Composite airfoil assembly for turbine engine

Abstract

A composite airfoil assembly for a turbine engine. A composite airfoil assembly includes a braided core having a core exterior. The braided core includes a first region, a second region, and a transition region. A skin is applied to at least a portion of the braided core at an exterior of the core.

Description

Technical Field

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

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

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

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

[0005] Figure 1 is a schematic cross-sectional view of a turbine engine.

[0006] Figure 2 is a schematic perspective view of an aircraft including an unducted or open rotor turbine engine.

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

[0008] Figure 4 According to an exemplary embodiment of the present disclosure, Figure 3 A schematic cross-sectional view taken along line IV-IV shows the interior of the composite airfoil assembly, including the braided core and laminated skins.

[0009] Figure 5 According to an exemplary embodiment of the present disclosure Figure 3 An enlarged view of the composite airfoil assembly.

[0010] Figure 6 According to an exemplary embodiment of the present disclosure, Figure 5 An enlarged schematic cross-section of a portion of a composite airfoil taken along line VI-VI, including a portion of a first region, a first transition region, and a portion of a second region. DETAILED DESCRIPTION

[0011] Aspects disclosed herein relate to a composite component to be used in an engine component of a turbine engine. The composite component is shown as a composite airfoil assembly having a composite airfoil including at least a core and a skin shown as a braided core and a laminated skin. The composite airfoil assembly may be located in a fan section of a turbine engine. The laminated skin has a skin outer surface, wherein a first portion of the skin outer surface defines a first portion of an airfoil outer surface of the composite airfoil assembly.

[0012] The benefits of the composite airfoil assembly include weight savings of the composite airfoil. The smooth outer surface of the airfoil provides less disruption to the air passing through the composite airfoil assembly. That is, the outer surface of the airfoil does not include positive bumps. The improved aerodynamics results in an overall better airflow, which improves the efficiency of the engine.

[0013] It should be understood that, in a non-limiting example, the present disclosure is applicable to other engine components of a turbine engine, not just airfoils, such as disks or combustor liners. In addition, although described in terms of a core for making an airfoil, it should be understood that the present disclosure is applicable to any other suitable environment.

[0014] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter references to refer to features in the drawings. Like or similar reference numbers in the drawings and description have been used to refer to like or similar parts of the present disclosure.

[0015] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. In addition, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

[0016] As used herein, the terms “first,” “second,” “third,” or “fourth” may be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0017] The term "at least one" in a context such as "at least one of A, B, and C" means only A, only B, only C, or any combination of A, B, and C.

[0018] The terms “turbomachine” or “turbomachine” refer to a machine that includes one or more compressors, a heat generating section (eg, a combustion section), and one or more turbines that together generate a torque output.

[0019] The term "turbine engine" refers to an engine having a turbine as all or part of its power source. Example turbine engines include turbofan engines, turboprop engines, turbojets, turboshaft engines, etc., as well as hybrid electric versions of one or more of these engines.

[0020] The terms "front" and "rear" refer to relative positions within a turbine engine or vehicle and refer to the normal operating attitude of the turbine engine or vehicle. For example, for a turbine engine, front refers to a position closer to the engine inlet, while rear refers to a position closer to the engine nozzle or exhaust.

[0021] As used herein, the term "upstream" refers to the direction opposite to the direction of fluid flow, while the term "downstream" refers to the direction in the same direction as the direction of fluid flow. The terms "front" or "forward" mean in front of something, and "rear" or "rearward" mean behind something. For example, when used in relation to fluid flow, front / forward may mean upstream, while rear / rearward may mean downstream.

[0022] As used herein, the terms "axial" and "axially" refer to directions and orientations extending substantially parallel to the centerline of a turbine engine. In addition, the terms "radial" and "radially" refer to directions and orientations extending substantially perpendicular to the centerline of a turbine engine. For example, in the context of a turbine engine as a whole, radial refers to directions along a ray extending between the central longitudinal axis of the engine and the periphery of the engine. In addition, as used herein, the terms "circumferential" and "circumferentially" refer to directions and orientations extending in an arc around the centerline of a turbine engine.

[0023] Unless otherwise specified herein, the terms "coupled," "fixed," "applied to," "attached to," and the like refer to both direct coupling, fixing, or attachment as well as indirect coupling, fixing, or attachment through one or more intermediate components or features.

[0024] All directional references (e.g., radial, axial, up, down, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, backward, etc.) are used only for identification purposes to help the reader understand the present disclosure, and do not create limitations, particularly limitations on the position, orientation, or use of aspects of the present disclosure described herein. Unless otherwise stated, connection references (e.g., attachment, connection, connection, and engagement) will be broadly interpreted and may include intermediate structural elements between sets of elements and relative movement between elements. Therefore, connection references do not necessarily infer that two elements are directly connected and have a fixed relationship to each other. The exemplary drawings are only for illustrative purposes, and the dimensions, positions, orders, and relative sizes reflected in the attached drawings may vary.

[0025] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise. Also, as used herein, the term "set" or a "set of" elements may refer to any number of elements, including only one.

[0026] As used herein, the term "stiffness" may be used to define the degree to which a structure resists deformation in response to a force. Stiffness may be defined as the ratio of the force to the displacement of an object under the action of a force. Stiffness may include resistance to deformation in response to forces applied from various directions, so that in non-limiting examples, stiffness may refer to axial stiffness, tensile stiffness, compressive stiffness, torsional stiffness, or shear stiffness.

[0027] As used herein, the term "composite" refers to a component having two or more materials. A composite can be a combination of at least two or more metals, non-metals, or a combination of metal and non-metal elements or materials. Examples of composite materials can be, but are not limited to, polymer matrix composites (PMC), ceramic matrix composites (CMC), metal matrix composites (MMC), carbon fibers, polymer resins, thermoplastics, bismaleimide (BMI) materials, polyimide materials, epoxy resins, glass fibers, and silicon matrix materials.

[0028] As used herein, a "composite" component refers to a structure or component that includes any suitable composite material. A composite component, such as a composite airfoil, may include several layers or several layups of composite materials. The stiffness, material, and dimensions of the layers or layups may vary to achieve a desired composite component or composite portion of a component having a predetermined weight, size, stiffness, and strength.

[0029] One or more adhesive layers may be used to form or join composite components. The adhesive may include resins and phenolics, where the adhesive may require curing at elevated temperatures or other hardening techniques.

[0030] In the present disclosure, when a layer is described as being "on" or "over" another layer or substrate, it should be understood that the layers may be directly in contact with each other or have another layer or feature between the layers unless explicitly stated to the contrary. Therefore, these terms simply describe the relative position of layers with respect to each other and do not necessarily mean "on top of...", as the relative position of above or below depends on the orientation of the device with respect to the viewer.

[0031] As used herein, PMC refers to a class of materials. As an example, PMC materials are partially defined by prepregs, which are reinforcing materials pre-impregnated with polymer matrix materials (e.g., thermoplastic resins). Non-limiting examples of processes for producing thermoplastic prepregs include: hot melt prepregs, in which fiber reinforcements are pulled through a molten bath of resin; and powder prepregs, in which resin is deposited on the fiber reinforcements, as a non-limiting example, electrostatically deposited on the fiber reinforcements, and then adhered to the fibers, as a non-limiting example, in an oven or with the help of heated rollers. Prepregs can be in the form of unidirectional tapes or woven fabrics, which are then stacked on top of each other to form the desired number of stacked layers of a part.

