Gas turbine engine with composite airfoil and preform core

Abstract

A gas turbine engine includes a fan section, a compressor section, a combustion section, and a turbine section arranged in series flow and defining an engine centerline extending between a forward direction and an aft direction. A composite component, such as an airfoil, extends between a root and a tip defining a spanwise direction therebetween, and extends between a leading edge and a trailing edge forming a chordwise direction therebetween. The composite component includes a preform core including a first set of fibers as a first set of warp fibers and a first set of weft fibers.

Description

Technical Field

[0001] The present disclosure relates generally to a core or preform thereof for a component of a gas turbine, and more particularly to a preform of an airfoil having a braided core geometry. Background Art

[0002] A turbine engine typically includes an engine core having a compressor section, a combustor section, and a turbine section arranged in a series flow. In a bypass turbine engine, a fan section may be provided upstream of the compressor section. The compressor section compresses air, which is directed to the combustor section where it is mixed with fuel and the mixture is then ignited to produce hot combustion gases. The combustion gases are directed to the turbine section, which extracts energy from the combustion gases to drive the compressor section and to produce useful work to propel an aircraft in flight or to power a load such as a generator.

[0003] Many parts of turbine engines are made of composite materials, which may include a fiber-reinforced matrix and have a high strength-to-weight ratio. Due to the high strength-to-weight ratio and moldability into relatively complex shapes, composite materials are used in various applications, such as turbine engines or aircraft. Parts made of composite materials can, for example, be mounted on or define a fuselage and / or a wing, rudder, manifold, airfoil, blade, vane or other part of an aircraft or turbine engine. 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 according to an exemplary aspect of the present disclosure.

[0006] Figure 2 is suitable for use according to an exemplary aspect of the present invention Figure 1 A schematic perspective view of a composite airfoil assembly and disk assembly within a turbine engine.

[0007] Figure 3 According to an exemplary aspect of the present invention Figure 2 Schematic cross-sectional view of a composite airfoil assembly including a core with a group of fibers.

[0008] Figure 4 According to an exemplary aspect of the present invention Figure 2Another schematic cross-sectional view of a composite airfoil assembly in FIG. 1 , the composite airfoil assembly comprising a composite airfoil assembly arranged at an angle Figure 3 A group of fibers.

[0009] Figure 5 is a flow chart depicting a method of forming a component according to an exemplary aspect of the invention. DETAILED DESCRIPTION

[0010] Aspects disclosed herein relate to a manufacturing core for an engine component, such as an airfoil. The core may be woven and may include additional woven layers that form the engine component. The core is used to create an engine component of a turbine engine or a preform thereof. For example, such an engine component may be an airfoil. However, it should be understood that the present disclosure is applicable to other engine components of a turbine engine, such as a combustor liner, a shroud, a nozzle, or a disk in a non-limiting example. In addition, although described in the form of a core for manufacturing an airfoil or an aircraft, it should be understood that the present disclosure is applicable to any other suitable environment, such as terrestrial or non-terrestrial, aerial, land-based or non-land-based applications.

[0011] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are shown in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. The same or similar designations in the drawings and description are used to refer to the same or similar parts of the present disclosure.

[0012] As used herein, the word "exemplary" means "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or better than other implementations. In addition, unless explicitly stated otherwise, all embodiments described herein should be considered exemplary.

[0013] As used herein, the terms “first,” “second,” “third,” and “fourth” may be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0014] As used herein, the term "upstream" refers to the direction opposite to the direction of fluid flow, and 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 "backward" mean behind something. For example, when used in relation to fluid flow, front / forward may refer to upstream, while rear / backward may refer to downstream.

[0015] The term "fluid" may be gas or liquid, or multiple phases.

[0016] As used herein, the term "camber line" refers to a measure of the curvature of an airfoil. The camber line may be an imaginary line located midway between the pressure side and the suction side, extending between and intersecting the leading edge and the trailing edge.

[0017] As used herein, the term "radial" or "radially" refers to directions away from a common center. For example, in the general context of turbine engines, radial refers to directions along a ray extending between a central longitudinal axis of the engine and an outer periphery of the engine.

[0018] All directional references that may be used herein (e.g., radial, axial, coaxial, front, rear, top, upstream, downstream, front, rear, etc.) are used only for identification purposes to help the reader understand the present disclosure and do not create limitations, particularly with respect to the position, orientation, or use of the aspects disclosed herein. Connection references (e.g., coupling, connection) should be interpreted broadly and, unless otherwise indicated, may include intermediate structural elements between sets of elements, as well as relative movement between elements. Therefore, a connection reference does not necessarily mean that the two elements are directly connected and fixed to each other. The exemplary drawings are for illustrative purposes only, and the dimensions, positions, orders, and relative sizes reflected in the drawings may vary.

[0019] The singular forms "a," "an," and "the" include plural referents unless the context clearly dictates otherwise. Also, as used herein, the term "set" or "a group of" elements may be any number of elements, including only one.

[0020] As used herein, the term "stiffness" may be used to define a mechanical property that defines the degree to which a structure resists deformation under the action of a force. Stiffness may be defined as the ratio of a force to the displacement of an object under the action of said force. Stiffness may include resistance to deformation under the action of forces applied from various directions, where, in non-limiting examples, stiffness may represent axial stiffness, tensile stiffness, compressive stiffness, torsional stiffness, or shear stiffness.

[0021] As used herein, the term "elasticity" may be used to define a mechanical property that defines the modulus of elasticity under tension or compression, and may relate to the elasticity of a particular material or structure made from such a material, such as an engine component as described herein. Elasticity may represent stress per unit area relative to its local strain or proportional deformation.

[0022] As used herein, the term "strength" may be used to define a mechanical property that defines the maximum load that a particular element or feature can withstand without breaking, cracking, or otherwise deforming due to the load experienced. In non-limiting examples, "strength" may include tensile strength, shear strength, torsional strength, or compressive strength.

[0023] As used herein, the term "hardness" may be used to define a mechanical property that defines resistance to deformation due to indentation or wear. In non-limiting examples, hardness may include scratch hardness, indentation hardness, or rebound hardness. Hardness may be measured in Newtons per square millimeter, or may be expressed in Brinell hardness numbers, Vickers hardness numbers, or Rockwell hardness numbers in non-limiting examples.

[0024] The term "composite material" as used herein refers to a component having two or more materials. A composite material can be a combination of at least two or more metals, non-metals, or 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, thermoplastic resins, bismaleimide (BMI) materials, polyimide materials, epoxy resins, glass fibers, and silicon matrix materials.

