Airfoil assembly with composite core

Abstract

An airfoil assembly and an engine component for a turbine engine having an engine core extending along an engine centerline, the airfoil assembly having an airfoil extending in a radial direction away from the engine centerline between a root and a tip to define a span length. The airfoil includes a composite core and a set of skins defining at least a portion of an exterior surface of the airfoil. The shank extends in a radial direction from the root toward the engine centerline and has a flared cross-section.

Description

Technical Field

[0001] The present disclosure relates generally to airfoil assemblies and, more particularly, to airfoil assemblies having a composite core. Background Art

[0002] A turbine engine (and in particular a gas or combustion turbine engine) is a rotary engine that extracts energy from a gas flow that passes through a fan having a plurality of fan blades, then through a series of compressor stages (which include pairs of rotating blades and stationary vanes), through a combustor, and then into the engine through a series of turbine stages (which include pairs of rotating blades and stationary vanes). The blades are mounted to a rotating disk, and the vanes are mounted to a stator disk.

[0003] During operation, air enters the compressor section through the fan section, is then pressurized in the compressor, and mixed with fuel in the combustor to generate hot combustion gases, which flow downstream through the turbine stage, where the air expands, and exits the exhaust section. The expansion of the air in the turbine section is used to drive the rotating sections of the fan section and the compressor section. The intake of air, the pressurization of air, and the expansion of air are accomplished in part by the rotation of various rotating blades mounted to respective disks throughout the fan section, the compressor section, and the turbine section, respectively.

[0004] Composite materials typically include a fiber reinforced matrix and exhibit a high strength-to-weight ratio. Due to the high strength-to-weight ratio and formability to 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 wing, rudder, manifold, airfoil, or other component of an aircraft or turbine engine, or define a portion thereof. Extreme loads, sudden forces, or heat can be applied to composite components of an aircraft or turbine engine. For example, during the intake of various materials by a turbine engine, extreme loads may occur to one or more airfoils. BRIEF DESCRIPTION OF THE DRAWINGS

[0005] 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:

[0006] Figure 1 is a schematic cross-sectional view of a turbine engine according to an exemplary embodiment of the present disclosure.

[0007] Figure 2 Is suitable for Figure 1 A schematic perspective view of an airfoil assembly and a disk assembly used within a turbine engine of FIG. 1 , the airfoil assembly including an airfoil and a dovetail.

[0008] Figure 3 is from Figure 2A schematic cross-sectional front view of a section of the airfoil assembly as seen through section line III-III, further illustrating a composite core, a foam core, and a set of skins.

[0009] Figure 4 is an enlarged schematic cross-section of a composite core and a foam core having a set of Z fibers according to aspects disclosed herein.

[0010] Figure 5 is an enlarged schematic cross-section of a composite core having a set of Z fibers, a foam core, and a set of skins according to another aspect disclosed herein.

[0011] Figure 6 is an enlarged schematic diagram of a composite core having a set of Z fibers, a foam core, and a set of skins according to yet another aspect disclosed herein. DETAILED DESCRIPTION

[0012] Aspects disclosed herein relate to a turbine engine including an airfoil assembly having a composite core and a set of skins defining at least a portion of a dovetail and an airfoil. The airfoil assembly can be used in one or more locations within a turbine engine. For example, the airfoil assembly is suitable as a fan blade in a fan section of a turbine engine, but other locations (such as a compressor section and a turbine section) are also conceivable. The airfoil assembly can be installed in a variety of ways. One such installation is to fix the blade directly or via a pitch control assembly to a rotator of the fan section. Regardless of where the airfoil assembly is located, a suitable mounting member is a disk assembly having complementary sockets to receive dovetails, wherein the sockets are circumferentially spaced around the periphery of the disk assembly. The airfoil assembly and the disk assembly can form a rotating assembly together so that the airfoil assembly is a composite blade assembly.

