TURBINE ENGINE AIRFOIL WITH A WOVEN CORE

A gas turbine engine includes a fan section, a compressor section, combustor section, and turbine section in serial flow arrangement, and defining an engine longitudinal axis. Sets of blades can be rotatably driven about the longitudinal axis in the fan section, the compressor section, and the turbine section. The airfoil structure defining one or more of the blades or vanes can include a structure which includes a woven or foam core, with a laminate skin, and an exterior coating.

TECHNICAL FIELD

The present disclosure relates generally to a core for a component of a gas turbine, and more specifically to an airfoil having a woven core geometry.

BACKGROUND

A turbine engine typically includes an engine core with a compressor section, a combustor section, and a turbine section in serial flow arrangement. A fan section can be provided upstream of the compressor section. The compressor section compresses air which is channeled to the combustor section where it is mixed with fuel, where the mixture is then ignited for generating hot combustion gases. The combustion gases are channeled to the turbine section which extracts energy from the combustion gases for powering the compressor section, as well as for producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator.

Forming engine components is commonly achieved by casting. Casting is a common manufacturing technique for forming various components of a gas turbine aviation engine. Casting a component involves a mold having a void in the form of a negative of the desired component shape, filling the void with a flowable material, letting the material harden, and removing the mold.

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 moldability to adopt relatively complex shapes, composite materials are utilized in various applications, such as a turbine engine or an aircraft. Composite materials can be, for example, installed on or define a portion of the fuselage and/or wings, rudder, manifold, airfoil, or other components of the aircraft or turbine engine. Extreme loading or sudden forces can be applied to the composite components of the aircraft or turbine engine. For example, extreme loading can occur to one or more airfoils during ingestion of various materials by the turbine engine.

DETAILED DESCRIPTION

Aspects of the disclosure herein are directed to a manufactured core used for an engine component, such as an airfoil. The core includes a woven core, and can include additional woven layers forming the engine component. The woven core is used to create an engine component for a turbine engine. Such an engine component can be an airfoil, for example. It should be understood, however, that the disclosure applies to other engine components of the turbine engine, not just an airfoil, such as a disk or combustor liner, in non-limiting examples. Further, while described in terms of a core used in the manufacture of an airfoil, it will be appreciated that the present disclosure is applied to any other suitable environment.

The terms “forward” and “aft” refer to relative positions within a turbine engine or vehicle, and refer to the normal operational attitude of the turbine engine or vehicle. For example, with regard to a turbine engine, forward refers to a position closer to an engine inlet and aft refers to a position closer to an engine nozzle or exhaust.

As used herein, the term “upstream” refers to a direction that is opposite the fluid flow direction, and the term “downstream” refers to a direction that is in the same direction as the fluid flow. The terms “fore” or “forward” mean in front of something and “aft” or “rearward” mean behind something. For example, when used in terms of fluid flow, fore/forward can mean upstream and aft/rearward can mean downstream.

The term “fluid” may be a gas or a liquid, or multi-phase.

Additionally, as used herein, the terms “radial” or “radially” refer 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 center longitudinal axis of the engine and an outer engine circumference.

The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise. Furthermore, as used herein, the term “set” or a “set” of elements can be any number of elements, including only one.

As used herein, the term “exterior” or “exteriorly” refers to an exterior position or location relative to a common interior or center. For example, as used herein, “located exteriorly” can refer to a location further from a core, relative to another aspect.

As used herein, the term “stiffness” may be used as defining the extent to which a structure resists deformation in response to force. Stiffness can be defined as the ratio of force to displacement of the object under said force. Stiffness can include resisting deformation in response to force applied from various directionalities, whereby the stiffness can represent an axial stiffness, tensile stiffness, compression stiffness, torsional stiffness, or shear stiffness in non-limiting examples.

As used herein, the term “elasticity” may be used as defining the modulus of elasticity (Young's modulus) under tension or compression, can may relate to an elasticity for a particular material or structure made of such material, such as the engine components described herein. The elasticity can represent the stress per unit area relative to the local strain or proportional deformation thereof.

The term “composite,” as used herein is, is indicative of a component having two or more materials A composite can be a combination of at least two or more metallic, non-metallic, or a combination of metallic and non-metallic elements or materials. Examples of a composite material can be, but not limited to, a polymer matrix composite (PMC), a ceramic matrix composite (CMC), a metal matrix composite (MMC), carbon fibers, a polymeric resin, a thermoplastic resin, bismaleimide (BMI) materials, polyimide materials, an epoxy resin, glass fibers, and silicon matrix materials.

As used herein, a “composite” component refers to a structure or a component including any suitable composite material. Composite components, such as a composite airfoil, can include several layers or plies of composite material. The layers or plies can vary in stiffness, material, and dimension to achieve the desired composite component or composite portion of a component having a predetermined weight, size, stiffness, and strength.

One or more layers of adhesive can be used in forming or coupling composite components. Adhesives can include resin and phenolics, wherein the adhesive can require curing at elevated temperatures or other hardening techniques.