[0025] 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 multiple layers or multiple layups of composite materials. The mechanical properties, materials, and dimensions of these layers or layups may be varied to achieve a desired composite component or composite portion of a component having a predetermined weight, size, or mechanical properties.

[0026] One or more layers of adhesive may be used to form or join the composite parts. The adhesive may include resins and phenolics, where the adhesive may need to be cured at elevated temperatures or employ other hardening techniques.

[0027] As used herein, PMC refers to a class of materials. For example, PMC materials are partially defined by prepregs, which are reinforcing materials pre-impregnated with a polymer matrix material (such as a thermoplastic resin). Non-limiting examples of processes for producing thermoplastic prepregs include hot melt prepregs, in which fiber reinforcements are pulled through a molten resin bath, and powder prepregs, in which resins are electrostatically deposited (as non-limiting examples) on fiber reinforcements and then adhered to the fibers in an oven or with the aid of heated rollers (as non-limiting examples). Prepregs can be in the form of unidirectional tapes or woven fabrics, which are then stacked together to form the number of stacked layers required for the component.

[0028] The multilayer prepreg is stacked to the appropriate thickness and orientation of the composite part, and then cured and solidified resin is formed into a fiber reinforced composite part. The resin for the matrix material of PMC can be generally divided into thermosetting resin or thermoplastic resin. Thermoplastic resin is generally classified as a polymer that can be repeatedly softened and flowed when heated due to physical changes rather than chemical changes and hardened when fully cooled. The significant example category of thermoplastic resin includes nylon, thermoplastic polyester, polyaryletherketone and polycarbonate resin. The specific example of high-performance thermoplastic resin considered for aerospace applications includes 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 significantly soften when heated, but thermally decompose when fully heated. The significant example of thermosetting resin includes epoxy resin, bismaleimide (BMI) and polyimide resin.

[0029] In another non-limiting example, in addition to using prepregs, in the case of using thermoplastic polymers, woven fabrics can be utilized. Woven fabrics can include, but are not limited to, dry carbon fibers woven together with thermoplastic polymer fibers or filaments. Non-prepreg woven structures can be made in a similar manner. In this way, the fiber volume of the component can be adjusted by specifying the relative concentrations of the thermoplastic fibers and reinforcing fibers that have been woven or braided together. Additionally, different types of reinforcing fibers can be woven or braided together to adjust the properties of the component at various concentrations. For example, glass fibers, carbon fibers, and thermoplastic fibers can all be woven together at various concentrations to adjust the properties of the component. Carbon fibers provide the strength of the system, glass fibers can be added to enhance impact performance, which is a design feature of components located near the engine inlet, and thermoplastic fibers provide bonding for the reinforcing fibers.

[0030] In another non-limiting example, resin transfer molding (RTM) can be used to form at least a portion of the composite component. Generally speaking, RTM includes applying dry fiber or matrix material to a mold or cavity. The dry fiber or matrix material can include prepreg, woven material, braided material, or any combination thereof.

[0031] The resin may be pumped or otherwise provided to the mold or cavity to impregnate the dry fibers or matrix material. The impregnated fiber or matrix material and resin combination is then cured and removed from the mold. Upon removal from the mold, the composite part may require a post-curing process.

[0032] 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.

[0033] The dry fiber or matrix material may be shaped to form a composite part or to introduce resin. Optionally, additional layers or reinforcement layers of a material different from the dry fiber or matrix material may also be included or added prior to heating or curing.

[0034] As used herein, CMC refers to a class of materials containing reinforcing fibers in a ceramic matrix. Generally speaking, the reinforcing fibers provide structural integrity to the ceramic matrix. Some examples of reinforcing fibers may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), non-oxide carbon-based materials (e.g., carbon), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates (such as mullite), or mixtures thereof), or mixtures thereof.

[0035] Some examples of ceramic matrix materials may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates, or mixtures thereof), or mixtures thereof. Optionally, the ceramic matrix may also include ceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite).

[0036] In general, a particular CMC can be referred to as its fiber type / matrix type combination. For example, C / SiC represents carbon fiber reinforced silicon carbide; SiC / SiC represents silicon carbide fiber reinforced silicon carbide, SiC / SiN represents silicon carbide fiber reinforced silicon nitride; SiC / SiC-SiN represents silicon carbide fiber reinforced silicon carbide / silicon nitride matrix mixture, etc. In other examples, the CMC can be composed of a matrix and reinforcing fibers, the reinforcing fibers including oxide-based materials such as aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates, and mixtures thereof. Aluminosilicates can include crystalline materials such as mullite (3Al2O3·2SiO2), as well as glassy aluminosilicates.

[0037] In certain 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 into reinforcement tapes, such as unidirectional reinforcement tapes. Multiple tapes may be laid together to form a preformed component. The fiber bundles may be impregnated with a slurry composition before or after forming the preform. The preform may then be heat treated and subsequently chemically treated to obtain a component formed of a CMC material having a desired chemical composition. For example, the preform may be cured or burned out to produce a high carbon residue in the preform, followed by melt infiltration with silicon, or cured or pyrolyzed to produce a silicon carbide matrix in the preform, followed by chemical vapor infiltration with silicon carbide. Before or after chemical vapor infiltration, other steps may be taken to improve the densification of the preform by injecting a liquid resin or polymer into the preform and then performing a heat treatment step to fill the voids with silicon carbide. The CMC materials 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.

[0038] Such materials, as well as certain monolithic ceramics (i.e., ceramic materials that do not contain reinforcement materials), are particularly suitable for high temperature applications. Additionally, these ceramic materials are lightweight compared to superalloys, but can still provide strength and durability to parts made therefrom. As a result, such materials are currently being considered for use in many gas turbine components used in the high temperature sections of gas turbine engines, such as airfoils (e.g., turbines and buckets), combustors, shrouds, and other similar components that would benefit from the lighter weight and higher temperature capabilities that these materials can provide.

[0039] The term "metal" as used herein refers to materials including metals, such as but not limited to titanium, iron, aluminum, stainless steel, 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.

[0040] The inventors practiced as described above by designing a core for manufacturing a component such as an airfoil, designing an airfoil with improved local mechanical properties or geometry, reduced weight, determining whether the component was manufactured as designed and met the component's purpose, and modifying the engine component with new or different characteristics in an iterative process when the engine component did not meet the component's purpose. Figure 1 This process is repeated when

[0041] Figure 11 is a schematic cross-sectional view of a turbine engine 10 for an aircraft. The turbine engine 10 has a generally longitudinally extending axis or engine centerline 12 extending from a forward direction 14 to an aft direction 16. The turbine engine 10 includes, in downstream serial flow relationship, 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.