[0013] The airfoil assembly also includes a foam core and a group of Z fibers in the composite core, which extend at least partially through the composite core, the group of skins and one of the foam cores. The airfoil assembly is designed to withstand the high mechanical stresses associated with the operation of the turbine engine while providing a lightweight structure. For the purpose of illustration, the present disclosure will be described with respect to an airfoil assembly for a turbine engine, specifically, with respect to an airfoil assembly disposed in a fan section of a turbine engine. However, it will be understood that the aspects of the present disclosure described herein are not limited thereto and may have general applicability in other engines or in other parts of a turbine engine. For example, the present disclosure may be applicable to airfoil assemblies in other engines or carriers, and may be used to provide benefits in industrial, commercial and residential applications. In addition, aspects of the present disclosure will be directed to a composite blade assembly including a dovetail. It should be understood that aspects of the present disclosure are directed to any airfoil assembly. Although shown as having a dovetail, the airfoil assembly may include a T-shaped flange root. In addition, the airfoil assembly is not limited thereto and may be a composite blade assembly or a composite blade assembly.

[0014] 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 "frontward" mean in front of something, and "rear" or "rearward" mean behind something. For example, when used in relation to fluid flow, front / front may refer to upstream, and rear / rearward may refer to downstream.

[0015] Furthermore, as used herein, the term "radial" or "radially" refers to a direction away from a common center. For example, in the overall context of a turbine engine, radial refers to a direction along a ray extending between a central longitudinal axis of the engine and an outer periphery of the engine. Furthermore, as used herein, the term "group" or a "group" of elements may be any number of elements, including only one.

[0016] Additionally, as used herein, the term "fluid" or iterations thereof may refer to any suitable fluid within a gas turbine engine, wherein at least a portion of the gas turbine engine is exposed to, for example, but not limited to, combustion gases, ambient air, a pressurized gas stream, a working gas stream, or any combination thereof. It is further contemplated that the gas turbine engine may be other suitable turbine engines, such as, but not limited to, a steam turbine engine or a supercritical carbon dioxide turbine engine. As non-limiting examples, the term "fluid" may refer to steam in a steam turbine engine, or to carbon dioxide in a supercritical carbon dioxide turbine engine.

[0017] All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, rear, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, backward, etc.) are used for identification purposes only to help the reader understand the present disclosure and do not create limitations, particularly limitations on the position, orientation, or use of the disclosed aspects described herein. Connection references (e.g., attach, couple, fix, fasten, connect, and engage) will be interpreted broadly and may include intermediate members between sets of elements and relative movement between elements, unless otherwise indicated. Therefore, a connection reference does not necessarily mean that the two elements are directly connected and fixed relative to each other. The exemplary drawings are for illustration purposes only, and the dimensions, positions, orders, and relative sizes reflected in the attached drawings may vary.

[0018] 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.

[0019] 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.

[0020] 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 require curing at elevated temperatures or other hardening techniques.

[0021] 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.

[0022] 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.

[0023] 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.

[0024] 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.

[0025] 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.

[0026] 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.

[0027] 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.

[0028] As used herein, CMC refers to a class of materials having reinforcing fibers in a ceramic matrix. Typically, the reinforcing fibers provide structural integrity to the ceramic matrix. Some examples of reinforcing fibers may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), non-oxide carbon-based materials (e.g., carbon), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates (such as mullite), or mixtures thereof), or mixtures thereof.

[0029] Some examples of ceramic matrix materials may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates, or mixtures thereof), or mixtures thereof. Optionally, ceramic particles (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.

[0030] In general, specific CMCs can be referred to as their fiber type / matrix type combination. For example, C / SiC is carbon fiber reinforced silicon carbide, SiC / SiC is silicon carbide fiber reinforced silicon carbide, SiC / SiN is silicon carbide fiber reinforced silicon nitride, SiC / SiC-SiN is silicon carbide fiber-reinforced silicon carbide / silicon nitride matrix mixture, etc. In other examples, CMCs can be composed of a matrix and reinforcing fibers comprising oxide-based materials such as aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates, and mixtures thereof. Aluminosilicates can include crystalline materials such as mullite (3Al2O3·2SiO2), as well as glassy aluminosilicates.

