Composite airfoil assembly with dovetail

Abstract

A composite airfoil assembly for a turbine engine includes an airfoil extending in a radial direction between a root and a tip to define a span length. The airfoil includes a composite core and a set of skins covering the composite core. The set of skins includes an outer skin defining at least a portion of an exterior surface of the airfoil. A dovetail extends radially below the root, the dovetail comprising a composite core and a set of skins, the dovetail further comprising a fan-shaped skin covering the outer skin near the root.

Description

Technical Field

[0001] The present disclosure relates generally to composite airfoil assemblies having dovetails, and more particularly, to composite airfoil assemblies having dovetails within turbine engines. 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 is 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. The rotation of the rotating blades applies mechanical stresses along various portions of the blades; in particular, mechanical stresses are applied along the locations where the blades are mounted to the disks.

[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 1A schematic perspective view of a composite airfoil assembly and a disk assembly for use within a turbine engine of the type shown in FIG. 1 , the composite airfoil assembly including an airfoil and a dovetail.

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

[0009] Figure 4 It is an enlarged schematic stereogram of a composite preform.

[0010] Figure 5 According to another aspect disclosed herein, there are a plurality of composite preforms. Figure 3 Schematic cross-sectional front view of a variation of a cross-section of a composite airfoil.

[0011] Figure 6 According to another aspect disclosed herein, a mid-plane epidermis Figure 5 Schematic cross-sectional front view of a variation of a cross-section of a composite airfoil. DETAILED DESCRIPTION

[0012] Aspects disclosed herein relate to a turbine engine including a composite airfoil assembly, the composite airfoil assembly including a dovetail and an airfoil. The composite airfoil can be used in one or more positions within a turbine engine. For example, the composite airfoil assembly is suitable as a fan blade in a fan section of a turbine engine. Although other positions (such as compressor sections and turbine sections) are also conceivable. The composite airfoil assembly can be installed in a variety of ways. One such installation is to fix the blade to the rotator of the fan section directly or via a pitch control assembly. Regardless of where the composite 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 composite airfoil assembly and the disk assembly can form a rotating assembly together, so that the composite airfoil assembly is a composite blade assembly.

[0013] The composite airfoil assembly also includes a composite core and a set of skins, the set of skins including at least an outer skin and a fan-shaped skin. The outer skin covers the composite core, and the fan-shaped skin covers at least a portion of the outer skin. The composite airfoil assembly is designed to withstand the high mechanical stress associated with the operation of the turbine engine. For the purpose of illustration, the present disclosure will be described with respect to a composite airfoil assembly for a turbine engine, specifically, with respect to a composite 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 composite 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 composite blade assemblies including dovetails. It should be understood that aspects of the present disclosure are directed to any composite airfoil assembly having a dovetail, including but not limited to 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) or same qualified resin transfer molding (SQ-RTM) can be used to form at least a portion of the composite component. Typically, RTM includes applying dry fibers or matrix materials to a mold or cavity. The dry fibers or matrix materials may include prepregs, braided materials, woven materials, or any combination thereof. The dry fibers may define an interpretation. The matrix material may define a precured preform.

[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] As used herein, a "preform" is a three-dimensional composite material formed by a plurality of yarns including warp yarns and weft yarns. In addition, at least one binder yarn partially or completely extends through the preform in a direction perpendicular to the warp yarns and the weft yarns. At least one binder yarn can be integrated into the preform by weaving, stitching, tufting, or any suitable manufacturing process. A preform as described herein is a fully formed part that is formed in a separate manufacturing process relative to the fully formed part.

[0032] 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.

[0033] 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.

[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, the term "modulus" refers to the resistance of an object or material to deformation when stress is applied.