[0042] The fan section 18 includes a fan case 40 surrounding the fan 20. The fan 20 includes a plurality of fan blades 42 radially arranged around the engine centerline 12. The HP compressor 26, the combustor 30, and the HP turbine 34 form an engine core 44 of the turbine engine 10, which produces combustion gases. The engine core 44 is surrounded by a core case 46, which may be coupled to the fan case 40.

[0043] 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 HP spool 48 and the LP spool 50 are rotatable about the engine centerline 12 and are coupled to a plurality of rotatable elements that may collectively define a rotor 51.

[0044] The LP compressor 24 and the HP compressor 26 each include a plurality 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 to compress or pressurize a fluid flow through the stage. In a single compressor stage 52, 54, the plurality of compressor blades 56, 58 may be arranged in a ring and may extend radially outward from a blade platform to a blade tip relative to the engine centerline 12, with the corresponding static compressor vanes 60, 62 positioned upstream of and adjacent to the rotating 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.

[0045] The compressor blades 56, 58 of the compressor stage may be mounted to (or integrated into) a disk 61 mounted to a respective one of the HP and LP spools 48, 50. The static compressor vanes 60, 62 of the compressor stage may be mounted to the core housing 46 in a circumferential arrangement.

[0046] The HP turbine 34 and the LP turbine 36 each include a plurality of turbine stages 64, 66, wherein 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 a fluid flow passing through the stage. In a single turbine stage 64, 66, the plurality of turbine blades 68, 70 may be arranged in a ring and may extend radially outward relative to the engine centerline 12, while the corresponding static turbine vanes 72, 74 are positioned upstream and adjacent to the rotating turbine blades 68, 70. Notably, Figure 1 The number of blades, buckets, and turbine stages shown in FIG. 5 are selected for illustration purposes only, and other numbers are possible.

[0047] The turbine blades 68, 70 of the turbine stage may be mounted to a disk 71 which in turn is mounted to a respective one of the HP and LP spools 48, 50. The static turbine buckets 72, 74 of the compressor stage may be mounted to the core housing 46 in a circumferential arrangement.

[0048] Complementing the rotor portion, the stationary portion of the turbine engine 10, such as the stationary blades 60, 62, 72, 74 in the compressor and turbine sections 22, 32, are also individually or collectively referred to as stators 63. Thus, the stator 63 may refer to the combination of non-rotating components in the entire turbine engine 10.

[0049] In operation, the airflow exiting the fan section 18 is split so that a portion of the airflow is directed into the LP compressor 24, which then supplies a pressurized airflow 76 to the HP compressor 26, which further pressurizes the air. The pressurized airflow 76 from the HP compressor 26 is mixed with fuel in the combustor 30 and ignited, thereby producing combustion gases. The HP turbine 34 extracts some work from these gases, thereby driving the HP compressor 26. The combustion gases are exhausted into the LP turbine 36, which extracts additional work to drive the LP compressor 24, and the exhaust gases are exhausted 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.

[0050] A portion of the pressurized air flow 76 may be extracted from the compressor section 22 as bleed air 77. The bleed air 77 may be extracted from the pressurized air flow 76 and provided to engine components that require cooling. The temperature of the pressurized air flow 76 entering the combustor 30 is significantly higher than the bleed air temperature. The bleed air 77 may be used to reduce the temperature of core components downstream of the combustor 30.

[0051] The remainder of the airflow exiting the fan section 18, referred to as bypass airflow 78, bypasses the LP compressor 24 and the engine core 44 and exits the turbine engine 10 through the stationary blade rows, and more specifically, exits the turbine engine 10 through an exit guide vane assembly 80 that includes a plurality of airfoil guide vanes 82 on the fan exhaust side 84. More specifically, a circumferential row of radially extending airfoil guide vanes 82 are used about the fan section 18 to exert some directional control on the airflow 78.

[0052] A portion of the air supplied by the fan 20 may bypass the engine core 44 and be used to cool portions of the turbine engine 10, particularly the hot portions, and / or to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are generally located downstream of the combustor 30, particularly the turbine section 32, with the HP turbine 34 being the hottest portion as it is located directly downstream of the combustion section 28. Other sources of cooling fluid may be, but are not limited to, fluid exhausted from the LP compressor 24 or the HP compressor 26.

[0053] Figure 2 Is suitable for Figure 1 1 is a schematic perspective view of a composite airfoil 100 and a disk 102 for use within a turbine engine 10 of FIG. Composite airfoil 100 may define a body 104. Disk 102 is suitable for use as disks 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 100 may be rotating or non-rotating, such that the composite airfoil 100 may include static compressor blades 60, 62 ( Figure 1 ), a set of compressor blades 56, 58 ( Figure 1 ), static turbine blades 72, 74 ( Figure 1 ), a set of turbine blades 68, 70 ( Figure 1 ) or a plurality of fan blades 42 ( Figure 1 ). As a non-limiting example, the composite airfoil 100 may be a composite fan blade.

[0054] The disk 102 may rotate or be stationary about an axis of rotation 106. The axis of rotation 106 may be aligned with the engine centerline (e.g., Figure 1 The disk 102 includes a plurality of slots 108 extending axially through a radially outer portion of the disk 102 and spaced circumferentially around the disk 102 relative to the axis of rotation 106.

[0055] Composite airfoil 100 and its body 104 extend between leading edge 114 and trailing edge 116 to define a chordwise direction and between root 118 and tip 120 to define a spanwise direction. Composite airfoil 100 includes pressure side 122 and suction side 124 and may include interior 126 defined within composite airfoil 100.

[0056] The composite airfoil 100 may be coupled to the disk 102 by inserting at least a portion of the composite airfoil 100 into a corresponding slot in the plurality of slots 108. The composite airfoil 100 may be secured in place by frictional contact with the slot 108, or may be coupled to the slot 108 by any suitable coupling method, such as, but not limited to, welding, bonding, fastening, etc. Although only a single composite airfoil 100 is shown, it should be understood that there may be any number of one or more composite airfoils 100 coupled to the disk 102. As a non-limiting example, there may be a number of composite airfoils 100 that corresponds to the total number of slots in the plurality of slots 108.

[0057] For ease of reference, a set of relative reference directions and coordinate systems may be applied to the composite airfoil 100. The axial direction (Ad) may extend from front to back and is shown extending at least partially into the page. The axial direction (Ad) may be arranged parallel to the rotation axis 106, which may also be arranged parallel to the engine centerline 12 ( Figure 1 ). The radial direction (Rd) extends perpendicular to the axial direction (Ad) and may extend perpendicular to the engine centerline 12. The circumferential direction (Cd) can be defined as being perpendicular to the radial direction (Rd), the circumferential direction (Cd) can be defined along a circumferential curvature around the axial direction (Ad), can be defined as a ray extending locally from the radial direction (Rd), and / or can be defined along the circumference of the turbine engine 10 relative to the engine centerline 12 or the rotation axis 106.