[0031] 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 bundles may be formed into reinforcing tapes, such as unidirectional reinforcing tapes. Multiple tapes may be stacked together to form a preform component. The fiber bundles may be impregnated with a slurry composition before or after forming the preform. The preform may then be subjected to heat treatment and subsequent chemical treatment to obtain a component formed of a CMC material having a desired chemical composition. For example, the preform may undergo curing or burnout to produce a high coke residue in the preform, and then be melt infiltrated with silicon, or undergo curing or pyrolysis to produce a silicon carbide matrix in the preform, and then be chemically infiltrated 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.

[0032] Such materials, along with certain monolithic ceramics (i.e., ceramic materials without reinforcement materials), are particularly useful for higher temperature applications. In addition, these ceramic materials are lightweight compared to superalloys, yet still provide strength and durability to components made therefrom. As a result, such materials are currently being considered for use in many gas turbine components used in the higher temperature sections of gas turbine engines (such as airfoils (e.g., turbines and buckets), combustors, shrouds, and other components) that would benefit from the lighter weight and higher temperature capabilities that these materials can provide.

[0033] As used herein, "foam core" is a rigid structural foam that has high resistance to shear and compression, even at elevated temperatures.

[0034] 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.

[0035] As used herein, "modulus" refers to the resistance of an object or material to deformation when stress is applied.

[0036] As used herein, "skin" refers to a layer of material having a plurality of composite layups or a plurality of composite layers.

[0037] Figure 1is 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 front portion 14 to an aft portion 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.

[0038] 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 generates combustion gases. The engine core 44 is surrounded by a core case 46, which may be coupled to the fan case 40.

[0039] 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 plurality of rotatable elements that may collectively define a rotor 51.

[0040] 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 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.

[0041] Compressor blades 56, 58 for one stage of the compressor may be mounted to (or integrated into) a disk 61 mounted to a corresponding one of the HP and LP spools 48, 50. Static compressor vanes 60, 62 for one stage of the compressor may be mounted to the core housing 46 in a circumferential arrangement.

[0042] 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 illustrative purposes only, and other numbers are possible.

[0043] Turbine blades 68, 70 for one stage of the turbine may be mounted to a disk 71 mounted to a corresponding one of the HP and LP spools 48, 50. Static turbine buckets 72, 74 for one stage of the compressor may be mounted to the core housing 46 in a circumferential arrangement.

[0044] 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 may refer to the combination of non-rotating elements throughout the turbine engine 10.

[0045] 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 the 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 the fuel in the combustor 30 and ignited, thereby generating combustion gases. The HP turbine 34 extracts some work from these gases, which drives the HP compressor 26. The combustion gases are exhausted into the LP turbine 36, which extracts additional work to drive the LP compressor 24, and the exhaust gas is ultimately 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.

[0046] 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 increased to above the bleed air temperature. The bleed air 77 may be used to reduce the temperature of core components downstream of the combustor 30.

[0047] The remaining portion of the pressurized airflow 76 bypasses the LP compressor 24 and the engine core 44 as a bypass airflow 78 and exits the turbine engine 10 through a stationary vane row at the fan exhaust side 84, and more specifically, exits the turbine engine 10 through an exit guide vane assembly 80 (including a plurality of airfoil guide vanes 82) at the fan exhaust side 84. More specifically, a circumferential row of radially extending airfoil guide vanes 82 are used adjacent the fan section 18 to exert some directional control on the airflow 78.

[0048] Some 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 case of a turbine engine, the hot portions of the engine are typically 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 discharged from the LP compressor 24 or the HP compressor 26.