[0032] Multilayer prepreg is stacked to the appropriate thickness and orientation of composite parts, and then resin is cured and solidified to provide fiber reinforced composite parts. The resin for PMC matrix material can be generally classified as thermosetting resin or thermoplastic resin. Thermoplastic resin is generally classified as a polymer that can be repeatedly softened and flowed when heated and can be hardened due to physical changes rather than chemical changes when fully cooled. The famous example category of thermoplastic resin includes nylon, thermoplastic polyester, polyaryletherketone and polycarbonate resin. The specific example of high-performance thermoplastic resin for aerospace applications has been envisioned to include polyetheretherketone (PEEK), polyetherketoneketone (PEKK), polyetherimide (PEI), polyaryletherketone (PAEK) and polyphenylene sulfide (PPS). On the contrary, once fully cured into a hard rigid solid, thermosetting resin will not experience significant softening when heated, but thermally decomposes when fully heated. The famous example of thermosetting resin includes epoxy resin, bismaleimide (BMI) and polyimide resin.

[0033] Instead of using prepreg, in another non-limiting example, by using thermoplastic polymers, braided fabrics can be utilized. Braided fabrics may include, but are not limited to, dry carbon fibers woven together with thermoplastic polymer fibers or filaments. Non-prepreg braided architectures can be made in a similar manner. In this way, the fiber volume of a part can be customized by specifying the relative concentrations of thermoplastic fibers and reinforcing fibers that have been woven or braided together. In addition, different types of reinforcing fibers can be braided or braided together at different concentrations to customize the properties of a part. For example, glass fibers, carbon fibers, and thermoplastic fibers can all be woven together at different concentrations to customize the properties of a part. Carbon fibers provide the strength of the system, glass fibers can be incorporated to enhance impact properties, which is a design feature of a part located near the engine inlet, and thermoplastic fibers provide bonding for reinforcing fibers.

[0034] In yet another non-limiting example, resin transfer molding (RTM) can be used to form at least a portion of a composite component. Typically, RTM includes applying a dry fiber or matrix material to a mold or cavity. The dry fiber or matrix material may include prepregs, woven materials, braided materials, or any combination thereof.

[0035] The resin may be pumped or otherwise provided to the mold or cavity to impregnate the dry fibers or matrix material. The combination of the impregnated fibers or matrix material and the resin is then cured and removed from the mold. When removed from the mold, the composite part may require a post-curing process.

[0036] It is contemplated that RTM may be a vacuum assisted process. That is, the air in the cavity or mold may be removed and replaced with resin prior to heating or curing. It is further contemplated that placement of dry fiber or matrix material may be manual or automated.

[0037] The dry fiber or matrix material may be shaped to form the composite part or to guide the resin. Optionally, additional layers or reinforcements of a material different from the dry fiber or matrix material may also be included or added prior to heating or curing.

[0038] As used herein, CMC refers to a class of materials having reinforcing fibers in a ceramic matrix. Typically, the reinforcing fibers provide structural integrity to the ceramic matrix. Some examples of reinforcing fibers may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), non-oxide carbon-based materials (e.g., carbon), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (Al2O3), and silicon nitride (SiO2). 2 O 3 ), silicon dioxide (SiO 2 ), aluminosilicates (such as mullite), or mixtures thereof), or mixtures thereof.

[0039] Some examples of ceramic matrix materials may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (Al 2 O 3 ), silicon dioxide (SiO 2 ), aluminosilicates or mixtures thereof), or mixtures thereof. Optionally, ceramic components (e.g., oxides of Si, Al, Zr, Y and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite and montmorillonite) may also be included in the ceramic matrix.

[0040] In general, specific CMCs can be referred to as their fiber type / matrix type combination. For example, C / SiC is carbon fiber reinforced silicon carbide, SiC / SiC is silicon carbide fiber reinforced silicon carbide, SiC / SiN is silicon carbide fiber reinforced silicon nitride, SiC / SiC-SiN is silicon carbide fiber-reinforced silicon carbide / silicon nitride matrix mixture, etc. In other examples, CMCs can be made of materials including oxide-based materials such as aluminum oxide (Al2O3). 2 O 3 ), silicon dioxide (SiO 2 ), aluminosilicates and mixtures thereof) and a matrix and reinforcing fibers. Aluminosilicates may include crystalline materials (e.g., mullite (3Al 2 O 3 ·2SiO 2 )), and glassy aluminosilicates.

[0041] In certain non-limiting examples, the reinforcing fibers may be bundled and / or coated before being included in the ceramic matrix. For example, the fiber bundle may be formed into a reinforcing tape, such as a unidirectional reinforcing tape. Multiple tapes may be stacked together to form a preform component. The fiber bundle may be impregnated with a slurry composition before or after forming the preform. The preform may then be subjected to heat treatment and subsequent chemical treatment (such as silicon melt infiltration) to obtain a component formed of a CMC material having a desired chemical composition. For example, the preform may undergo solidification or burnout to produce a high coke residue in the preform, and then undergo melt infiltration with silicon, or undergo solidification or pyrolysis to produce a silicon carbide matrix in the preform, and then undergo chemical vapor infiltration with silicon carbide. Additional steps may be taken to improve the densification of the preform, by injecting the preform with a liquid resin or polymer before or after chemical vapor infiltration, and then performing a heat treatment step to fill the voids with silicon carbide. CMC materials as used herein may be formed using any known or later developed method including, but not limited to, melt infiltration, chemical vapor infiltration, polymer infiltration pyrolysis (PIP), or any combination thereof.

[0042] The reinforcing fibers may be at least some portions of individual filaments or strands. As used herein, "ceramic fiber tow," "fiber tow," or simply "tow" refers to a bundle of multiple individual fibers, filaments, or loose strands. The filaments of the tow may be randomly mixed or arranged in a pattern, and / or may be continuous or discontinuous. For example, the tow may include broken filaments or filament segments. As another example, the filaments of the tow may be substantially parallel, twisted, or otherwise arranged. The tow may function in substantially the same manner as a single or individual filament. It will also be understood that "individual ceramic filaments" or simply "individual filaments" used herein refers to a single or non-bundled elongated ceramic member.

[0043] Such materials, along with certain monolithic ceramics (i.e., ceramic materials without reinforcement materials), are particularly suitable for higher temperature applications. In addition, these ceramic materials are lightweight compared to superalloys, but still provide strength and durability to the parts made from them. As a result, such materials are currently being considered for use in many turbine components used in the higher temperature sections of turbine engines (such as airfoils (e.g., turbine blades and vanes), combustors, shrouds, etc.), which would benefit from the lighter weight and higher temperature capabilities that these materials can provide.

[0044] The term "metal" as used herein refers to materials including metals such as, but not limited to, titanium, iron, aluminum, stainless steel, brass, copper, and nickel alloys. A metal material or alloy may be a combination of at least two or more elements or materials, at least one of which is a metal. As used herein, the term "additive manufacturing" generally refers to a manufacturing process in which a feedstock of a material in the form of a particulate powder or wire is aggregated to form a three-dimensional component. The feedstock material is then melted by applying heat or other curing treatments to form an integral single component that may have various integral subcomponents. As used herein, integral refers to a single structure in which the individual layers lose their characteristics in the final single structure by virtue of the material of each layer being melted to or with the material of an adjacent layer and lacking an interface or joint.