[0058] Figure 3 A schematic cross-sectional view of a composite core 110 of a composite airfoil 100 is shown, and illustrating the interior 126. The composite core 110 includes a preform core 150 and a skin 152, with a coating 154 positioned exterior to the skin 152. In one example, the composite core 110 may be formed from the preform core 150 without the skin 152.

[0059] The preform core 150 may be made at least in part of a braided structure. In a non-limiting example, the braided structure may include a first group of fibers 160, which may include a first group of warp fibers 162 and a first group of weft fibers 164. The first group of warp fibers 162 and the first group of weft fibers 164 may be interlaced or interwoven to at least partially form the preform core 150. In a non-limiting example, a jacquard loom or a 3D braiding machine may be used to create a complex braided structure, which may include interweaving one or more materials or composite materials to form the preform core 150. In a non-limiting example, such materials or composite materials may include carbon or carbon fibers, glass or fiberglass, nylon, rayon or other aramid fibers, however other materials such as nickel, titanium or ceramic composites are also contemplated.

[0060] In one non-limiting example, the first weaving pattern may be a three-dimensional weaving pattern. More specifically, the weaving structure may be arranged along the axial direction (Ad), the radial direction (Rd), and the circumferential direction (Cd) ( Figure 2 ) combination. The axial direction (Ad) can be aligned with the engine centerline 12 ( Figure 1 ) are collinear, and the radial direction (Rd) may be a line defined by diverging perpendicular to the engine centerline 12. It should be understood that the braid pattern may be independent of the turbine engine 10 ( Figure 1 ) and their directions (Ad, Rd, Cd) are 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 three mutually orthogonal planes may include a first plane defined between the spanwise direction and the chordwise direction, and two additional mutually orthogonal planes, both of which are orthogonal to the first plane and to each other.

[0061] The skin layer 152 is positioned outside the preform core 150 and can be disposed on or applied to the preform core 150. In one non-limiting example, the skin layer 152 can be formed as a plurality of laminated plies 170 on the preform core 150. In another non-limiting example, the skin layer 152 can be a woven layer, similar to the woven layer of the preform core 150. In such an example, the weaving patterns for the preform core 150 and the skin layer 152 can be similar, the same, or different, or have discrete areas that can be similar, the same, or different from other parts of the preform core 150 or the skin layer 152. Such a woven layer includes a second group of fibers 172, which includes a second group of warp fibers 174 and a second group of weft fibers 176. In such an example, the skin layer 152 can be formed of a composite material using a jacquard loom or a 3D weaving machine, which can include a weaving pattern or structure similar or different from the preform core 150. In additional non-limiting examples, the skin 152 may be composed of a composite material similar to or different from the preform core 150, such as carbon or carbon fibers, glass or fiberglass, nylon, rayon, or aramid fibers, while in non-limiting examples, the skin 152 may be composed of other materials such as nickel, titanium, or ceramic composites. The size and shape of the skin 152 may form the outer airfoil shape, such as defining a greater or lesser local thickness.

[0062] In yet another example, the preform core 150, the skin 152, or both may be woven into a braided braided pattern or a braided braided pattern. The braided or braided braided pattern may be formed into a braided structure and may include a braided pattern comprising three or more interwoven fibers, tows, yarns, or strands woven in a repeating pattern, such as a first group of fibers 160 or a second group of fibers 172. In another non-limiting example, the braided geometry may include a first group of fibers 160 or a second group of fibers 172 laid in sequence to define the braided geometry. In one example, the braided geometry or pattern may be repeated over the entire skin 152, or only over a portion thereof, wherein one or more additional braids or braided geometries define the skin 152. In yet another example, the braided pattern of the skin 152 may be a braided geometry different from the preform core 150, although both are braided. In a non-limiting example, this difference may include differences in materials, braided patterns, arrangements, arrangements, or offsets from one another.

[0063] The first group of fibers 160 and the second group of fibers 172 forming the preform core 150 and the skin 152 may be arranged in a first direction 180 and a second direction 182. For example, in one non-limiting example, the first group of warp fibers 162 and / or the second group of warp fibers 174 may be arranged in a first direction 180, wherein the first direction 180 may be aligned with a radial direction (Rd) and the second direction 182 may be aligned with a span direction. In another example, the first direction 180 may be aligned with the engine centerline 12 ( Figure 1 ). In another non-limiting example, the first direction 180 may be defined as extending between a midpoint of a camber line of the composite airfoil 100 at the root 118 and a midpoint of a camber line at the tip 120. In another non-limiting example, the first direction 180 may be defined as projected onto a common plane with the first group of fibers 160 and aligned with at least one of the first group of warp fibers 162 or the first group of weft fibers 164.

[0064] The first group of weft fibers 164 and / or the second group of weft fibers 176 can be arranged in a second direction 182 that is perpendicular to the first direction 180. In one non-limiting example, the second direction 182 can be aligned with the axial direction (Ad). In additional non-limiting examples, it is contemplated that the second direction 182 is aligned with the radial direction (Rd), the circumferential direction (Cd), or the chord direction.

[0065] The coating 154 may form an outer wall. It is contemplated that one or more additional inner or outer layers may be provided inside or outside the coating 154, between the coating 154 and the skin 152, inside the skin 152, or between the skin 152 and the preform core 150. The coating 154 may be a barrier coating, for example, to prevent erosion due to impact with an object, while additional properties may be contemplated, such as hydrophobic or ice-repellent properties, corrosion-resistant or erosion-resistant properties, or UV-resistant properties. Additional finishing layers or materials may be provided on the coating 154 as desired, such as an anti-oxidation or corrosion-resistant coating or paint.

[0066] During manufacturing, the preform core 150 and / or the skin 152 can be formed to define a specific braided structure. The first group of fibers 160 and the second group of fibers 172 are arranged in two mutually orthogonal directions, illustrated as a first direction 180 and a second direction 182, the first direction 180 being aligned with the radial direction (Rd) or the span direction, and the second direction 182 being aligned with the axial direction (Ad) or the chord direction. The specific braided structure can be specified as having a predetermined geometry and arrangement of the first group of fibers 160 and the second group of fibers 172, and / or can be cut or otherwise sized and shaped after the braided structure is manufactured, such as by cutting or grinding the preform core 150 into a desired shape. Such cutting or sizing can be used to create or define the shape of the composite core 110 or a preform, which can include the preform core 150, the skin 152, or both.