[0049] Figure 2 Is suitable for Figure 1 Schematic perspective view of an airfoil assembly 104 and a disk assembly 102 for use in a turbine engine 10 of FIG. The disk assembly 102 is suitable for use as a disk 61, 71 ( Figure 1 ) or any other disk, such as, but not limited to, the fan section 18 ( Figure 1 )、Compressor section 22( Figure 1 ) or turbine section 32 ( Figure 1 ) within the disk. The airfoil assembly 104 may be rotating or non-rotating, such that the airfoil assembly 104 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 a plurality of fan blades 42 ( Figure 1 ). As a non-limiting example, the airfoil assembly 104 may be a composite fan blade assembly.

[0050] The disk assembly 102 may rotate or be stationary about an axis of rotation 106. The axis of rotation 106 may be aligned with an engine centerline (e.g., Figure 1The disk assembly 102 includes a plurality of slots 108 extending axially through a radially outer portion of the disk assembly 102 and spaced circumferentially about the disk assembly 102 relative to the rotational axis 106. The rotational axis 106 may define an axial direction (denoted as "A"). A radial direction (denoted as "R") extends radially outward and perpendicular to the axial direction A. In addition, a circumferential direction (denoted as "C") relative to the rotational axis 106 is shown, and the circumferential direction C extends perpendicular to the axial direction A and the radial direction R.

[0051] The airfoil assembly 104 may be any engine component having a shank, or a part by which an engine component may be attached to another engine component. As a non-limiting example, the engine component is an airfoil 110, and the shank is a dovetail 112 extending from the airfoil 110. The shank may be any suitable attachment geometry, including a T-shaped flange root. The airfoil 110 extends between a leading edge 114 and a trailing edge 116 to define a chord-wise direction. The airfoil 110 extends between a root 118 and a tip 120 to define a span length (denoted as "SL") extending in a radial direction R. The root defines 0% of the span length SL, and the tip defines 100% of the span length SL. The airfoil 110 includes an airfoil exterior surface 144 defining a pressure side 122 and a suction side 124 of the airfoil 110. The dovetail 112 extends between a first end 126 and a second end 128 in a radial direction R. The first end 126 represents a transition between the dovetail 112 and the airfoil 110. As a non-limiting example, the first end 126 coincides with the root 118 of the airfoil 110. The dovetail 112 and the airfoil 110 may be integrally or non-integrally formed with one another.

[0052] The airfoil assembly 104 is coupled to the disk assembly 102 by inserting at least a portion of the dovetail 112 into a corresponding slot in the plurality of slots 108 such that the second end 128 faces the disk assembly 102. The airfoil assembly 104 is held in place by frictional contact with the slot 108, or may be coupled to the slot 108 via any suitable coupling method, such as, but not limited to, welding, bonding, fastening, etc. Although only a single airfoil assembly 104 is shown, it should be understood that there may be any number of one or more composite airfoil assemblies 104 coupled to the disk assembly 102. As a non-limiting example, there may be a number of composite airfoil assemblies 104 corresponding to the total number of slots in the plurality of slots 108. The airfoil assemblies 104 as discussed herein allow for effective load transfer from the airfoil 110 to the disk assembly 102 through the dovetail 112.

[0053] Figure 3 is from Figure 2Schematic cross-sectional front view of a portion of the airfoil assembly 104 as viewed along section line III-III. The airfoil assembly 104 extends from the first end 126 of the dovetail 112 to the tip 120 ( Figure 2 ) extends from a centerline axis CA. The centerline axis CA is equidistant from respective opposing portions of the airfoil assembly 104. The centerline axis CA may be linear or non-linear. The airfoil assembly 104 may be symmetrical or asymmetrical about the centerline axis CA. The centerline axis CA may also be a body axis, which may not extend along the center of the body in the case of an asymmetrical airfoil. The dovetail 112 may have a flared cross-section 134 relative to the centerline axis CA.

[0054] The airfoil assembly 104 includes a composite core 130 that defines and extends between the dovetail 112 and portions of the airfoil 110. The composite core 130 may extend from the second end 128 to the tip 120 ( Figure 2 ). The foam core 138 is located within the composite core 130. The composite core 130 may define various portions of the airfoil 110 and the dovetail 112, while the foam core 138 defines the core of the airfoil 110.