[0045] Suitable additive manufacturing techniques according to the present disclosure include, for example, directed energy deposition (DED), fused deposition modeling (FDM), selective laser sintering (SLS), 3D printing such as by inkjet and laser jetting, stereolithography (SLA), direct selective laser sintering (DSLS), electron beam sintering (EBS), electron beam melting (EBM), laser engineered net shaping (LENS), laser net shape manufacturing (LNSM), direct metal deposition (DMD), digital light processing (DLP), direct selective laser melting (DSLM), selective laser melting (SLM), direct metal laser melting (DMLM) and other known processes.

[0046] Digital light processing (DLP) can include a 3D DLP printer having a transparent vat or tank, a building platform, and a light assembly. The vat or tank can contain, for example, a photopolymer resin.

[0047] The DLP build platform may be coupled to, for example, a motor or other mechanism that allows the build platform to move in one or more dimensions, such as raising or lowering the build platform from or toward resin in a vat or tank.

[0048] The DLP printing component can be coupled to a lower portion of the building platform facing the barrel or tank. The lighting assembly is at least partially located below the barrel or tank. The lighting assembly may include at least one light source and at least one optical reflector or refractor, such as a deflecting mirror or at least one lens.

[0049] A controller coupled to or included in a DLP printer can control one or more aspects of the DLP printer, such as the position of a DLP build platform or the intensity, duration, or orientation of a light source.

[0050] In addition to using a direct metal laser sintering (DMLS), direct metal laser melting (DMLM) process, or electron beam melting (EBM) process in which an energy source is used to selectively sinter or melt portions of a powder layer, it will be appreciated that, according to alternative aspects of the present disclosure, the additive manufacturing process may be a "binder jetting" process. In this regard, binder jetting involves the continuous deposition of layers of additive powder in a manner similar to that described above. However, instead of using an energy source to generate an energy beam to selectively melt or fuse the additive powder, binder jetting involves selectively depositing a liquid binder onto each powder layer. The liquid binder may be, for example, a photocurable polymer or another liquid binder. Other suitable additive manufacturing methods and variations are intended to fall within the scope of the present subject matter.

[0051] Figure 1 is a schematic cross-sectional view of a turbine engine, particularly an open rotor or turbine engine 10 for an aircraft. The turbine engine 10 has a generally longitudinally extending axis or engine centerline 12 extending from a forward end 14 to an aft end 16. The turbine engine 10 includes, in downstream serial flow relationship, a set of circumferentially spaced blades or propellers defining: a fan section 18 including a fan 20; a compressor section 22 including a supercharger or low pressure (LP) compressor 24 and a high pressure (HP) compressor 26; a combustion section 28 including a combustor 30; a turbine section 32 including an HP turbine 34 and an LP turbine 36; and an exhaust section 38. The turbine engine 10 as described herein is a non-limiting example, and other architectures are possible, such as, but not limited to, a steam turbine engine, a supercritical carbon dioxide turbine engine, or any other suitable turbine engine.

[0052] The exterior surface of the turbine engine 10 defined by a casing (such as a nacelle 40) extends from the front end 14 of the turbine engine 10 toward the rear end 16 of the turbine engine 10 and covers at least a portion of the compressor section 22, the combustion section 28, the turbine section 32, and the exhaust section 38. The fan section 18 may be positioned at the front of the nacelle 40 and extends radially outward from the nacelle 40 of the turbine engine 10, specifically, the fan section 18 extends radially outward from the nacelle 40. The fan section 18 includes a set of fan blades 42 and a set of stationary fan blades 82 downstream of the set of fan blades 42, both of which are radially arranged around the engine centerline 12. The turbine engine 10 includes any number of one or more sets of rotating blades or propellers (e.g., the set of fan blades 42) arranged upstream of the set of stationary fan blades 82. As a non-limiting example, the turbine engine 10 may include multiple sets of fan blades 42 or the set of stationary fan blades 82. Therefore, the turbine engine 10 is further defined as a non-ducted single fan turbine engine. The 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 of the combustion section 28, downstream of the combustion section 28, or aligned with the axial positioning of the combustion section 28.

[0053] The compressor section 22 , the combustion section 28 , and the turbine section 32 are collectively referred to as an engine core 44 , which generates combustion gases. The engine core 44 is surrounded by an engine casing 46 operably coupled to a portion of a nacelle 40 of the turbine engine 10 .

[0054] An HP shaft or spool 48 disposed coaxially about the engine centerline 12 of the turbine engine 10 drivingly connects the HP turbine 34 to the HP compressor 26. An LP shaft or spool 50 disposed coaxially about the engine centerline 12 of the turbine engine 10 within the larger diameter annular HP spool 48 drivingly connects the LP turbine 36 to the LP compressor 24 and the fan 20. The spools 48, 50 are rotatable about the engine centerline 12 and are coupled to a set of rotatable elements that collectively define a rotor 51.

[0055] It should be appreciated that turbine engine 10 is a direct drive or integral drive engine utilizing a reduction gearbox coupling LP shaft or spool 50 to fan 20 .

[0056] The LP compressor 24 and the HP compressor 26 each include a set of compressor stages 52, 54, wherein a set of compressor blades 56, 58 rotate relative to a corresponding set of static compressor vanes 60, 62 (also referred to as nozzles) to compress or pressurize the fluid flow passing through the stage. In a single compressor stage 52, 54, a plurality of compressor blades 56, 58 are arranged in a ring and extend radially outward from a blade platform to a blade tip relative to the engine centerline 12, while the corresponding static compressor vanes 60, 62 are positioned upstream of and adjacent to the compressor blades 56, 58. Notably, Figure 1 The numbers of blades, buckets, and compressor stages shown in FIG. 5 are selected for illustration purposes only, and other numbers are possible.

[0057] Compressor blades 56, 58 for one stage of the compressor are mounted to disks 61 mounted to respective ones of the HP spool 48 and the LP spool 50, with each stage having its own disk 61. Static compressor vanes 60, 62 for one stage of the compressor are mounted to the engine case 46 in a circumferential arrangement.

[0058] The HP turbine 34 and the LP turbine 36 each include a set of turbine stages 64, 66 in which a set of turbine blades 68, 70 rotate relative to a corresponding set of static turbine vanes 72, 74 (also referred to as nozzles) to extract energy from the fluid flow passing through the stage. In a single turbine stage 64, 66, a plurality of turbine blades 68, 70 are arranged in a ring and extend radially outward from the blade platform to the blade tip relative to the engine centerline 12, while the corresponding static turbine vanes 72, 74 are positioned upstream of and adjacent to the turbine blades 68, 70. Notably, Figure 1 The number of blades, buckets, and turbine stages shown in FIG. 5 are selected for illustrative purposes only, and other numbers are possible.

[0059] Turbine blades 68, 70 for one stage of turbine section 32 are mounted to disks 71 mounted to a respective one of HP spool 48 and LP spool 50, with each stage having a dedicated disk 71. Static turbine buckets 72, 74 for one stage of turbine section 32 are mounted to engine casing 46 in a circumferential arrangement.

[0060] The rotating portions of turbine engine 10, such as blades 56, 58, 68, 70 in compressor section 22 and turbine section 32, are also referred to individually or collectively as rotors 51. Thus, a rotor refers to the combination of rotating elements throughout turbine engine 10.

