestion events.

[0170] When applied using discrete components, the cladding shown as the first, second, and third cladding can be customized based on size, application, environment, or location within the turbine engine.

[0171] Preparing one or more surfaces of the cladding before it is adhered to the composite airfoil improves the adhesion of the cladding when it is applied to the composite airfoil.

[0172] Within the scope not described herein, various features and structures of the various embodiments may be combined or substituted for each other as needed. The fact that a feature is not shown in all embodiments is not to be construed as meaning it cannot be shown in this way, but is done for the sake of brevity. Therefore, various features of different embodiments may be mixed and matched as needed 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 this disclosure.

[0173] This written description uses examples to disclose this disclosure, including best practices, and also enables any person skilled in the art to practice this disclosure, including making and using any device or system and performing any combination of methods. The patentable scope of this disclosure is defined by the claims, but may include other examples that would occur to a person skilled in the art. Such other examples are intended to fall within the scope of the claims if they include structural elements that are not indistinguishable from the literal language of the claims, or if they include equivalent structural elements that are not substantially different from the literal language of the claims.

[0174] Further details are provided by the following topics:

[0175] A composite airfoil assembly for a gas turbine engine, the composite airfoil assembly comprising: a core, the core including a composite structure defining an exterior of the core; a skin having an inner skin surface and an outer skin surface connected to at least a portion of the exterior of the core; and a cladding having an inner cladding surface connected to the outer skin surface, wherein at least a portion of the inner cladding surface has an arithmetic mean roughness (Ra) of at least 100.

[0176] The composite airfoil assembly according to any of the foregoing clauses further includes an adhesive that bonds a portion of the inner surface of the cladding to a portion of the outer surface of the skin.

[0177] According to any of the preceding clauses, the composite airfoil assembly is wherein the cladding is attached to the outer surface of the skin adjacent to at least one of the trailing edge, leading edge, root, or tip of the composite airfoil defined by the core and the skin.

[0178] According to any of the preceding clauses, the composite airfoil assembly includes a first cladding having a first inner surface and a second cladding having a second inner surface, wherein the Ra of the first inner surface is different from the Ra of the second inner surface.

[0179] According to any of the preceding clauses, the composite airfoil assembly is wherein the cladding is attached to the outer surface of the skin adjacent to at least one of the trailing edge, leading edge, root, or tip of the composite airfoil defined by the core and the skin.

[0180] According to any of the foregoing clauses, the arithmetic mean roughness (Ra) of the inner surface of the cladding is in the range of 150 to 400.

[0181] The composite airfoil assembly according to any of the foregoing clauses, wherein the cladding comprises a metallic material.

[0182] The composite airfoil assembly according to any of the foregoing clauses, wherein the metallic material includes titanium, nickel, nickel alloy, steel, tin, aluminum, brass, copper or any combination thereof.

[0183] The composite airfoil assembly according to any of the foregoing clauses, wherein the core is a woven core comprising a three-dimensional woven pattern.

[0184] In any of the preceding clauses, the composite airfoil assembly is wherein the skin is a laminated skin applied to the braided core.

[0185] A method of forming a composite airfoil assembly, the method comprising: forming a core using a composite material, wherein the core includes a woven structure; applying a skin to at least a portion of the exterior of the core to define a composite airfoil, the skin having an inner skin surface and an outer skin surface; curing the composite airfoil; forming a cladding having an outer cladding surface and an inner cladding surface; preparing at least a portion of the inner cladding surface to have an arithmetic mean roughness (Ra) of 100 or greater; and adhering the cladding to a pressure side or a suction side at a trailing edge, leading edge, or tip of the composite airfoil, wherein the inner cladding surface faces a portion of the outer skin surface.

[0186] According to any of the foregoing clauses, the method of preparing at least a portion of the inner surface of the cladding includes preparing at least a portion of the inner surface of the cladding to have an arithmetic mean roughness (Ra) in the range of 150 to 400.

[0187] According to the method described in any of the foregoing clauses, the at least portion of the inner surface of the cladding includes a material removal depth in the range of 0.025 mm to 0.127 mm.