[0055] The foam core 138 may extend along the centerline axis CA and define at least a portion of the airfoil 110. The foam core 138 may include a tapered end point 139 that is spaced radially outward from the root 118 and within the composite core 130. The tapered end point 139 is located between 0% and 5% of the span length SL of the airfoil. In one non-limiting example, the foam core 138 may extend from the root 118 to the tip 120 ( Figure 2 ), in other words, extending the entire length of the airfoil 110. In another non-limiting example, the foam core 138 extends from 20% to 80% of the span length SL.

[0056] A set of skins 140 covers the composite core 130. As a non-limiting example, the skins referred to herein may be defined as laminated skins, woven or braided skins, or any combination thereof. The set of skins 140 may be any suitable type of skin. The set of skins 140 may be skin layers that are layered in direct contact with adjacent skins in the airfoil 110. The set of skins 140 may define various portions of the airfoil 110 and the dovetail 112. The set of skins 140 may define at least a portion of an airfoil exterior surface 144. The airfoil 110 extends a certain thickness (indicated as "T") between the pressure side 122 and the suction side 124.

[0057] Figure 4 Is from Figure 3FIG. 1 is an enlarged view of the composite core 130 and the foam core 138 of section IV in FIG. The middle plane MP is a plane oriented at the center of the airfoil 110 and extending along the centerline axis CA. Although shown as extending along the radial direction R, the middle plane MP can be curved. A group of Z fibers 136 passes through the foam core 138 and extends through the foam core 138 into the composite core 130. The group of Z fibers 136 may include at least one first direction fiber 142a extending linearly through the middle plane MP along a first direction (denoted as "a") at a first fiber angle α. The first fiber angle α is measured between the middle plane MP and the at least one first direction fiber 142a. The first fiber angle α is in the range of 30° to 90° relative to the middle plane MP. The group of Z fibers may also include at least one second direction fiber 142b extending linearly through the middle plane MP along a second direction (denoted as "b") at a second fiber angle β. The second fiber angle β is in the range of -30° to 90° relative to the middle plane MP.

[0058] The group of Z fibers 136 includes fibers that extend through the mid-plane MP and span 5% to 100% of the thickness T of the airfoil 110. For example, the group of Z fibers 136 may include at least one full-span fiber 142f (or a group of fibers 142f, also referred to below as a fully extended group of fibers) that extends through the mid-plane MP and spans the entire thickness T or 100% of the thickness T. By further example, the group of Z fibers 136 may include at least one partial-span fiber 142p that extends through the mid-plane MP while spanning only less than 100% of the thickness T (such as 5%, 20%, 30%, 50%, or 70%). It is further contemplated that the group of Z fibers 136 includes any number of partial-span fibers 142p that extend through the mid-plane MP and span 5% to less than 100% of the thickness T. The partial-span fiber 142p (or a group of fibers 142p) may extend through the mid-plane MP and span all or only a portion of the thickness of the foam core 138 to reinforce the foam core 138. Partial span fibers 142p (or a group of fibers 142p, also referred to hereinafter as a group of fibers extending intermediately) may extend through the intermediate plane MP and span all or only a portion of the thickness of the composite core 130, both reinforcing the foam core and connecting the foam core to the composite core. In addition, partial span fibers 142p (or a group of fibers 142p) may extend through the intermediate plane MP and span a portion of the thickness of the group of skins 140 to reinforce the foam core 138 and connect the foam core 138, the composite core 130, and the group of skins 140 together.

[0059] It should be understood that any of the first direction fibers 142a may also be full span fibers 142f or partial span fibers 142p. In addition, any of the second direction fibers 142b may also be full span fibers 142f or partial span fibers 142p.