[0061] Complementing the rotor portion, the stationary portion of the turbine engine 10, such as the stationary blades 60, 62, 72, 74 in the compressor section 22 and the turbine section 32, are also individually or collectively referred to as stators 63. Thus, the stator 63 refers to the combination of non-rotating elements throughout the turbine engine 10.

[0062] The nacelle 40 is operably coupled to the turbine engine 10 and covers at least a portion of the engine core 44, the engine case 46, or the exhaust section 38. At least a portion of the nacelle 40 extends axially forward or upstream of the position shown. For example, the nacelle 40 extends axially forward so that a portion of the nacelle 40 covers or covers a portion of the fan section 18 or the supercharger section (not shown) of the turbine engine 10.

[0063] During operation of the turbine engine 10, a free stream airflow 80 flows against the front of the turbine engine 10. A portion of the free stream airflow 80 enters the annular region 25 defined by the swept area between the outer surface of the nacelle 40 and the tips of the blades, where the 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.

[0064] More specifically, a working airflow 76 flows into the LP compressor 24, which then pressurizes the working airflow 76, thereby defining a pressurized airflow that is supplied to the HP compressor 26, which further pressurizes the air. The working airflow 76 or pressurized airflow from the HP compressor 26 is mixed with fuel in the combustor 30 and ignited, thereby generating combustion gases. The HP turbine 34 extracts some work from these gases, which drives the HP compressor 26. The combustion gases are discharged into the LP turbine 36, which extracts additional work to drive the LP compressor 24, and the working airflow 76 or exhaust gas is ultimately discharged from the turbine engine 10 via the exhaust section 38. The drive of the LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP compressor 24. The working airflow 76, including the pressurized airflow and the combustion gases, defines a working airflow that flows through the compressor section 22, the combustion section 28, and the turbine section 32 of the turbine engine 10.

[0065] The inlet airflow 78 flows through the set of fan blades 42 and over the nacelle 40 of the turbine engine 10. The inlet airflow 78 then flows over at least a portion of the set of stationary fan blades 82, which directs the inlet airflow 78 so that it is transversely toward the engine centerline 12. The inlet airflow 78 then follows the curvature of the nacelle 40 and flows through the set of stationary fan blades 82 toward the exhaust section 38. The pylon 84 mounts the turbine engine 10 to an external structure (e.g., a fuselage, wing, tail, etc. of an aircraft).

[0066] The working airflow 76 and at least some of the inlet airflow 78 merge downstream of the exhaust section 38 of the turbine engine 10. Together, the working airflow 76 and the inlet airflow 78 contribute to the overall thrust of the turbine engine 10.

[0067] It is envisioned that a portion of the working airflow 76 is extracted as bleed air 77 (e.g., from the compressor section 22). The bleed air 77 provides airflow to engine components that require cooling. The temperature of the working airflow 76 leaving the combustor 30 is significantly increased relative to the working airflow 76 within the compressor section 22. Therefore, the cooling provided by the bleed air 77 is necessary for operating these engine components in an elevated temperature environment or in the hot part of the turbine engine 10. In the case of a turbine engine, the hot part of the engine is typically downstream of the combustor 30, particularly the turbine section 32, with the HP turbine 34 being the hottest part because it is located directly downstream of the combustion section 28. Other sources of cooling fluid are, but are not limited to, fluid discharged from the LP compressor 24 or the HP compressor 26.

[0068] Figure 2 is suitable for use as Figure 1 Schematic perspective view of an aircraft 86 with a generic non-ducted turbine engine 88 of the turbine engine 10. Aircraft 86 includes a fuselage 90 having an exterior surface. At least one wing 92 and an empennage 94 extend from fuselage 90. Emitter 94 is operably coupled to and spaced apart from fuselage 90 via an empennage pylon 96. Non-ducted turbine engine 88 is operably coupled to the exterior surface of fuselage 90 via pylon 98. Non-ducted turbine engine 88 includes a set of circumferentially spaced fan blades 100. A set of stationary fan blades 102 are disposed downstream of the set of circumferentially spaced fan blades 100. Fuselage 90 extends between a nose 104 and a tail 106 and includes a fuselage centerline 108 extending therebetween.

[0069] Additionally, while the empennage 94 is a T-wing empennage (e.g., as shown in the figure), other conventional empennages are contemplated, such as a cruciform empennage, an H-tail, a triple tail, a V-tail, an inverted tail, a Y-tail, a double tail, a boom tail, or a ring tail, all of which are referred to herein as the empennage 94.

[0070] Figure 3 Is suitable for Figure 1 The turbine engine 10 or Figure 2 A schematic perspective view of a composite airfoil assembly 110 and a disk assembly 112 for use within a non-ducted turbine engine 88. The disk assembly 112 is suitable for use as a disk 61, 71 ( Figure 1) or any other disk, such as, but not limited to, a disk within the fan section 18, the compressor section 22, or the turbine section 32 of the turbine engine 10. The composite airfoil assembly 110 may be rotating or non-rotating, such that the composite airfoil assembly 110 may include static compressor blades 60, 62 ( Figure 1 ), the set of compressor blades 56, 58 ( Figure 1 ), static turbine blades 72, 74 ( Figure 1 ), the group of turbine blades 68, 70 ( Figure 1 ) or the set of fan blades 42 ( Figure 1 ). As a non-limiting example, the composite airfoil assembly 110 may be a composite fan blade assembly.

[0071] The disk assembly 112 may be rotatable or stationary about an axis of rotation 114. The axis of rotation 114 may be aligned with an engine centerline (e.g., Figure 1 The disk assembly 112 includes a plurality of slots 116 extending axially through a radially outer portion of the disk assembly 112 and spaced circumferentially about the disk assembly 112 relative to the rotational axis 114.

[0072] The composite airfoil assembly 110 includes a composite airfoil 118 and a cladding 120. The composite airfoil 118 extends between a leading edge 124 and a trailing edge 126 opposite the leading edge 124 to define a chord-wise direction. The composite airfoil 118 extends between a root 128 and a tip 130 to define a spanwise direction. An airfoil outer surface 142 of the composite airfoil assembly 110 is defined by a pressure side 132 and a suction side 134 opposite the pressure side 132. The pressure side 132 and the suction side 134 extend radially from the root 128 to the tip 130 in a radial direction. The suction side 134 may mirror the pressure side 132 and have the same surface area. However, it is contemplated that the surface area of ​​the suction side may be less than or greater than the surface area of ​​the pressure side 132.

[0073] Leading edge 124 and trailing edge 126 extend radially from root 128 to tip 130. A dovetail portion 138 may extend from composite airfoil 118.

[0074] The composite airfoil assembly 110 is coupled to the disk assembly 112 by inserting at least a portion of the dovetail portion 138 into a corresponding slot in the plurality of slots 116. The composite airfoil assembly 110 is held in place by frictional contact with the slots 116, or may be coupled to the slots 116 via any suitable coupling method, such as, but not limited to, welding, adhering, fastening, etc. Although only a single composite airfoil assembly 110 is shown, it should be understood that there may be any number of composite airfoil assemblies 110 coupled to the disk assembly 112. As a non-limiting example, there may be a number of composite airfoil assemblies 110 corresponding to the total number of slots in the plurality of slots 116.