[0188] The method according to any of the foregoing clauses further includes etching the at least portion of the cladding before preparing the at least portion of the inner surface of the cladding.

[0189] According to any of the foregoing clauses, the method of preparing at least a portion of the inner surface of the cladding includes preparing at least a portion of the inner surface of the cladding to have an arithmetic mean roughness (Ra) in the range of 150 to 400.

[0190] The method according to any of the foregoing clauses further includes applying an undercoat to one or more surfaces of the cladding before applying the cladding or after adhering the cladding.

[0191] According to any of the foregoing provisions, the method of adhering the cladding includes applying an adhesive to the inner surface of the cladding, the outer surface of the epidermis, or both the inner surface of the cladding and the outer surface of the epidermis.

[0192] The method according to any of the foregoing clauses, wherein forming the cladding includes additive manufacturing, machining or casting of metallic material to form the cladding.

[0193] According to any of the foregoing clauses, the preparation of at least a portion of the inner surface of the cladding results in the morphology of the first inner surface of the cladding, the second inner surface of the cladding, the third inner surface of the cladding, the fourth inner surface of the cladding, the inner surface of the sheath, or the inner surface of the root having at least 10 random peaks per 645 square millimeters (approximately 1 square inch).

[0194] According to any of the foregoing provisions, forming the coating includes forming a first coating and a second coating, and adhering the coating includes adhering the first coating at the trailing edge to the pressure side and adhering the second coating at the trailing edge to the suction side.

[0195] According to any of the foregoing provisions of the method, applying the skin to the exterior of the core includes applying a laminated skin.

[0196] The composite airfoil assembly according to any of the foregoing clauses, wherein the cladding has a cladding bulk modulus greater than the core bulk modulus.

[0197] The composite airfoil assembly according to any of the foregoing clauses, wherein the cladding comprises titanium, nickel, steel, tin, aluminum, brass, copper, or any combination thereof.

[0198] According to any of the preceding clauses, the composite airfoil assembly includes a first cladding and a second cladding located at the trailing edge of the composite airfoil, wherein the first cladding and the second cladding are located on opposite sides of the composite airfoil.

[0199] The composite airfoil assembly according to any of the foregoing clauses further includes a third cladding layer located at the tip of the composite airfoil.

[0200] The composite airfoil assembly according to any of the foregoing clauses, wherein the cladding includes a leading-edge sheath located at the leading edge of the composite airfoil.

[0201] The composite airfoil assembly according to any of the foregoing clauses, wherein the leading edge sheath comprises one or more of a composite material, a thermoplastic material, or a metallic material.

[0202] The composite airfoil assembly according to any of the foregoing clauses, wherein the core is a woven core comprising a three-dimensional woven pattern.

[0203] In any of the preceding clauses, the composite airfoil assembly wherein the skin is a laminated skin.

[0204] According to any of the foregoing clauses, the composite airfoil assembly, wherein the three-dimensional woven pattern includes a woven pattern.

[0205] The method according to any of the foregoing clauses further includes impregnating the braided core with resin prior to curing.

Claims

1. A composite airfoil assembly for a gas turbine engine, characterized by, The composite airfoil assembly comprises: a core comprising a composite structure defining a core exterior; a skin having a skin interior surface coupled to at least a portion of the core exterior and a skin exterior surface; and a cap layer having a cap layer interior surface coupled to the skin exterior surface, wherein at least a portion of the cap layer interior surface has an arithmetic average roughness (Ra) of at least 100, wherein the cap layer comprises a first cap layer having a first interior surface and a second cap layer having a second interior surface, wherein the arithmetic average roughness (Ra) of the first interior surface is different than the arithmetic average roughness (Ra) of the second interior surface.

2. The composite airfoil assembly of Claim 1, wherein, further comprising an adhesive bonding the cap layer interior surface to a portion of the skin exterior surface.

3. The composite airfoil assembly of Claim 2, wherein, wherein, the cap layer is coupled to the skin exterior surface proximate at least one of a trailing edge, a leading edge, a root, or a tip of the composite airfoil defined by the core and the skin.