[0067] Figure 4 Another schematic cross-sectional view of the composite core 110 of the composite airfoil 100 is shown. The first group of fibers 160 further includes a group of chord-wise fibers 190 that are aligned with the first direction 180 and the second direction 182 and have other fibers 192 arranged at a first angle 184 and a second angle 186 relative to the first direction 180. The fibers 192 arranged at the first angle 184 and the second angle 186 may be offset from the first direction 180 or the second direction 182, or both. In a non-limiting example, the first angle 184 may vary between thirty-five degrees and sixty-five degrees (35° to 65°), and the second angle 186 may vary between negative thirty-five degrees and negative sixty-five degrees (-35° to -65°). It should be understood that negative angles represent extension in a local forward direction and positive angles represent extension in a local rearward direction, wherein the radial direction (Rd) represents a zero degree value. Alternatively, a negative angle may represent an orientation in a direction opposite to the orientation of a positive angle relative to a common axis, as can be understood by one of ordinary skill in the art.

[0068] In one non-limiting example, at least 20% of the fiber volume of the first group of fibers 160 forming the preform core 150 may be arranged at the first angle 184, and at least 20% of the fiber volume of the preform core 150 may be arranged at the second angle 186. In such an example, at least 40% of the fiber volume of the first group of fibers 160 may be arranged at the first angle 184 and the second angle 186.

[0069] In another non-limiting example, at least some of the second group of fibers 172 forming the surface layer 152 can be arranged at a third angle 194 and a fourth angle 196, such that the third angle 194 of the second group of warp fibers 174 can vary between thirty-five degrees and sixty-five degrees (35° to 65°), and the fourth angle 196 of the second group of weft fibers 176 can vary between negative thirty-five degrees and negative sixty-five degrees (-35° to -65°).

[0070] In another non-limiting example, at least 20% of the fiber volume of the skin layer 152 may be arranged at the third angle 194, and at least 20% of the fiber volume of the skin layer 152 may be arranged at the fourth angle 196. In yet another non-limiting example, at least 40% of the total volume or mass of the skin layer 152 may be arranged at the third angle 194 or the fourth angle 196.

[0071] In one non-limiting example, additional layers, sheets, volumes, or regions may include additional groups of fibers similar to the first group of fibers 160 or the second group of fibers 172. For example, the first region 200 has some fibers of the first group of fibers 160 arranged at the first angle 184 and the second angle 186 that are added to, arranged with, passed through, located in, or covered on other fibers of the first group of fibers 160 that are aligned with the spanwise and chordwise directions but are not arranged at the first angle 184 and the second angle 186. The additional fibers added at the first angle 184 and the second angle 186 may define the first region 200 and may provide a localized increase in thickness, which may be used to define the three-dimensional shape of the preform core 150. In another non-limiting example, the second region 202 may include additional fibers arranged at the first angle 184 and the second angle 186. Such additional fibers within the second region 202 may be arranged at the first angle 184 and the second angle 186 similar to the first region 200, although different angles between different regions are contemplated. The second region 202 includes at least some of the first group of fibers 160 that are not arranged at the first angle 184 and the second angle 186 in addition to the first group of fibers 160 arranged at the first angle 184 and the second angle 186. These local additions can determine the local geometry of the composite airfoil 100 or the composite core 110. Additionally, the additions that define the first region 200 or the second region 202 can be used to define the mechanical properties of the composite airfoil 100 or the composite core 110, including, but not limited to, increased, decreased, or changed local thickness, stiffness, elasticity, hardness, strength, center of rotation, or center of mass, in non-limiting examples. The determination of these mechanical properties can be achieved by using different materials, geometries, weave patterns, or positions, orientations, or amounts of the first group of fibers 160, the first group of warp fibers 162, or the first group of weft fibers arranged at the first angle 184 and the second angle 186.

[0072] Similarly, additional groups of fibers, such as the second group of fibers 172 arranged at the third angle 194 and the fourth angle 196, may be locally added to the third region 204 or the fourth region 206, wherein other fibers of the second group of fibers 172 that are not arranged at the third angle 194 and the fourth angle 196 are disposed in the third region 204 or the fourth region 206, pass through, are located in, or cover the other fibers. These additional fibers may also be used to determine the local geometry of the skin 152 of the composite airfoil 100 or the composite core 110, such as defining mechanical properties, including, but not limited to, increased, decreased, or altered local thickness, stiffness, elasticity, strength, center of rotation, or center of mass, in non-limiting examples.

[0073] In one non-limiting example, the composite core 110 can be formed as a preform. As used herein, a "preform" is a three-dimensional composite material formed by a plurality of fibers or yarns, including warp fibers or yarns and weft fibers or yarns. In addition, at least some of the fibers or yarns are arranged at an offset angle with the warp and weft fibers or yarns, and the offset angle can vary between 35° to 65° or -35° to -65°. The fibers or yarns, or at least some of them, can be integrated into the preform by weaving, stitching, tufting or any suitable manufacturing process. The preform referred to herein is a fully formed component that is formed in a separate manufacturing process relative to the fully formed component.

[0074] The composite core 110 as a preform may include a preform core 150 and may or may not include a skin layer 152 or other layers (not shown). The preform may be formed with a first set of fibers 160 and a second set of fibers 172 at a first angle 184, a second angle 186, a third angle 194, and / or a fourth angle 196 and may then be otherwise sized, cut, finished, or processed to form the composite core 110. Using the composite core 110 as a preform may provide a core for an airfoil or other engine component that may be finished to form a composite airfoil or engine component. In non-limiting examples, the preform may include a resin transfer molding, a precured resin transfer molding, a prepreg laminate, or an inter-molding and prepreg formed from the same qualified resin transfer molding.