[0075] For ease of reference, a set of relative reference directions and coordinate systems may be applied to the composite airfoil assembly 110. The circumferential direction (Cd) may extend from front to back and is shown as extending at least partially into the page. The circumferential direction (Cd) may be arranged parallel to the axis of rotation 114. The radial direction (Rd) extends perpendicular to the circumferential direction (Cd) and may extend perpendicular to the engine centerline 12. The axial direction (Ad) may be defined as being perpendicular to the radial direction (Rd) and may be defined relative to the engine centerline 12 ( Figure 1 ) along the turbine engine 10 ( Figure 1 )’s circumference.

[0076] Figure 4 It is along Figure 3 The schematic cross-sectional view taken along line IV-IV of FIG. 1 shows the interior of the composite airfoil 118 including the braided core 144 with the laminated skin 150 . That is, Figure 4 A cross-section of the composite airfoil assembly 110 is shown with the cladding 120 removed for ease of understanding.

[0077] The composite airfoil 118 includes a core, shown as a braided core 144, and a skin, shown as a laminate skin 150, wherein the laminate skin 150 is disposed on the braided core 144. That is, the laminate skin 150 surrounds or wraps at least a portion of a core exterior 156, wherein the core exterior 156 is an exterior surface of the braided core 144.

[0078] In one non-limiting example, the braided core 144 can be dry, without additional materials, or alternatively, impregnated with resin and cured. The braided core 144 can include a composite structure and / or be made of a braided structure. Such a braided structure can be a three-dimensional braided structure. More specifically, the braided structure can be circumferentially (Cd), radially (Rd), and axially (Ad) ( Figure 3 ) combination, it should be understood that the weaving pattern can be combined with the turbine engine 10 ( Figure 1) are separately formed and defined so that the weaving pattern is woven in any three mutually orthogonal planes to define a three-dimensional object relative to the planes. In a non-limiting example, the woven structure may include a three-dimensional weaving that includes a plurality of warp fibers 162 and a plurality of weft fibers 164 that can be woven in three directions to form a three-dimensional structure of the woven core 144. The three directions of the warp fibers 162 and the weft fibers 164 can be along the circumferential direction (Cd), the radial direction (Rd), and the axial direction (Ad) ( Figure 3 ) is defined by, or angled relative to. In one non-limiting example, a jacquard loom or 3D loom can be used to create a complex three-dimensional woven structure, which can include interweaving one or more composite materials to form a braided core 144. The braided core 144 can be composed of a composite material such as carbon or carbon fiber, glass or glass fiber, nylon, rayon, poly(p-phenylene terephthalamide) fiber or other aramid fiber, while other materials such as nickel, steel, titanium, metal fibers, or ceramic composites are also contemplated in non-limiting examples.

[0079] It is further contemplated that the braided core 144 can be formed into a three-dimensional braided structure having a woven or braided geometry or pattern. For example, the woven or braided geometry or pattern can include a braided pattern that includes three or more interwoven fibers woven in a repeating pattern. In another non-limiting example, the woven geometry can include a group of fibers or strands that are sequentially stacked on top of each other to define the woven geometry. The woven or braided geometry or pattern can be repeated for the entire braided core 144, or only for a portion thereof. Such additional woven geometries can be similar, where the arrangement of the fibers is the same but the orientation is different, or where the arrangement of the fibers is different and the orientation can be similar or dissimilar. The woven geometry or pattern can be formed with a composite material by a jacquard loom or a 3D loom. The three-dimensional braided structure can include a plurality of fibers in three dimensions (such as the circumferential direction (Cd), the radial direction (Rd), and the axial direction (Ad) ( Figure 3 )'s combination).

[0080] The laminate skin 150 can be formed as a set of laminate layers disposed around or around the braided core 144. In a non-limiting example, the laminate skin 150 can be a pre-impregnated, fiber-placed, or dry fiber laminate layer. In a non-limiting example, such laminate layers forming the laminate skin 150 can be formed by resin transfer molding (RTM), partial RTM, same qualified resin transfer molding (SQRTM), or out of autoclave. The braided core 144 can be made of a different material than the laminate skin 150.

[0081] Laminate skin 150 may include skin outer surface 152 and skin inner surface 154. Skin inner surface 154 may at least partially contact at least a portion of core exterior 156 of braided core 144. In other words, laminate skin 150 may be applied to at least a portion of core exterior 156 of braided core 144.

[0082] The skin thickness 158 may be measured from the skin outer surface 152 to the skin inner surface 154. Although shown as uniform, the skin thickness 158 may be in the circumferential direction (Cd), radial direction (Rd), or axial direction (Ad) ( Figure 3 ) changes.

[0083] The airfoil width 160 may be measured between the leading edge 124 and the trailing edge 126 of the composite airfoil 118. Although shown as being measured in the circumferential direction (Cd), in different and non-limiting examples, it is contemplated that the airfoil width 160 may be measured as an average length, a weighted average, a minimum length, or a maximum length measured in the axial direction (Ad) or the circumferential direction (Cd) from the leading edge 124 and the trailing edge 126 of the composite airfoil 118. In yet another different and non-limiting example, it is further contemplated that the airfoil width 160 may be measured along an average camber line 161 extending from the leading edge 124 and the trailing edge 126, wherein the average camber line is equidistant from the pressure side 132 and the suction side 134.

[0084] For example, the braided core thickness 146 may be measured between the pressure surface 133 and the suction surface 135 of the braided core 144. The pressure surface 133 is a portion of the core exterior 156 that faces the pressure side 132. The suction surface 135 is a portion of the core exterior 156 that faces the suction side 134. The braided core thickness 146 may be in the range of 5% to 30% of the airfoil width 160. The range of core thickness 146 provides a desired aerodynamic profile while maintaining weight benefits.

[0085] The skin thickness 158 may be in a range of 1% to 40% of the braided core thickness 146 . More specifically, the skin thickness 158 may be in a range of 2% to 30% of the braided core thickness 146 .

[0086] Figure 5 is an enlarged view of the composite airfoil 118. That is, Figure 5 is an enlarged view of the composite airfoil assembly 110, with the cladding 120 removed for ease of understanding ( Figure 3). The airfoil length 140 may be measured from the root 128 to the tip 130 of the composite airfoil 118. As an example, the airfoil length 140 is shown as a maximum length measured from the root 128 to the tip 130. In different and non-limiting examples, it is contemplated that the airfoil length 140 may be an average length, a weighted average, or a minimum length measured from the root 128 to the tip 130 of the composite airfoil 118 in a radial direction (Rd) defined from the root 128 to the tip 130.

[0087] The composite airfoil 118 may include a first region 170, a second region 172, a third region 174, a first transition region 176 between the first region 170 and the second region 172, and a second transition region 178 between the second region 172 and the third region 174. The first region 170 may include woven fibers formed of a first material. The first material may include one or more of carbon or carbon fibers, or non-carbon fibers such as glass or fiberglass, nylon, rayon, poly(p-phenylene terephthalamide) fibers or other aramid fibers, nickel, steel, titanium, metal fibers, or ceramic composites.

[0088] The first region 170 has a first stiffness. The first stiffness is determined by material properties, one or more braid patterns, or material properties and the braided core 144 ( Figure 4 ) is defined by a combination of one or more weave patterns of the first portion of the fabric. Material properties may include, but are not limited to, material stiffness or elastic modulus, material density, or material physical dimensions.