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

[0037] Figure 1 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 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 airflow exiting the remainder of the fan section 18 (bypass airflow 78) bypasses the LP compressor 24 and the engine core 44 and exits the turbine engine 10 through a stationary blade 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 is used adjacent the fan section 18 to exert some directional control on the bypass 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 a composite airfoil assembly 104 and a disk assembly 102 for use within 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, a disk within the fan section 18, the compressor section 22, or the turbine section 32 of the turbine engine 10. The composite airfoil assembly 104 may be rotating or non-rotating, such that the composite 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 composite 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 a first direction or axial direction (denoted as "A"). A radial direction (denoted as "R") extends radially outward and is perpendicular to the axial direction A. In addition, a circumferential direction (denoted as "C") is shown that extends generally in a circumferential direction relative to the rotational axis 106 and is perpendicular to the axial direction A and the radial direction R.

[0051] The composite airfoil assembly 104 includes an airfoil 110 and a dovetail 112 extending from the airfoil 110. The airfoil 110 extends between a leading edge 114 and a trailing edge 116 to define a chordwise 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 that defines 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 the 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 composite 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 composite 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 composite 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 composite 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 2 Schematic cross-sectional front view of a portion of the composite airfoil assembly 104 as viewed along section line III-III. The composite airfoil assembly 104 extends from the second end 128 of the dovetail 112 to the tip 120 of the airfoil 110 ( Figure 2) extends from a centerline axis 132. The centerline axis 132 is equidistant from respective opposing portions of the composite airfoil assembly 104. The centerline axis 132 may be linear or non-linear. The composite airfoil assembly 104 may be symmetrical or asymmetrical about the centerline axis 132. The centerline axis 132 may also be a body axis, which may not extend along the center of the body in the case of an asymmetrical airfoil.

[0054] The composite 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 ). As non-limiting examples, the composite core 130 described herein may be included as a braided fabric, a woven fabric, a composite preform, or any combination thereof.

[0055] A set of skins 140 covers the composite core 130. The set of skins may be a multi-layer skin 148 including an outer skin 142 and a scalloped skin 146. The scalloped skin 146 covers at least a portion of the outer skin 142. The scalloped skin 146 may cover the outer skin 142 near the root 118 to define the first end 126 of the dovetail 112. The scalloped skin 146 may cover the outer skin 142 radially above the first end 126 by less than or equal to the span length SL ( Figure 2 ). The scalloped skin 146 may further define a dovetail outer surface 152 of the dovetail 112. As non-limiting examples, the skins described herein may be defined as laminated skins, woven skins, or braided skins, or any combination thereof. The set of skins 140 may be any suitable type of skin, and include at least one woven skin or at least one braided skin. Additionally, the set of skins 140 may include at least one of a filler ply or a machine ply.

[0056] The composite core 130 and the set of skins 140 may define various portions of the airfoil 110 and the dovetail 112. The outer skin 142 may define at least a portion of the airfoil exterior surface 144. The multi-layer skin 148 may include a skin layer between the composite core 130 and the outer skin 142. The multi-layer skin 148 may extend into at least a portion of the airfoil 110 and the dovetail 112. The multi-layer skin 148 may be delaminated in direct contact with adjacent skins 148a, 148b in the airfoil 110.

[0057] At least one insert 134 forms a portion of the dovetail 112 having a flared cross-section 150. The at least one insert 134 is positioned between adjacent skins 148a, 148b to define the flared cross-section 150, wherein the multiple layers of skins 148 fan outwardly away from each other and away from the centerline axis 132. The at least one insert 134 may have a defined skin ply angle θ s 148b. The at least one insert 134 may be formed to provide a triangular cross-sectional shape to provide an increased amount of spacing between adjacent skins 148a, 148b. In other words, the multi-layer skin 148 may transition from being in direct contact with adjacent skins 148a, 148b at the first end 126 to being gradually spaced an increasing amount between the same adjacent skins 148a, 148b at the second end 128. The at least one insert 134 is formed to provide a predefined spacing (denoted as "S") between adjacent skins 148a, 148b. The amount of the predefined spacing S is measured along the circumferential direction C and is maximized at the second end 128. The at least one insert 134 may be a single filler material, a composite material, a composite preform, or any combination of materials.