[0075] In a non-limiting example, utilizing one or more groups of fibers, such as the first group of fibers 160 and the second group of fibers 172 at the first angle 184, the second angle 186, the third angle 194, and / or the fourth angle 196, a variable local thickness, material composition, geometry, orientation, physical properties, or mechanical properties, such as stiffness, elasticity, hardness, strength, center of rotation, or center of mass, may be achieved. Adding the first group of fibers 160 and the second group of fibers 172 at the first angle 184, the second angle 186, the third angle 194, and / or the fourth angle 196 may combine additional materials to define the local thickness, geometry, or orientation of the fibers, thereby changing one or more mechanical properties. Furthermore, utilizing different weave patterns, angles of fibers, the number of angled fibers, or different materials may change the mechanical properties of the material forming the composite airfoil 100, thereby changing the stiffness, hardness, elasticity, or strength of the composite core 110. In addition to arranging at least 20% of the fiber volume at the first angle 184, the second angle 186, the third angle 194, and / or the fourth angle 196, local physical or mechanical properties may also be altered by changing the materials used in the fibers. In another example, the first group of fibers 160 and the second group of fibers 172 at the first angle 184, the second angle 186, the third angle 194, and / or the fourth angle 196 may change the natural frequency of the composite airfoil 100 or its composite core 110. Adjusting the vibration frequency of the composite airfoil 100 may reduce vibration or change the vibration frequency to a frequency outside of the operating conditions of the composite airfoil 100. Such adjustment of the natural frequency may reduce susceptibility to vibration wear or fatigue, as well as increase component life and reduce overall maintenance. Furthermore, changing or moving the center of rotation or center of mass of the composite airfoil 100 by adding the first and second groups of fibers 160 , 172 at the first, second, third, and / or fourth angles 184 , 186 , 194 , and / or 196 may provide enhanced or improved balance for the composite airfoil 100 , which may increase operating efficiency and component life.

[0076] Figure 5 A flow chart depicting a method 220 of forming a component, such as the composite airfoil 100 or composite core 110 described herein, is shown. The method 220 includes, at 222, forming a preform core of the component, such as the preform core 150 ( Figure 3-4 The preform core 150 may be at least partially formed of fibers, such as fibers in a first direction 180 and a second direction 182 ( Figure 3 ) extends on the first group of fibers 160 ( Figure 3-4 In one non-limiting example, the fibers may be woven to define a woven preform core. Forming the preform core 150 may further include, for example, weaving the fibers in a first direction 180 ( Figure 3 ) or the second direction 182( Figure 3) is applied to the first set of fibers. In a non-limiting example, a machine such as a jacquard loom or a 3D braiding machine can be used to create the preform core 150.

[0077] At 224, the method 220 may include rotating the preform core 150. The rotation of the preform core 150 may be accomplished by rotating the machine that forms the preform core 150, or by the machine or a person operating the machine rotating a single component (i.e., the preform core 150), or by other rotations to orient the preform core 150 at a certain angular position. This rotation of the preform core 150 defines the angle of the fibers applied to the preform core 150, such as the first group of fibers 160 ( Figure 4 ) at a first angle 184( Figure 4 ) or the second angle 186( Figure 4 ) orientation.

[0078] At 226, the method 220 may include applying fibers to the preform core 150 at an angle. After rotating the preform core 150 at 224, the first group of fibers 160 (including the first group of warp fibers 162 and the first group of weft fibers 164) are applied at a first angle 184 and / or a second angle 186, and the first group of fibers 160 are arranged at the first angle 184 and / or the second angle 186. The first angle 184 may vary between thirty-five degrees and sixty-five degrees (35° to 65°), and the second angle 186 may vary between negative thirty-five degrees and negative sixty-five degrees (-35° to -65°). In a non-limiting example, at least 20% of the fiber volume of the preform core 150 may be arranged at the first angle 184, and at least 20% of the fiber volume of the preform core 150 may be arranged at the second angle 186. In addition, it is contemplated that at least a portion of the preform core 150 is formed using a braided geometry.

[0079] At 228, method 220 may include applying a skin layer, such as skin layer 152 ( Figure 3-4 In one non-limiting example, the skin layer 152 can be formed as a set of plies or a plurality of plies. In another non-limiting example, the skin layer 152 can be applied as a set of fibers, such as the second set of fibers 172 ( Figure 3-4), including a second group of warp fibers 174 and a second group of weft fibers 176. At least some of the second group of fibers 172 can be aligned with the first direction 180 and the second direction 182, and at least some of the second group of fibers 172 can be arranged at a third angle 194 and a fourth angle 196. The third angle 194 of the second group of warp fibers 174 can vary between thirty-five degrees and sixty-five degrees (35° to 65°), and the fourth angle 196 of the second group of weft fibers 176 can vary between negative thirty-five degrees and negative sixty-five degrees (-35° to -65°). In a non-limiting example, at least 20% of the fiber volume of the surface layer 152 can be arranged at the third angle 194, and at least 20% of the fiber volume of the surface layer 152 can be arranged at the fourth angle 196. In addition, it is contemplated that at least a portion of the surface layer 152 is formed using a woven geometry.

[0080] At 230, method 220 may include applying a coating, such as coating 154 ( Figure 3 ). Coating 154 may be applied to the exterior of surface layer 152. Coating 154 may form an outer wall. Coating 154 may be a barrier coating, for example, to prevent erosion due to impact with an object, however additional properties may be envisioned, such as hydrophobic or ice-phobic properties, anti-corrosion or anti-erosion properties, or UV-resistant properties. Additional finishing layers or materials may be provided on coating 154 as desired, such as an anti-oxidation or anti-corrosion coating or paint.

[0081] The benefit of using a preform core 150 or skin 152 that includes fibers arranged at an angle is that variable local thickness, material, geometry, orientation, physical properties, or mechanical properties can be achieved. In another example, using the first group of fibers 160 and the second group of fibers 172 at the first, second, third, and / or fourth angles 184, 186, 194, 196 can adjust the local vibration frequency of the composite airfoil 100 or composite core 110, and the natural frequency of the composite airfoil 100 can be changed. This adjustment of the natural frequency can be achieved using at least 40% of the fibers oriented at an angle relative to the remaining fibers, or using at least 20% of the warp fibers and 20% of the weft fibers, totaling at least 40% of the total fibers. Changing the natural frequency can adjust the composite airfoil 100 so that the composite airfoil 100 is less susceptible to vibration wear or fatigue, and extends component life and reduces overall maintenance. Changing the natural frequency may be achieved by using the first set of fibers 160 and the second set of fibers 172 at first, second, third, and / or fourth angles 184, 186, 194, 196. Utilizing angled fibers may move the center of rotation or center of mass, thereby providing greater or better balance to the composite airfoil 100 or composite core 110, which may improve operating efficiency and component life.

[0082] This written description uses examples to disclose the disclosure, including the best mode, and also to enable those skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that do not differ substantially from the literal language of the claims.