[0089] It is contemplated that the first stiffness may be determined based on or equal to the first material stiffness of the first material.

[0090] The first region 170 has a first length 190. The first length 190 may be measured from the root 128 to the first transition region 176 in a radial direction (Rd).

[0091] The second region 172 may include woven fibers formed of a second material. The second material may include one or more of carbon or carbon fibers, or non-carbon fibers such as glass or glass fibers, nylon, rayon, poly(p-phenylene terephthalamide) fibers or other aramid fibers, nickel, steel, titanium, metal fibers, or ceramic composites, wherein the second material is different from the first material. That is, the second region 172 includes at least one or more materials that are different from the first material in the first region 170.

[0092] The second region 172 has a second stiffness. The second stiffness is defined by a material property, one or more weaving patterns, or a combination of a material property and one or more weaving patterns forming the first portion of the braided core 144. The second stiffness is different from the first stiffness. That is, the first stiffness is greater than the second stiffness. However, in different and non-limiting examples, it is contemplated that the first stiffness may be less than the second stiffness.

[0093] It is contemplated that the second stiffness may be determined based on or equal to the second material stiffness of the second material.

[0094] The second region 172 has a second length 192. The second length 192 may be measured from the first transition region 176 to the second transition region 178 in a radial direction (Rd).

[0095] The third region 174 may include woven fibers formed of a third material. The third material may include one or more of carbon or carbon fibers, or non-carbon fibers such as glass or glass fibers, nylon, rayon, poly(p-phenylene terephthalamide) fibers or other aramid fibers, nickel, steel, titanium, metal fibers, or ceramic composites, wherein the third material is different from the second material. That is, the third region 174 includes at least one or more materials that are different from the second material in the second region 172.

[0096] The third region 174 has a third stiffness. The third stiffness is defined by a material property, one or more weaving patterns, or a combination of a material property and one or more weaving patterns forming the second portion of the braided core 144. The third stiffness is different from the second stiffness. That is, the second stiffness is greater than the third stiffness. However, in different and non-limiting examples, it is contemplated that the second stiffness may be less than the third stiffness. It is contemplated that the third stiffness may be determined based on the third material stiffness of the third material or equal to the third material stiffness of the third material.

[0097] The third region 174 has a third length 194. The third length 194 may be measured from the second transition region 178 to the tip 130 in a radial direction (Rd).

[0098] The first transition region 176 couples the first region 170 to the second region 172. That is, the first transition region 176 extends from the first region 170 to the second region 172. The first transition region 176 is defined as a portion of the composite airfoil 118 between the first region 170 and the second region 172 where the stiffness varies by more than 2% in the radial direction (Rd).

[0099] The first transition region 176 may include a first material and a second material. That is, the first transition region 176 is a portion of the composite airfoil 118 in which the braided core 144 ( Figure 4 ) fibers are interwoven and transition from a first material to a second material.

[0100] As a non-limiting example, the first material can be coupled to the second material. The coupling can include, but is not limited to, one or more of entanglement, interweaving, twisting, abutting, fastening, weaving, wrapping, splicing, braiding, or crossing of the first and second materials. The coupling of the first and second materials in the first transition region 176 can produce a gradual change in stiffness along the transition length 196 between the first radial inner boundary 173 and the first radial outer boundary 175. The first radial inner boundary 173 is located at the intersection of the first region 170 and the first transition region 176. The first radial outer boundary 175 is located at the intersection of the first transition region 176 and the second region 172. In different and non-limiting examples, it is contemplated that the gradual change in stiffness can be between the first radial inner boundary 173 and the first radial outer boundary 175 along the circumferential direction (Cd), radial direction (Rd), or axial direction (Ad) ( Figure 3 ) or any direction angled thereto. Although shown as being generally linear, the first radially inner boundary 173 and / or the first radially outer boundary 175 may be oriented in the circumferential direction (Cd), radial direction (Rd), or axial direction (Ad) ( Figure 3 ) includes contours in one or more directions.

[0101] The transition length 196 may be measured from the first region 170 to the second region 172 in the radial direction (Rd). The transition length 196 may be in the range of 1% to 200% of the first length 190 or the second length 192. As shown, as an example, the transition length 196 may be less than the first length 190 or the second length 192. That is, the transition length 196 may be in the range of 1% to less than 100% of the first length 190 or the second length 192. Alternatively, in a different and non-limiting example, the transition length 196 may be equal to or greater than the first length 190 or the second length 192. That is, the transition length 196 may be in the range of 100% to 200% of the first length 190 or the second length 192.

[0102] While the stiffness in the first transition region 176 varies, it is contemplated that the average stiffness of the first transition region 76 may be within a range between a first stiffness of the first region 170 and a second stiffness of the second region 172. The first transition region 176 provides a stiffness transition between the first region 170 and the second region 172.

[0103] As a non-limiting example, the average stiffness may be calculated by averaging the stiffness measured at predetermined intervals over the transition length 196 and / or measured at predetermined intervals from the first radially inner boundary 173 and the first radially outer boundary 175. As a further non-limiting example, the average may be calculated by averaging the stiffness at the first radially inner boundary 173 and the first radially outer boundary 175.

[0104] The second transition region 178 couples the second region 172 to the third region 174. That is, the second transition region 178 extends in the radial direction (Rd) from the second region 172 to the third region 174. The second transition region 178 is defined as a portion of the composite airfoil 118 between the second region 172 and the third region 174 where the stiffness varies by more than 2% in the radial direction (Rd).

[0105] The second transition region 178 may include the second material and the third material. That is, the second transition region 178 is a portion of the composite airfoil 118 in which the braided core 144 ( Figure 4 ) fibers are interwoven and transition from the second material to the third material.

[0106] As a non-limiting example, the second material can be coupled to the third material. The coupling can include, but is not limited to, one or more of entanglement, interweaving, twisting, abutting, fastening, weaving, wrapping, splicing, braiding, or crossing of the first and second materials. The coupling of the second material and the third material in the second transition region 178 can produce a gradual change in stiffness along the transition length 198, for example, between the second radial inner boundary 177 and the second radial outer boundary 179. The second radial inner boundary 177 is located at the intersection of the second region 172 and the second transition region 178. The second radial outer boundary 179 is located at the intersection of the second transition region 178 and the third region 174. In different and non-limiting examples, it is contemplated that the gradual change in stiffness can be between the second radial inner boundary 177 and the second radial outer boundary 179 along the circumferential direction (Cd), radial direction (Rd), or axial direction (Ad) ( Figure 3 ) or angled relative to one or more of the directions. Although shown as being generally linear, the second radially inner boundary 177 and / or the second radially outer boundary 179 may be oriented in the circumferential direction (Cd), the radial direction (Rd), or the axial direction (Ad) ( Figure 3 ) includes contours in one or more directions.

[0107] The transition length 198 may be measured from the second region 172 to the third region 174 in the radial direction (Rd). The transition length 198 may be in the range of 1% to 200% of the first length 190, the second length 192, or the third length 194. As shown, as an example, the transition length 198 may be less than the first length 190, the second length 192, or the third length 194. That is, the transition length 198 may be in the range of 1% to less than 100% of the first length 190, the second length 192, or the third length 194. Alternatively, in a different and non-limiting example, the transition length 198 may be equal to or greater than the first length 190, the second length 192, or the third length 194. That is, the transition length 198 may be in the range of 100% to 200% of the first length 190, the second length 192, or the third length 194.