[0058] At least one insert 134 is a composite preform 154 that defines at least a portion of the flared cross-section 150. The composite preform 154 can be located between any skins in the set of skins 140. In one non-limiting example, the composite preform 154 is located between the outer skin 142 and the scalloped skin 146. The scalloped skin 146 transitions from being in direct contact with the outer skin 142 at the first end 126 to being gradually spaced apart an increasing amount toward the second end 128. As shown, the scalloped skin 146 scallops outwardly away from the centerline axis 132, articulated from the first end 126 to define a dovetail outer surface 152. The two inserts 134 can be a pair of composite preforms 154a, 154b that define two sides of the flared cross-section 150 relative to the centerline axis 132.

[0059] The composite preform 154 is formed to provide the majority of the fan-out for the composite airfoil assembly 104. The composite preform 154 provides a maximum amount of separation (denoted as “S ”) between the outer skin 142 and the scalloped skin 146 measured along a line perpendicular to the centerline axis 132 and at the second end 128. M ”). Maximum spacing S M Less than or equal to 2.5 inches (S M < 2.5in) or less than or equal to 6.4cm(S M < 6.4 cm). This reduces the skin ply angle θ when transitioning from the airfoil 110 to the dovetail 112 sIn other words, the predefined spacing S is reduced by adding the composite preform 154. A line 153 perpendicular to the axial direction A and extending tangentially from the dovetail outer surface 152 is at a sector angle θ with respect to the centerline axis 132. f The fan angle θ is defined as the amount of fanning outward and away from the center line 132. f It can be between 0 and 45° (0<θ f < 45°). It is further contemplated that the outer skin 142 and the sector skin 146 are integrally formed, or are a continuous piece, and that the outer skin 142 extends from the first end 126 at a sector angle θ f Bend outwards.

[0060] Go to Figure 4 , showing Figure 3 154. Any of the at least one insert 134 disclosed herein may embody the same features as the illustrated single composite preform 154. The composite preform 154 may have a triangular cross-sectional shape 160 with a vertex 162 having a relative fan angle θ f The congruent angles θ c . The composite preform 154 may include at least one material 156 oriented in three different directions, which are generally referred to as warp orientation, weft orientation, and binder orientation when the composite preform 154 is formed using a weaving process. Although described as a weaving process, it should be understood that any process that results in the orientation described herein is contemplated. The at least one material 156 can be a single type of material oriented in three different directions, or multiple types of materials oriented in three different directions. Specifically, the composite preform 154 is formed from at least one material 156, at least one material 156 having a binder orientation in the axial direction A and having the lowest modulus (denoted as "E3") of the composite preform 154. The at least one material 156 is further oriented in the warp direction along a direction parallel to the dovetail outer surface 152 ( Figure 3 ) and has a highest modulus (denoted as "E1"). At least one material 156 is further woven in a weft orientation along a third direction (denoted as "T") perpendicular to the second direction P and parallel to the thickness (denoted as "Th") of the composite preform 154, and is associated with an intermediate modulus (denoted as "E2"). In other words, the highest modulus E1 is greater than the intermediate modulus E2, and the intermediate modulus E2 is greater than the lowest modulus E3 (E1>E2>E3).

[0061] From the dovetail outer surface 152 ( Figure 3 ) toward the centerline axis 132 ( Figure 3) moves, the second direction P becomes closer and closer to being parallel to the radial direction R. Similarly, the third direction T becomes closer and closer to being parallel to the circumferential direction C ( Figure 3 In other words, when at least one insert 134 has the same properties as the composite preform 154 and is positioned closer to the centerline axis 132, the warp orientation is close to or parallel to the radial direction R. Likewise, the weft orientation is close to or parallel to the circumferential direction C.

[0062] Go to Figure 5 , a composite airfoil assembly 204 is shown according to another aspect disclosed herein. Composite airfoil assembly 204 is similar to composite airfoil assembly 104, and thus, like parts will be identified with like numerals increased by 100, it being understood that the description of like parts of composite airfoil assembly 104 applies to composite airfoil assembly 204 unless otherwise noted.