[0083] The following topics provide further aspects:

[0084] A gas turbine engine comprises: a fan section, a compressor section, a combustion section and a turbine section, wherein the fan section, the compressor section, the combustion section and the turbine section are in a serial flow arrangement and define an engine centerline extending between a forward direction and a rearward direction; and a composite airfoil, wherein the composite airfoil is arranged in one of the fan section, the compressor section or the turbine section, wherein the composite airfoil extends between a root and a tip, wherein a span direction is defined between the root and the tip, and extends between a leading edge and a trailing edge, wherein a chord direction is defined between the leading edge and the trailing edge, wherein the composite airfoil comprises: a preform core, wherein the preform core comprises a first group of fibers, wherein the first group of fibers comprises a first group of warp fibers and a first group of weft fibers; wherein at least 40% of the volume of the first group of fibers is arranged at a first angle relative to the span direction, and wherein the first angle varies between thirty-five degrees and sixty-five degrees (35° to 65°), or varies between negative thirty-five degrees and negative sixty-five degrees (-35° to -65°).

[0085] A gas turbine engine according to any of the preceding clauses, wherein the angle of the first set of warp fibers varies between thirty-five degrees and sixty-five degrees (35° to 65°), and the angle of the first set of weft fibers varies between negative thirty-five degrees and negative sixty-five degrees (-35° to -65°).

[0086] A gas turbine engine as claimed in any preceding clause, wherein the first group of warp fibres forms at least 20% by volume of the first group of fibres, and wherein the first group of weft fibres forms at least 20% by volume of the first group of fibres.

[0087] A gas turbine engine as claimed in any preceding clause, wherein the first set of warp fibres which are not arranged at an angle are aligned with the spanwise direction.

[0088] A gas turbine engine as claimed in any preceding clause, wherein the first set of weft fibres which are not arranged at an angle are aligned with the chord-wise direction.

[0089] A gas turbine engine as claimed in any preceding clause, wherein the first set of fibres of the preform core comprises a braided weave pattern.

[0090] A gas turbine engine as described in any preceding clause, wherein the composite airfoil further comprises a skin positioned externally of the preform core.

[0091] A gas turbine engine as claimed in any preceding clause, wherein the skin further comprises a second set of fibres comprising a second set of warp fibres and a second set of weft fibres.

[0092] A gas turbine engine according to any of the preceding clauses, wherein 40% by volume of the second group of fibers are arranged at an angle relative to the span direction, wherein the angle varies between thirty-five and sixty-five degrees (35° to 65°), or between negative thirty-five and negative sixty-five degrees (-35° to -65°).

[0093] A gas turbine engine according to any of the preceding clauses, wherein the angle of the second set of warp fibers varies between thirty-five degrees and sixty-five degrees (35° to 65°), and the angle of the second set of weft fibers varies between negative thirty-five degrees and negative sixty-five degrees (-35° to -65°).

[0094] A gas turbine engine as claimed in any preceding clause, wherein the second group of warp fibres forms at least 20% by volume of the second group of fibres, and wherein the second group of weft fibres forms at least 20% by volume of the second group of fibres.

[0095] A gas turbine engine as claimed in any preceding clause, wherein the skin layer is a woven layer.

[0096] A gas turbine engine as claimed in any preceding clause, wherein the second set of fibres of the skin layer comprises a braided weave pattern.

[0097] A gas turbine engine as described in any preceding clause, wherein the composite airfoil further comprises a coating positioned externally of the skin.

[0098] A gas turbine engine as claimed in any preceding clause, wherein the preform core comprises a first region comprising at least some of the first set of fibres and defining a first localised increased thickness.

[0099] A gas turbine engine as claimed in any preceding clause, wherein the preform core comprises a second region comprising at least some of the first set of fibres.

[0100] A gas turbine engine as claimed in any preceding clause, wherein the angle of the first set of fibres in the second region is different from the angle of the first set of fibres in the first region.

[0101] A gas turbine engine as claimed in any preceding clause, wherein the localised increased thickness defines a three-dimensional shape of the preform core.

[0102] A gas turbine engine as claimed in any preceding clause, wherein the skin comprises a third region comprising at least some of the second set of fibres and defining a second localised increased thickness.

[0103] A gas turbine engine as claimed in any preceding clause, wherein the skin comprises a fourth region comprising at least some of the second set of fibres.

[0104] A gas turbine engine as claimed in any preceding clause, wherein the angle of the second set of fibres in the third region is different from the angle of the second set of fibres in the fourth region.

[0105] A gas turbine engine as claimed in any preceding clause, wherein at least one of the first local increased thickness or the second local increased thickness changes a mechanical property of the preform core or skin.

[0106] A composite airfoil comprises: a body extending between a root and a tip defining a spanwise direction between the root and the tip, and extending between a leading edge and a trailing edge defining a chordwise direction between the leading edge and the trailing edge; and a preform core comprising a first group of warp fibers, a first group of weft fibers, a second group of warp fibers, and a second group of weft fibers; wherein the first group of warp fibers extend in a first direction, and the first group of weft fibers extend in a second direction; wherein the second group of warp fibers are arranged at a first angle relative to the first direction, the first angle varying between thirty-five and sixty-five degrees (35° to 65°); and wherein the second group of weft fibers are arranged at a second angle relative to the first direction, the second angle varying between negative thirty-five and negative sixty-five degrees (-35° to -65°).

[0107] A composite airfoil as described in any preceding clause wherein at least 20% by volume of the preform core is formed from the second set of warp fibres.

[0108] A composite airfoil as claimed in any preceding clause, wherein at least 20% by volume of the preform core is formed by the second set of weft fibres.

[0109] A composite airfoil as described in any preceding clause, wherein the first direction is aligned with the spanwise direction.

[0110] A composite airfoil as described in any preceding clause, wherein the second direction is aligned with the chord-wise direction.

[0111] A composite airfoil as claimed in any preceding clause, wherein at least 40% by volume of the preform core is formed by the second set of warp fibres and the second set of weft fibres.

[0112] A method of forming a composite core for a composite airfoil for a turbine engine, the composite core extending in a spanwise direction between a root and a tip and in a chordwise direction between a leading edge and a trailing edge, the method of forming the composite core comprising: applying a first group of fibers, wherein the first group of fibers comprises a first group of warp fibers arranged in a first direction and a first group of weft fibers arranged in a second direction perpendicular to the first direction; and applying a second group of fibers, wherein the second group of fibers comprises a second group of warp fibers arranged at a first angle relative to the first direction, the first angle varying between thirty-five and sixty-five degrees (35° to 65°), and comprises a second group of weft fibers arranged at a second angle relative to the first direction, the second angle varying between negative thirty-five and negative sixty-five degrees (-35° to -65°).

[0113] A method according to any preceding clause, wherein the second plurality of warp fibers forms at least 20% by volume of the composite core.

[0114] A method according to any preceding clause, wherein the second set of weft fibers forms at least 20% by volume of the composite core.

[0115] A method as in any preceding clause wherein the second group of fibers forms at least 40% by volume of the composite core.