[0060] Figure 5 Is from Figure 4 FIG. 1 is an enlarged schematic diagram of the composite core 130, the foam core 138, and the set of skins 140 of section V at the bottom of FIG. As can be more easily seen, the set of skins 140 is a plurality of layers 148. The plurality of layers 148 may include at least one laminated skin or at least one woven skin. The plurality of layers 148 may also include at least one set of filler plies or a set of machined plies. The plurality of layers 148 may also include a prepreg laminate.

[0061] The composite core 130 may include a composite layer 132 surrounding a foam core 138. The composite layer 132 may be a 3D composite layer in the form of a fiber preform or a braided core or a woven composite preform. The composite layer 132 may be an interference preform or a precured preform. When a braiding process is used to form the composite core 130, the composite layer 132 may include at least one material 133 oriented in three different directions (commonly referred to as warp, weft, and binder orientation). Although described as a braiding process, it should be understood that any process that results in the orientation described herein is contemplated. The composite layer 132 may be combined with the set of skins 140 by resin transfer molding (RTM) or same qualified resin transfer molding (SQ-RTM) to form a single composite component.

[0062] The set of Z fibers 136 includes a fully extended set of fibers 170, a first intermediately extended set of fibers 172, and a second intermediately extended set of fibers 174. The fully extended set of fibers 170 (e.g., full span fibers 142f) extend through the entire foam core 138, the composite core 130, and the set of skins 140. The first intermediately extended set of fibers 172 (e.g., partial span fibers 142p) extend through the entire foam core 138 and the composite core 130. The second intermediately extended set of fibers 174 (e.g., partial span fibers 142p) extend through the entire foam core 138 and the composite core 130.

[0063] Figure 61 is an enlarged view of a composite core 130a, a foam core 138a, and a set of skins 140a according to another aspect disclosed herein. In another exemplary arrangement, the set of Z fibers 136 can extend through any combination of the foam core 138a, the composite core 130a, and the set of skins 140a. The first fibers 150 can be second direction fibers 142b and partial span fibers 142p extending through the foam core 138a and the composite core 130a. The second fibers 152 can be first direction fibers 142a and full span fibers 142f extending through the foam core 138a, the composite core 130a, and the set of skins 140a. The third fibers 154 can be first direction fibers 142a and partial span fibers 142p extending through the foam core 138a, the first side 162 of the composite core 130a, and the first side 164 of the set of skins 140a. The fourth fiber 156 may be a first direction fiber 142a and a partial span fiber 142p extending through the foam core 138a, the composite core 130a, and the second side 166 of the set of skins 140a. The fifth fiber 158 may be a first direction fiber 142a and a partial span fiber 142p extending through the foam core 138a and the composite core 130a. The sixth fiber 160 may be a second direction fiber 142b and a partial span fiber 142p extending through the foam core 138a, the second side 168 of the composite core 130a, and the second side 166 of the set of skins 140a. It should be understood that any combination of the first, second, third, fourth, fifth, and sixth fibers 150, 152, 154, 156, 158, 160 may be contemplated. Furthermore, it should be understood that the first, second, third, fourth, fifth, and sixth fibers 150, 152, 154, 156, 158, 160 are exemplary and are not meant to be limiting.

[0064] Benefits associated with using the airfoil assemblies described herein include having a lighter airfoil assembly, without sacrificing the performance of the airfoil assembly, compared to a non-composite (e.g., cast) airfoil assembly. In other words, the material used for the airfoil assembly is lighter than the material used for the non-composite airfoil assembly. In addition, the materials used in the airfoil assemblies described herein do not sacrifice the ability to perform as intended within the turbine engine. The reduced weight, in turn, means an increased efficiency of the turbine engine when compared to a conventional turbine engine that includes a non-airfoil assembly.