[0063] The composite core 230 and a set of skins 240 may define portions of the airfoil 210 and the dovetail 212. The dovetail 212 may extend between a first end 226 and a second end 228 in a radial direction R. The first end 226 represents a transition between the dovetail 212 and the airfoil 210. As a non-limiting example, the first end 226 coincides with the root 218 of the airfoil 110. The dovetail 212 and the airfoil 210 may be integrally or non-integrally formed with one another. The set of skins 240 may be a multi-layer skin 248 including an outer skin 242 and a scalloped skin 246. In the dovetail 212, the multi-layer skins 248 splay outwardly away from one another and away from the centerline axis 232. In other words, the multi-layer skin 248 can transition from being in direct contact with adjacent skins 248a, 248b at the first end 226 of the dovetail 212 to being gradually spaced apart by increasing amounts between the same adjacent skins 248a, 248b at the second end 228. The scalloped skin 246 covers at least a portion of the outer skin 242. The outer skin 242 can define at least a portion of the airfoil exterior surface 244. A line 253 that is perpendicular to the axial direction A and extends tangentially from the dovetail exterior surface 252 is at a scalloped angle θ with respect to the centerline axis 232. f intersect, thereby defining the amount of fanning outward and away from the centerline 232. It is further contemplated that the outer skin 242 and the scalloped skin 246 are integrally formed, or formed as one continuous piece, wherein the outer skin 242 is fanned at an angle θ f Curved outward from the first end 226 .

[0064] The plurality of inserts 234 form a portion of the dovetail 212 having a flared cross-section 250. As shown, the plurality of inserts 234 are positioned between adjacent skins 248a, 248b to define the flared cross-section 250 and provide a predefined spacing (indicated as "S"). Each of the plurality of inserts 234 may have a triangular cross-sectional shape 260 that defines a skin layup angle θ. s To provide a predefined spacing S between adjacent skins 248a, 248b. Skin ply angle θ s Can be smaller than the sector angle θ f (θ s <θ f ). Skin ply angle θ s Can be varied between multiple inserts 234. Skin ply angle θ s Can be 0 degrees, 45 degrees, or between 0 and 45 degrees (0° < θ s < 45°).

[0065] The plurality of inserts 234 may each be a composite preform 254 formed of materials oriented in three different directions. Specifically, the composite preform 254 formed of preform materials is similar to the composite preform 154 previously described herein. The binder orientation is associated with the lowest modulus E3, the warp orientation is associated with the highest modulus E1, and the weft orientation is associated with the intermediate modulus E2, wherein the highest modulus E1 is greater than the intermediate modulus E2, which in turn is greater than the lowest modulus E3 (E1>E2>E3). Moving from the dovetail outer surface 252 toward the centerline axis 232, the warp orientation extends along the second direction P, which becomes increasingly closer to being parallel to the radial direction R for the composite preform 254 closer to the centerline. Similarly, the weft orientation extends along the third direction T, which becomes increasingly closer to being parallel to the circumferential direction C. In other words, the composite preform 254 closest to the centerline axis 232 has the highest modulus E1 oriented close to or parallel to the radial direction R. Likewise, the composite preform 254 closest to the centerline axis 232 has an intermediate modulus E2 oriented close to or parallel to the circumferential direction C. The composite preform 254 is formed to have a skin ply angle θ s The plurality of composite preforms 254 enables the same transition from the airfoil 110 to the dovetail 112 discussed previously herein while also reducing layup time.

[0066] At least one composite preform 254c provides a maximum spacing amount (denoted as “S”) between adjacent skins of the multi-layer skin 248 measured along a line perpendicular to the centerline axis 232. M ”). Maximum spacing S M Less than or equal to 2.5 inches (SM < 2.5in) or less than or equal to 6.4cm(S M < 6.4cm).

[0067] Go to Figure 6 , showing a composite airfoil assembly 304 according to another aspect of the disclosure. The composite airfoil assembly 304 is similar to Figure 5 Composite airfoil assembly 204. Composite airfoil assembly 304 extends along a centerline axis 332. Centerline axis 332 is equidistant from respective opposing portions of composite airfoil assembly 304. Centerline axis 332 may be linear or non-linear. Composite airfoil assembly 304 may be symmetrical or asymmetrical about centerline axis 332. Centerline axis 332 may also be a body axis, which may not extend along the center of the body in the case of an asymmetric airfoil. Composite core 330 may extend radially through composite airfoil assembly 304 along centerline axis 332. A set of skins 340 may include a midplane skin 338 located within composite core 330. Midplane skin 338 may define centerline axis 332. Midplane skin 338 may include material oriented in the warp orientation, weft orientation, and binder orientation previously described herein. Although shown as a composite airfoil assembly 204 ( Figure 5 ), but it should be understood that the midplane skin 338 may also be provided on the composite airfoil assembly 104 ( Figure 3 )middle.