4. The composite airfoil assembly of Claim 1, wherein, wherein, the cap layer is coupled to the skin exterior surface proximate at least one of a trailing edge, a leading edge, a root, or a tip of the composite airfoil defined by the core and the skin.

5. The composite airfoil assembly of Claim 1, wherein, wherein, the arithmetic average roughness (Ra) of the cap layer interior surface is in a range of 150 to 400.

6. The composite airfoil assembly of Claim 1, wherein, wherein, the cap layer comprises a metallic material.

7. The composite airfoil assembly of Claim 6, wherein, wherein, the metallic material comprises titanium, nickel, a nickel alloy, steel, tin, or aluminum.

8. The composite airfoil assembly of Claim 6, wherein, wherein, the metallic material comprises brass.

9. The composite airfoil assembly of Claim 6, wherein, wherein, the metallic material comprises copper.

10. The composite airfoil assembly of Claim 1, wherein, wherein, the core is a woven core comprising a three-dimensional weave pattern.

11. The composite airfoil assembly of Claim 10, wherein, wherein, the skin is a laminated skin applied to the woven core.

12. The composite airfoil assembly of Claim 1, wherein, wherein, the first interior surface or the second interior surface comprises a topography having at least 10 random peaks per 645 square millimeters.

13. A method of forming a composite airfoil assembly, characterized by, The method comprises: forming a core using a composite material, wherein the core comprises a woven structure; applying a skin to at least a portion of an exterior of the core to define a composite airfoil, the skin having a skin interior surface and a skin exterior surface; curing the composite airfoil; forming a cap layer having a cap layer exterior surface and a cap layer interior surface; preparing at least a portion of the cap layer interior surface to have an arithmetic average roughness (Ra) of 100 or greater, wherein preparing the at least a portion of the cap layer interior surface results in a topography of a first cap layer interior surface, a second cap layer interior surface, a third cap layer interior surface, a fourth cap layer interior surface, a sheath interior surface, or a root interior surface having at least 10 random peaks per 645 square millimeters; and adhering the cap layer to a pressure side or a suction side at a trailing edge, a leading edge, or a tip of the composite airfoil, wherein the cap layer interior surface faces a portion of the skin exterior surface.

14. The method of claim 13, wherein, wherein, preparing the at least a portion of the cap layer interior surface comprises preparing the at least a portion of the cap layer interior surface to have an arithmetic average roughness (Ra) in a range of 150 to 400.

15. The method of claim 13, wherein, wherein, the at least a portion of the cap layer interior surface comprises a material removal depth in a range of 0.025 millimeters to 0.127 millimeters.

16. The method of claim 13, wherein, further comprising etching the at least a portion of the cap layer prior to preparing the at least a portion of the cap layer interior surface.

17. The method of claim 16, wherein, wherein, preparing the at least a portion of the cap layer interior surface comprises preparing the at least a portion of the cap layer interior surface to have an arithmetic average roughness (Ra) in a range of 150 to 400.

18. The method of claim 13, wherein, further comprising applying a primer to one or more surfaces of the cap layer prior to applying the cap layer or after adhering the cap layer.

19. The method of claim 13, wherein, wherein, adhering the cap layer comprises applying an adhesive to the cap layer interior surface, the skin exterior surface, or the cap layer interior surface and the skin exterior surface.

20. The method of claim 13, wherein, wherein, forming the cap layer comprises additive manufacturing, machining, or casting a metallic material to form the cap layer.

21. The method of claim 13, wherein, wherein, forming the cap layer comprises forming a first cap layer and a second cap layer, and adhering the cap layer comprises adhering the first cap layer at the trailing edge to the pressure side and adhering the second cap layer at the trailing edge to the suction side.

22. A composite airfoil assembly for a gas turbine engine, characterized by, The composite airfoil assembly comprises: a core comprising a composite structure defining a core exterior; a skin having a skin interior surface coupled to at least a portion of the core exterior and a skin exterior surface; and a cap layer having a cap layer interior surface coupled to the skin exterior surface, wherein at least a portion of the cap layer interior surface has an arithmetic average roughness (Ra) of at least 100, wherein the at least a portion of the cap layer interior surface results in a topography having at least 10 random peaks per 645 square millimeters.

Images