[0065] Further benefits associated with engine components (such as the airfoil assemblies described herein) include the ability of the airfoil assemblies to withstand mechanical stresses associated with the operation of the airfoil assemblies when compared to conventional airfoil assemblies. The airfoil assemblies disclosed herein having a set of skins with high in-plane modulus and strength for carrying in-plane bending loads are most effective when subjected to bending loads. The foam core and composite core resist lateral shear forces. In addition, the lateral modulus of the foam core and composite core affects the overall airfoil bending stiffness. The disclosure herein provides Z fibers to increase the foam core modulus and strength. In addition, the set of skins, composite core, and foam core are connected together via Z fibers, which improves the resistance of the airfoil under loads such as foreign matter ingestion by effectively transferring the load from the set of skins to the foam core in the thickness direction.

[0066] 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.

[0067] This written description uses examples to describe the aspects of the disclosure described herein, including the best mode, and also to enable any person skilled in the art to practice aspects of the disclosure, including making and using any device or system and performing any combined methods. The patentable scope of aspects of the disclosure is defined by the claims, and may include other examples that occur to one skilled in the art. These other examples are intended to fall within the scope of the claims if they have 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.

[0068] Further aspects are provided by the subject matter of the following clauses:

[0069] An airfoil assembly for a turbine engine having an engine core extending along an engine centerline, the airfoil assembly comprising: an airfoil extending in a radial direction away from the engine centerline between a root and a tip to define a span length, the airfoil comprising a composite core having a set of Z fibers extending therethrough, and a set of skins covering the composite core and defining at least a portion of an exterior surface of the airfoil; and a shank extending radially downwardly in the radial direction from the root toward the engine centerline between a first end and a second end, the shank comprising the composite core and the set of skins and having a flared cross-section.

[0070] An airfoil assembly as recited in any preceding clause, wherein said set of Z fibers extends through at least a portion of said foam core and said composite core.

[0071] An airfoil assembly for a turbine engine having an engine core extending along an engine centerline, the airfoil assembly comprising: an airfoil extending in a radial direction away from the engine centerline between a root and a tip to define a span length, the airfoil comprising a composite core and a set of skins covering the composite core and defining at least a portion of an exterior surface of the airfoil; a set of Z fibers extending through the middle of the airfoil; and a shank extending in the radial direction from the root toward the engine centerline between a first end and a second end, the shank comprising the composite core and the set of skins and having a flared cross-section.

[0072] An airfoil assembly as recited in any preceding clause, further comprising a foam core within the composite core, wherein the set of Z fibers extend through the foam core and at least a portion of the composite core.

[0073] An airfoil assembly according to any preceding clause, wherein a mid-plane extends along a centerline axis CA of the airfoil center, and wherein the set of Z fibers includes first direction fibers extending through the airfoil mid-plane in a first direction at a first fiber angle, the first fiber angle being measured between an individual fiber and the mid-plane.

[0074] An airfoil assembly as described in any preceding clause, wherein the first fiber angle is between 30 and 90 degrees.

[0075] An airfoil assembly as in any preceding clause, wherein the set of Z fibers comprises second direction fibers extending through the airfoil midplane in a second direction at a second fiber angle measured between a single first fiber and the midplane.

[0076] An airfoil assembly as claimed in any preceding clause, wherein the first fiber angle is between 30 and 90 degrees and the second fiber angle is between -30 and 90 degrees.

[0077] An airfoil assembly as claimed in any preceding clause, wherein the airfoil defines a thickness and the set of Z fibres extends through a mid-plane at the centre of the airfoil and spans between 5% and 100% of the thickness.

[0078] An airfoil assembly as described in any preceding clause, wherein the set of Z fibers extends through at least a portion of the set of skins.

[0079] An airfoil assembly as described in any preceding clause, wherein the set of Z fibers extend through the entire thickness of the airfoil.

[0080] An airfoil assembly as described in any preceding clause, wherein the composite core comprises 3D composite layers.

[0081] An airfoil assembly according to any preceding clause, wherein the 3D composite layer is a preform or pre-cured preform and is bonded with the set of skins by resin transfer molding (RTM) or same quality resin transfer molding (SQ-RTM) to form a single composite part.

[0082] An airfoil assembly as claimed in any preceding clause, wherein the set of skins comprises at least one laminate skin or at least one woven skin.