[0116] A method as in any preceding clause, further comprising rotating the composite core.

[0117] A method as in any preceding clause, wherein the rotating composite core defines a first angle and a second angle.

[0118] A method as described in any preceding clause, further comprising applying a skin layer external to the composite core.

[0119] A method as in any preceding clause, wherein the surface layer comprises a third set of fibers having a third set of warp fibers and a third set of weft fibers.

[0120] A method as in any preceding clause, wherein at least 20% by volume of the third group of warp fibers are arranged at a third angle, and at least 20% by volume of the third group of weft fibers are arranged at a fourth angle.

[0121] A method according to any of the preceding clauses, wherein a third group of warp fibers are arranged at a third angle relative to the first direction, the third angle varying between thirty-five and sixty-five degrees (35° to 65°), and a third group of weft fibers are arranged at a fourth angle relative to the first direction, the fourth angle varying between negative thirty-five and negative sixty-five degrees (-35° to -65°).

[0122] A method as in any preceding clause, wherein the second set of fibers comprises at least one of carbon, glass, nylon, rayon, aramid, nickel, titanium, or ceramic.

[0123] A method as in any preceding clause, wherein the first angle and the second angle define a stiffness, elasticity or strength of the composite core.

[0124] A method as in any preceding clause, wherein the first set of fibers and the second set of fibers form a preform.

[0125] A method as in any preceding clause, wherein the preform comprises a resin transfer moulding, a precured resin transfer moulding, a prepreg laminate, or a dry form and prepreg formed from the same qualified resin transfer moulding.

[0126] A preform for a component in a turbine engine, the preform comprising: a preform core extending in a first direction and a second direction perpendicular to the first direction, the preform core comprising: a first group of warp fibers, the first group of warp fibers are aligned with the first direction; a first group of weft fibers, the first group of weft fibers are aligned with the second direction; a second group of warp fibers, the second group of warp fibers are arranged at a first angle relative to the first direction, the first angle varying between thirty-five and sixty-five degrees (35° to 65°); and a second group of weft fibers, the second group of weft fibers are arranged at a second angle relative to the first direction, the second angle varying between negative thirty-five and negative sixty-five degrees (-35° to -65°).

[0127] A preform as claimed in any preceding clause wherein at least 20% by volume of the preform core is formed by the second set of warp fibres.

[0128] A preform as claimed in any preceding clause wherein at least 20% by volume of the preform core is formed by the second set of weft fibres.

[0129] A preform as claimed in any preceding clause wherein at least 40% by volume of the preform core is formed by the second set of warp fibres and the second set of weft fibres.

[0130] A preform according to any preceding clause, further comprising a skin layer arranged externally to the preform core.

[0131] A preform according to any preceding clause, wherein the skin further comprises a third group of fibres comprising a third group of warp fibres and a third group of weft fibres.

[0132] A preform according to any of the preceding clauses, wherein 40% by volume of the third group of fibers are arranged at a third angle relative to the first direction, and wherein the third angle varies between thirty-five degrees and sixty-five degrees (35° to 65°) or between negative thirty-five degrees and negative sixty-five degrees (-35° to -65°) relative to the first direction.

[0133] A preform according to any of the preceding clauses, wherein the third angle of the third group of warp fibers varies between thirty-five degrees and sixty-five degrees (35° to 65°), and the third angle of the third group of weft fibers varies between negative thirty-five degrees and negative sixty-five degrees (-35° to -65°).

[0134] A preform according to any preceding clause, wherein the third group of warp fibers forms at least 20% by volume of the third group of fibers, and wherein the third group of weft fibers forms at least 20% by volume of the third group of fibers.

[0135] A preform as claimed in any preceding clause, wherein the skin comprises a braided weave pattern.

[0136] A preform according to any preceding clause, wherein the preform core is at least partially formed from a resin transfer moulding, a precured resin transfer moulding, a prepreg laminate, or a preform and prepreg formed from the same qualified resin transfer moulding.

Claims

1. A gas turbine engine, characterized in that: include: a fan section, a compressor section, a combustion section, and a turbine section, the fan section, the compressor section, the combustion section, and the turbine section being in a serial flow arrangement and defining an engine centerline extending between a forward direction and an aft direction; as well as A composite airfoil disposed in one of the fan section, the compressor section, or the turbine section, the composite airfoil extending between a root and a tip defining a spanwise direction therebetween, and extending between a leading edge and a trailing edge defining a chordwise direction therebetween, the composite airfoil comprising: a preform core comprising a first set of fibers including a first set of warp fibers and a first set of weft fibers; Wherein at least 40% of the volume of the first group of fibers is arranged at a first angle relative to the spanwise direction, and wherein the first angle varies between thirty-five degrees and sixty-five degrees (35° to 65°), or between negative thirty-five degrees and negative sixty-five degrees (-35° to -65°).

2. The gas turbine engine according to claim 1, characterized in that The angle of the first group of warp fibers varies between thirty-five and sixty-five degrees (35° to 65°), and the angle of the first group of weft fibers varies between negative thirty-five and negative sixty-five degrees (-35° to -65°).

3. The gas turbine engine according to claim 2, characterized in that wherein said first group of warp fibers forms at least 20% of said volume of said first group of fibers, and wherein said first group of weft fibers forms at least 20% of said volume of said first group of fibers.

4. The gas turbine engine according to claim 1, characterized in that The first group of warp fibers that are not arranged at the angle are aligned with the spanwise direction.

5. The gas turbine engine according to claim 4, characterized in that The first group of weft fibers that are not arranged at the angle are aligned with the chord-wise direction.

6. The gas turbine engine according to claim 1, characterized in that Wherein the first group of fibers of the preform core comprises a braided weave pattern.

7. The gas turbine engine according to claim 1, characterized in that Wherein the composite airfoil further comprises a skin positioned outside the preform core.

8. The gas turbine engine according to claim 7, characterized in that The surface layer further comprises a second group of fibers, wherein the second group of fibers comprises a second group of warp fibers and a second group of weft fibers.

9. The gas turbine engine according to claim 8, characterized in that Wherein 40% of the volume of the second group of fibers is arranged at a second angle relative to the spanwise direction, and wherein the second angle varies between thirty-five degrees and sixty-five degrees (35° to 65°), or between negative thirty-five degrees and negative sixty-five degrees (-35° to -65°).

10. The gas turbine engine according to claim 9, characterized in that The angle of the second group of warp fibers varies between thirty-five and sixty-five degrees (35° to 65°), and the angle of the second group of weft fibers varies between negative thirty-five and negative sixty-five degrees (-35° to -65°).

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