[0068] Benefits associated with using a composite airfoil assembly include a lighter airfoil assembly, as compared to a non-composite (e.g., cast) airfoil assembly, without sacrificing performance of the airfoil assembly. In other words, the material used for the composite airfoil assembly is lighter than the material used for the non-composite airfoil assembly, without sacrificing the ability to perform as intended within the turbine engine. The reduced weight, in turn, means an increase in the efficiency of the turbine engine when compared to a conventional turbine engine including a non-composite airfoil assembly.

[0069] Further benefits associated with the composite airfoil assembly include the ability of the airfoil assembly to withstand mechanical stresses associated with the operation of the composite airfoil assembly as compared to conventional composite airfoil assemblies. Specifically, the highest modulus is generally in the direction of the highest transferred load between the airfoil and the dovetail, i.e., in the spanwise direction of the airfoil. Likewise, the intermediate or second highest modulus is generally in the direction of the second highest transferred load between the airfoil and the dovetail, i.e., in the chordwise direction of the airfoil. Purposefully forming a composite preform having these orientations contributes to a stronger composite airfoil assembly.

[0070] 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.

[0071] 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.

[0072] Further aspects are provided by the subject matter of the following clauses:

[0073] A composite airfoil assembly for a turbine engine, the composite airfoil assembly comprising: an airfoil extending in a radial direction 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, the set of skins comprising an outer skin defining at least a portion of an outer surface of the airfoil; and a dovetail extending radially between a first end and a second end below the root, the dovetail comprising the composite core, at least one insert comprising a composite preform, and the set of skins, the set of skins comprising a fan-shaped skin covering the at least one insert; wherein the composite preform defines at least a portion of a flared cross-section of the dovetail.

[0074] A composite airfoil assembly as described in any preceding clause, wherein said at least one insert comprises a triangular cross-sectional shape defining a sector angle.

[0075] A composite airfoil assembly as claimed in any preceding clause, wherein the scalloped skin extends away from the outer skin at the scalloped angle.

[0076] A composite airfoil assembly as described in any preceding clause, wherein the sector angle is less than or equal to 45 degrees.

[0077] A composite airfoil assembly as described in any preceding clause, wherein the set of skins comprises a plurality of skins.

[0078] A composite airfoil assembly as described in any preceding clause, wherein the at least one insert is a plurality of inserts.

[0079] A composite airfoil assembly as described in any preceding clause, wherein the plurality of inserts separate adjacent skins from one another.

[0080] A composite airfoil assembly as in any preceding clause, wherein the root defines 0% of the span length, the tip defines 100% of the span length, and the scalloped skin extends radially over the root an amount less than or equal to 5% of the span length.

[0081] A composite airfoil assembly as described in any preceding clause, wherein the composite core comprises a braided fabric, a woven fabric or a composite preform.

[0082] A composite airfoil assembly according to any preceding clause, wherein the set of skins comprises at least one laminate skin or at least one woven skin.

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

[0084] A composite airfoil assembly according to any preceding clause, wherein the composite preform is formed from at least one material oriented in a first direction associated with a lowest modulus, a second direction associated with a highest modulus, and a third direction associated with an intermediate modulus.

[0085] A composite airfoil assembly as in any preceding clause, wherein the first direction is oriented along an axial direction of the dovetail, the third direction is oriented along a thickness of the dovetail, and the second direction is perpendicular to the first and third directions.

[0086] A composite airfoil assembly according to any preceding clause, wherein the set of skins are prepregs.

[0087] A composite airfoil assembly according to any preceding clause, wherein the composite preform is a preform or a precured preform and is combined with the prepreg by resin transfer molding (RTM) or same quality resin transfer molding (SQ-RTM).