[0083] An airfoil assembly as described in any preceding clause, wherein the set of skins comprises at least one of a filler layup or a machined layup.

[0084] An airfoil assembly as claimed in any preceding clause, wherein the set of skins are prepreg laminates.

[0085] An airfoil assembly as claimed in any preceding clause, wherein the 3D composite layer comprises a woven composite preform formed from at least one material oriented in three different directions.

[0086] An airfoil assembly as claimed in any preceding clause, wherein the foam core comprises a tapered end point located radially outwardly of the root portion and within the composite core.

[0087] An airfoil assembly as recited in any preceding clause, wherein said root defines 0% of said span length, said tip defines 100% of said span length, and said tapered end point is between 0% and 5% of said span length.

[0088] An airfoil assembly according to any preceding clause, wherein the composite core and at least a portion of the set of skins each comprise a composite material, the composite material comprising at least one of a polymer matrix composite, a ceramic matrix composite, a metal matrix composite, carbon fiber, a polymer resin, a thermoplastic, a bismaleimide, a polyimide, an epoxy resin, a glass fiber, or a silicon matrix.

[0089] An engine component for a turbine engine, the engine component comprising: a composite core; a foam core, the foam core being located within the composite core; a set of skins, the set of skins covering the composite core and defining at least a portion of an exterior surface of the component; and a shank, the shank extending radially below the root between a first end and a second end, the shank including the composite core and the set of skins and having a flared cross-section.

[0090] An engine component as recited in any preceding clause, further comprising a set of Z fibers extending through at least a portion of said foam core and said composite core.

Claims

1. An airfoil assembly for a turbine engine having an engine core extending along an engine centerline, characterized in that: The airfoil assembly comprises: an airfoil extending in a radial direction away from the engine centerline between a root and a tip to define a span length, the airfoil comprising a composite core, a foam core disposed within the composite core and having a set of Z fibers extending therethrough, and a set of skins covering the composite core and defining at least a portion of an exterior surface of the airfoil; and A shank portion extends in the radial direction from the root portion toward the engine centerline between a first end and a second end, the shank portion including the composite core and the set of skins and having a flared cross-section.

2. The airfoil assembly according to claim 1, wherein: in, The set of Z fibers extends through at least a portion of the foam core and the composite core.

3. The airfoil assembly according to claim 2, wherein: in, A midplane extends along a centerline axis of the airfoil center, and wherein the set of Z fibers includes first direction fibers extending through the airfoil midplane in a first direction at a first fiber angle measured between an individual fiber and the midplane.

4. The airfoil assembly according to claim 3, characterized in that in, The first fiber angle is 30 to 90 degrees.

5. The airfoil assembly according to claim 3, characterized in that: in, The set of Z fibers includes second direction fibers extending in a second direction through the airfoil midplane at a second fiber angle measured between a single first fiber and the midplane.

6. The airfoil assembly according to claim 5, characterized in that in, The first fiber angle is 30 to 90 degrees, and the second fiber angle is -30 to 90 degrees.

7. The airfoil assembly according to claim 2, wherein: in, The airfoil defines a thickness, and the set of Z fibers extends through a mid-plane at a center of the airfoil and spans 5% to 100% of the thickness.

8. The airfoil assembly according to claim 2, wherein: in, The set of Z fibers extends through at least a portion of the set of epidermis.

9. The airfoil assembly according to claim 8, wherein: in, The set of Z fibers extends through the entire thickness of the airfoil.

10. An airfoil assembly for a turbine engine having an engine core extending along an engine centerline, characterized in that: The airfoil assembly comprises: an airfoil extending in a radial direction away from the engine centerline between a root and a tip to define a span length, the airfoil including a composite core and a set of skins covering the composite core and defining at least a portion of an exterior surface of the airfoil; a set of Z fibers extending through the middle of the airfoil; and A shank portion extends in the radial direction from the root portion toward the engine centerline between a first end and a second end, the shank portion including the composite core and the set of skins and having a flared cross-section.

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