[0108] While the stiffness in the second transition region 178 varies, it is contemplated that the average stiffness of the second transition region 178 may be within a range between the second stiffness of the second region 172 and the third stiffness of the third region 174. The second transition region 178 provides a stiffness transition between the second region 172 and the third region 174.

[0109] As a non-limiting example, the average stiffness may be calculated by averaging the stiffnesses measured at predetermined intervals over transition length 198 and / or measured at predetermined intervals from second radially inner boundary 177 and second radially outer boundary 179. As a further non-limiting example, the average may be calculated by averaging the stiffnesses at second radially inner boundary 177 and second radially outer boundary 179.

[0110] The gradual change in stiffness of the first transition region 176 or the second transition region 178 (which may be a smooth change or a step change in stiffness) may reduce fatigue and increase the life of the composite airfoil. In addition, the smooth change in stiffness may increase the distribution of energy from an impact event throughout the composite airfoil.

[0111] Figure 6 It is along Figure 5 An enlarged schematic cross-section of a portion of the composite airfoil 118 taken along line VI-VI further illustrating a portion of the first region 170 , the first transition region 176 , and a portion of the second region 172 .

[0112] The first group of warp fibers 180 includes a first material. The first group of warp fibers 180 is located in the first region 170, the first transition region 176, or the first region 170 and the first transition region 176. As shown, as an example, a subset 182 of the first group of warp fibers 180 is located in the first transition region 176 or extends from the first region 170 to the first transition region 176.

[0113] As an example, the first set of warp fibers 180 are shown extending from the leading edge 124 to the trailing edge 126. However, in different and non-limiting examples, it is contemplated that the first set of warp fibers 180 may extend along a circumferential direction (Cd), a radial direction (Rd), or an axial direction (Ad) ( Figure 3 ) is defined by or angled relative to one or more directions in the .

[0114] The first region 170 has a first weaving pattern. As a non-limiting example, the first weaving pattern may include a portion of a first group of warp fibers 180 that extends from the leading edge 124 to the trailing edge 126 and is located in a plane perpendicular to the plurality of weft fibers 164. The first group of warp fibers 180 may be woven in an alternating pattern above and below a group of through-thickness fibers 166 to form a first weaving pattern. Although shown as patterned, the first weaving pattern may have any weaving pattern, or an overlap, connection or winding of the first group of warp fibers 180, a plurality of weft fibers 164, and the group of through-thickness fibers 166. The first group of warp fibers 180, a plurality of weft fibers 164, and the group of through-thickness fibers 166 are woven or otherwise combined in three dimensions to form a three-dimensional structure of the braided core 144.

[0115] The first weave pattern has a warp-to-weft ratio. The warp-to-weft ratio in the first weave pattern may be in the range of 40:60 to 60:40.

[0116] The second group of warp fibers 184 includes a second material. The second group of warp fibers 184 is located in the second region 172, the first transition region 176, or the second region 172 and the first transition region 176. As shown, as an example, a subset 186 of the second group of warp fibers 184 extends to or is located in the first transition region 176.

[0117] The second set of warp fibers 184 are shown as being angled relative to the leading edge 124 and the trailing edge 126. However, in different and non-limiting examples, it is contemplated that the second set of warp fibers 184 may be arranged along a circumferential direction (Cd), a radial direction (Rd), or an axial direction (Ad) ( Figure 3 ) is defined by or angled relative to one or more directions in the .

[0118] The second region 172 has a second weaving pattern. In a non-limiting example, a portion of the second group of warp fibers 184 can extend from the leading edge 124 to the trailing edge 126 at an angle so that the second group of warp fibers 184 are angled and woven above and then below the plurality of weft fibers 164, or above and then below the group of through-thickness fibers 166. As an example, the second group of warp fibers 184 are shown as being located in a plane perpendicular to the plurality of weft fibers 164. Although shown as random, the second weaving pattern can have any weaving pattern, or overlap, coupling or entanglement of the second group of warp fibers 184, the plurality of weft fibers 164 and the group of through-thickness fibers 166. The second group of warp fibers 184, the plurality of weft fibers 164 and the group of through-thickness fibers 166 are woven or otherwise combined in three dimensions to form a three-dimensional structure of the woven core 144.

[0119] The second weaving pattern has a warp-to-weft ratio. The warp-to-weft ratio of the second weaving pattern may be in the range of 40:60 to 60:40.

[0120] As an example, the plurality of weft fibers 164 are shown as being generally made of the same material. It is contemplated that the plurality of weft fibers 164 may have a first plurality of weft fibers comprising a first material and a second plurality of weft fibers comprising a second material, wherein the transition from the first plurality of weft fibers to the second plurality of weft fibers may occur in a transition zone similar to the transition of the first group of warp fibers 180 and the second group of warp fibers 184 in the first transition region 176.

[0121] As an example, the group of through-thickness fibers 166 is shown as being generally made of the same material. It is contemplated that the group of through-thickness fibers 166 may have a first group of through-thickness fibers made of a first material and a second group of through-thickness fibers made of a second material, wherein the transition from the first group of through-thickness fibers to the second group of through-thickness fibers may occur in a transition zone similar to the transition of the first group of warp fibers 180 and the second group of warp fibers 184 in the first transition region 176.

[0122] The first transition region 176 has a transition weave pattern. The transition weave pattern may include at least some portions of the first weave pattern and the second weave pattern. However, in different and non-limiting examples, it is contemplated that the subset 182 of the first group of warp fibers 180, the subset 186 of the second group of warp fibers 184, the plurality of weft fibers 164, and the group of through-thickness fibers 166 may have a weave pattern that is different from the first weave pattern or the second weave pattern.

[0123] The transitional weave pattern may have an average warp-to-weft ratio, as the warp-to-weft ratio may vary in the radial direction (Rd), the circumferential direction (Cd), or the axial direction (Ad). The average warp-to-weft ratio in the first transition region 176 may be in the range of 40:60 to 60:40 in the radial direction (Rd).

[0124] When at least one fiber in the first group of warp fibers 180 is coupled or crossed with at least one fiber in the second group of warp fibers 184, a crossover point 188 is formed. The coupling may include, but is not limited to, one or more of entanglement, interweaving, twisting, abutting, fastening, braiding, encircling (wrapping), splicing, braiding, or crossing of at least one fiber in the subset 182 of the first group of warp fibers 180 and at least one fiber in the subset 186 of the second group of warp fibers 184.

[0125] It is contemplated that intersection 188 may define a connection location between the first weave pattern and the second weave pattern.

[0126] The intersections 188 of the fibers or weave pattern may provide control over the location of energy transfer within the composite airfoil 118 during an impact event.

[0127] The transitional weave pattern may have a warp to weft ratio. In a non-limiting example, the warp to weft ratio may vary in a radial direction (Rd) or from a circumferential direction (Cd), but it is contemplated that the warp to weft ratio may vary in any direction.

[0128] Benefits associated with the first transition region 176 and the second transition region 178 described herein include adjusting the composite airfoil assembly 110 ( Figure 5 ). Adjusting the frequency of the composite airfoil assembly 110 may increase part life and reduce fatigue.

[0129] Changing the material in the first transition region 176 or the second transition region 178 smooths the transition in stiffness. A smooth transition in stiffness from one zone or region to another can improve and control the dissipation of energy, for example, during an impact event.