[0088] A composite airfoil assembly according to any preceding clause, wherein at least a portion of 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 polymeric resin, a thermoplastic, a bismaleimide, a polyimide, an epoxy resin, a glass fiber, or a silicon matrix.

[0089] A composite airfoil assembly according to any preceding clause, wherein the airfoil is a blade and the turbine engine comprises a fan section, a compressor section, a combustion section and a turbine section in a serial flow arrangement, wherein the composite airfoil assembly is disposed within the fan section.

[0090] A composite airfoil assembly as claimed in any preceding clause, wherein the scalloped skin defines a further portion of the flared cross-section.

[0091] A composite airfoil assembly as claimed in any preceding clause, wherein the scalloped skin overlies the outer skin at the first end.

[0092] A composite airfoil assembly as claimed in any preceding clause, wherein the scalloped skin extends away from the outer skin at a scalloped angle.

[0093] A composite airfoil assembly as described in any preceding clause, wherein the sector angle is less than or equal to 30 degrees.

[0094] A composite airfoil assembly as described in any preceding clause, wherein the at least one insert is a plurality of inserts, each insert comprising a composite preform.

[0095] A composite airfoil assembly as described in any preceding clause, wherein said plurality of inserts together with said set of skins define said flared cross-section of said dovetail.

[0096] A composite airfoil assembly for a turbine engine, the composite airfoil assembly comprising: an airfoil extending in a radial direction 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, the set of skins comprising an outer skin defining at least a portion of an outer surface of the airfoil; and a dovetail extending radially below the root and having a flared cross-section, the dovetail comprising the composite core, at least one insert comprising a composite preform, and the set of skins, the set of skins comprising a fan-shaped skin covering the at least one insert and extending away from the outer skin at a fan angle; wherein the fan angle is less than or equal to 30 degrees.

[0097] A composite airfoil assembly as described in any preceding clause, wherein the set of skins comprises a plurality of skins.

[0098] A composite airfoil assembly as described in any preceding clause, wherein the at least one insert is a plurality of inserts, and wherein the plurality of inserts separate adjacent skins from one another.

[0099] A composite airfoil assembly as claimed in any preceding clause, wherein the scalloped skin overlies the outer skin at the first end.

[0100] A composite airfoil assembly as claimed in any preceding clause, wherein the composite preform defines at least a portion of a flared cross-section of the dovetail and the scalloped skin defines another portion of the flared cross-section.

Claims

1. A composite airfoil assembly for a turbine engine, characterized in that: The composite airfoil assembly comprises: an airfoil extending in a radial direction 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, the set of skins including an outer skin defining at least a portion of an exterior surface of the airfoil; and a dovetail extending radially below the root between a first end and a second end, the dovetail comprising the composite core, at least one insert comprising a composite preform, and the set of skins comprising a scalloped skin covering the at least one insert; wherein the composite preform defines at least a portion of a flared cross-section of the dovetail.

2. The composite airfoil assembly according to claim 1, wherein: in, The at least one insert includes a triangular cross-sectional shape defining a sector angle.

3. The composite airfoil assembly according to claim 2, wherein: in, The sector skin extends away from the outer skin at the sector angle.

4. The composite airfoil assembly according to claim 3, wherein: in, The sector angle is less than or equal to 45 degrees.

5. The composite airfoil assembly of claim 1, wherein: in, The set of epidermis includes multiple layers of epidermis.

6. The composite airfoil assembly of claim 1, wherein: in, The at least one insert is a plurality of inserts.

7. The composite airfoil assembly according to claim 6, wherein: in, The plurality of inserts separate adjacent skins from one another.

8. The composite airfoil assembly of claim 1, wherein: in, The root defines 0% of the span length, the tip defines 100% of the span length, and the scalloped skin extends radially over the root an amount less than or equal to 5% of the span length.

9. The composite airfoil assembly of claim 1, wherein: in, The composite core includes a braided fabric, a woven fabric or a composite preform.

10. The composite airfoil assembly of claim 1, wherein: in, The set of skins includes at least one laminated skin or at least one woven skin.

Images