[0130] Changing the braid pattern in the first transition region 176 or the second transition region 178 may also provide a smooth transition in stiffness from the first region 170 to the second region 172 or from the second region 172 to the third region 174 .

[0131] Additional benefits include the ability to enhance control of the zone of weakness by varying stiffness through changes in the material or weave pattern of the composite airfoil 118 so that the composite airfoil 118 bends, deforms, or separates at predetermined desired locations during an ingestion event.

[0132] To the extent not yet described, the different features and structures of the various embodiments may be used in combination as desired, or may be interchanged with one another. Not showing a feature in all embodiments does not mean that it is interpreted as not being able to be shown in this way, but rather is done for the sake of brevity of description. Therefore, the various features of the different embodiments may be mixed and matched as desired to form new embodiments, whether or not the new embodiments are explicitly described. All combinations or permutations of the features described herein are covered by the present disclosure.

[0133] This written description uses examples to disclose the present disclosure, including the best mode, and also to enable any person skilled in the art to practice the present disclosure, including making and using any device or system and performing any combined method. The patentable scope of the present disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. These other examples are intended to fall within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that do not differ substantially from the literal language of the claims.

[0134] Further aspects are provided by the subject matter of the following clauses:

[0135] A composite airfoil assembly for a turbine engine, the composite airfoil assembly comprising a braided core and a skin applied to at least a portion of the exterior of the core, the braided core comprising: a first region of the braided core, the first region comprising a first material having a first stiffness; a second region of the braided core, the second region comprising a second material having a second stiffness, wherein the second stiffness is different from the first stiffness; and a transition region, the transition region connecting the first region and the second region, the transition region having the first material and the second material.

[0136] A composite airfoil assembly as described in any preceding clause, wherein the first stiffness is greater than the second stiffness.

[0137] A composite airfoil assembly as described in any preceding clause, wherein the first material is different from the second material.

[0138] A composite airfoil assembly as described in any preceding clause, wherein the first stiffness of the first material is greater than the second stiffness of the second material.

[0139] A composite airfoil assembly according to any preceding clause, wherein the first region has a first length measured in a radial direction and the transition region has a transition length measured in the radial direction, wherein the transition length is in the range of 1% to 200% of the first length or the second length.

[0140] A composite airfoil assembly as recited in any preceding clause, further comprising a third region having a third material different from the first material and the second material and a third stiffness different from the first stiffness and the second stiffness.

[0141] A composite airfoil assembly as recited in any preceding clause, wherein the transition region is a first transition region and further comprises a second transition region coupling the third region and the second region, wherein the second transition region comprises the first material and the second material.

[0142] A composite airfoil assembly as in any preceding clause, wherein the stiffness of the first and second transition regions varies by more than 2% in the radial direction across the transition length of each of the first and second transition regions.

[0143] A composite airfoil assembly as described in any preceding clause, wherein the braided core comprises a three-dimensional braid pattern.

[0144] A composite airfoil assembly as described in any preceding clause, wherein the first material or the second material comprises carbon fibres.

[0145] A composite airfoil assembly as described in any preceding clause, wherein the first material or the second material comprises metal fibers.

[0146] A composite airfoil assembly as described in any preceding clause, wherein the first material or the second material comprises non-carbon fibers.

[0147] A composite airfoil assembly as described in any preceding clause, wherein the skin is a laminate skin having a skin thickness measured from a skin inner surface to a skin outer surface.

[0148] A composite airfoil assembly as described in any preceding clause, wherein the braided core has a braided core thickness, wherein the braided core thickness is measured from a pressure surface to a suction surface.

[0149] A composite airfoil assembly as described in any preceding clause, wherein the skin thickness is in the range of 1% to 40% of the braided core thickness.

[0150] A composite airfoil assembly as claimed in any preceding clause, wherein the braided core thickness is in the range of 5% to 30% of an airfoil width of the composite airfoil assembly.

[0151] A composite airfoil assembly according to any preceding clause, wherein the first material and the second material are coupled in the transition region, wherein coupling comprises one or more of entanglement, interweaving, twisting, abutting, fastening, weaving, winding, splicing, braiding or crossing of the first material and the second material.

[0152] A composite airfoil assembly according to any preceding clause, wherein the transition region includes an intersection defined by the first material being coupled to the second material, wherein the coupling of the first material and the second material includes one or more of entanglement, interweaving, twisting, abutting, fastening, weaving, winding, splicing, braiding, or crossing of the first material and the second material.

[0153] A composite airfoil assembly as described in any preceding clause, wherein the first material and the second material are interwoven within the transition region.

[0154] A composite airfoil assembly as described in any preceding clause, wherein the first material and the second material comprise the same material.

[0155] A composite airfoil assembly as recited in any preceding clause, wherein the first material has a first weave pattern in the first region and the second material has a second weave pattern in the second region, wherein the first weave pattern is different from the second weave pattern.

[0156] A composite airfoil assembly as recited in any preceding clause, wherein the transition region has a transition weave pattern, wherein the transition weave pattern includes at least some portions of the first weave pattern and the second weave pattern.

[0157] A composite airfoil assembly according to any preceding clause, wherein the transitional weave pattern has an intersection, wherein the intersection includes a connection location between the first weave pattern and the second weave pattern, wherein the connection includes one or more of entanglement, interweaving, twisting, abutting, fastening, weaving, wrapping, splicing, braiding or crossing of the first weave pattern and the second weave pattern.

Claims

1. A composite airfoil assembly for a turbine engine, characterized in that: The composite airfoil assembly comprises: A braided core defining a core exterior, the braided core comprising: a first region comprising a first material having a first stiffness; a second region comprising a second material having a second stiffness, wherein the second stiffness is different than the first stiffness; and a transition region connecting the first region and the second region, the transition region comprising the first material and the second material; and A skin is applied to at least a portion of the exterior of the core.

2. The composite airfoil assembly according to claim 1, wherein: in, The first stiffness is greater than the second stiffness.

3. The composite airfoil assembly according to claim 2, wherein: in, The first stiffness of the first region is greater than the second stiffness of the second region.

4. The composite airfoil assembly of claim 1, wherein: in, The first region has a first length measured in a radial direction, the second region has a second length measured in the radial direction, and the transition region has a transition length measured in the radial direction, wherein the transition length is in the range of 1% to 200% of the first length or the second length.

5. The composite airfoil assembly of claim 1, wherein: Further included is a third region having a third material different from the first material and the second material and a third stiffness different from the first stiffness and the second stiffness.

6. The composite airfoil assembly according to claim 5, wherein: in, The transition region is a first transition region, and the composite airfoil assembly further includes a second transition region coupling the second region and the third region, wherein the second transition region includes the second material and the third material.

7. The composite airfoil assembly according to claim 6, wherein: in, The stiffness of each of the first transition region and the second transition region varies by more than 2% in a radial direction across a transition length of each of the first transition region and the second transition region.

8. The composite airfoil assembly of claim 1, wherein: in, The braided core includes a three-dimensional braid pattern.

9. The composite airfoil assembly according to any one of claims 1 to 8, characterized in that: in, The first material or the second material includes carbon fibers.

10. The composite airfoil assembly according to any one of claims 1 to 8, characterized in that: in, The first material or the second material includes metal fibers.

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