As used herein, PMC refers to a class of materials. By way of example, the PMC material is defined in part by a prepreg, which is a reinforcement material pre-impregnated with a polymer matrix material, such as thermoplastic resin. Non-limiting examples of processes for producing thermoplastic prepregs include hot melt pre-pregging in which the fiber reinforcement material is drawn through a molten bath of resin and powder pre-pregging in which a resin is deposited onto the fiber reinforcement material, by way of non-limiting example electrostatically, and then adhered to the fiber, by way of non-limiting example, in an oven or with the assistance of heated rollers. The prepregs can be in the form of unidirectional tapes or woven fabrics, which are then stacked on top of one another to create the number of stacked plies desired for the part.

Multiple layers of prepreg are stacked to the proper thickness and orientation for the composite component and then the resin is cured and solidified to render a fiber reinforced composite part. Resins for matrix materials of PMCs can be generally classified as thermosets or thermoplastics. Thermoplastic resins are generally categorized as polymers that can be repeatedly softened and flowed when heated and hardened when sufficiently cooled due to physical rather than chemical changes. Notable example classes of thermoplastic resins include nylons, thermoplastic polyesters, polyaryletherketones, and polycarbonate resins. Specific example of high performance thermoplastic resins that have been contemplated for use in aerospace applications include, polyetheretherketone (PEEK), polyetherketoneketone (PEKK), polyetherimide (PEI), polyaryletherketone (PAEK), and polyphenylene sulfide (PPS). In contrast, once fully cured into a hard rigid solid, thermoset resins do not undergo significant softening when heated, but instead thermally decompose when sufficiently heated. Notable examples of thermoset resins include epoxy, bismaleimide (BMI), and polyimide resins.

Instead of using a prepreg, in another non-limiting example, with the use of thermoplastic polymers, it is possible to utilize a woven fabric. Woven fabric can include, but is not limited to, dry carbon fibers woven together with thermoplastic polymer fibers or filaments. Non-prepreg braided architectures can be made in a similar fashion. With this approach, it is possible to tailor the fiber volume of the part by dictating the relative concentrations of the thermoplastic fibers and reinforcement fibers that have been woven or braided together. Additionally, different types of reinforcement fibers can be braided or woven together in various concentrations to tailor the properties of the part. For example, glass fibers, carbon fibers, and thermoplastic fibers could all be woven together in various concentrations to tailor the properties of the part. The carbon fibers provide the strength of the system, the glass fibers can be incorporated to enhance the impact properties, which is a design characteristic for parts located near the inlet of the engine, and the thermoplastic fibers provide the binding for the reinforcement fibers.

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

Resin can be pumped into 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 are then cured and removed from the mold. When removed from the mold, the composite component can require post-curing processing.

It is contemplated that RTM can be a vacuum assisted process. That is, the air from the cavity or mold can be removed and replaced by the resin prior to heating or curing. It is further contemplated that the placement of the dry fibers or matrix material can be manual or automated.

The dry fibers or matrix material can be contoured to shape the composite component or direct the resin. Optionally, additional layers or reinforcing layers of a material differing from the dry fiber or matrix material can also be included or added prior to heating or curing.

As used herein, CMC refers to a class of materials with reinforcing fibers in a ceramic matrix. Generally, the reinforcing fibers provide structural integrity to the ceramic matrix. Some examples of reinforcing fibers can 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 oxycarbides, silicon oxynitrides, aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates such as mullite, or mixtures thereof), or mixtures thereof.

Some examples of ceramic matrix materials can 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 oxycarbides, silicon oxynitrides, 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) can also be included within the ceramic matrix.

Generally, particular CMCs can be referred to as their combination of type of fiber/type of matrix. For example, C/SiC for carbon-fiber-reinforced silicon carbide; SiC/SiC for silicon carbide-fiber-reinforced silicon carbide, SiC/SiN for silicon carbide fiber-reinforced silicon nitride; SiC/SiC—SiN for silicon carbide fiber-reinforced silicon carbide/silicon nitride matrix mixture, etc. In other examples, the CMCs can be comprised 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.

In certain non-limiting examples, the reinforcing fibers may be bundled and/or coated prior to inclusion within the ceramic matrix. For example, bundles of the fibers may be formed as a reinforced tape, such as a unidirectional reinforced tape. A plurality of the tapes may be laid up together to form a preform component. The bundles of fibers may be impregnated with a slurry composition prior to forming the preform or after formation of the preform. The preform may then undergo thermal processing, and subsequent chemical processing to arrive at a component formed of a CMC material having a desired chemical composition. For example, the preform may undergo a cure or burn-out to yield a high char residue in the preform, and subsequent melt-infiltration with silicon, or a cure or pyrolysis to yield a silicon carbide matrix in the preform, and subsequent chemical vapor infiltration with silicon carbide. Additional steps may be taken to improve densification of the preform, either before or after chemical vapor infiltration, by injecting it with a liquid resin or polymer followed by a thermal processing step to fill the voids with silicon carbide. CMC material as used herein may be formed using any known or hereinafter developed methods including but not limited to melt infiltration, chemical vapor infiltration, polymer impregnation pyrolysis (PIP), or any combination thereof.

Such materials, along with certain monolithic ceramics (i.e., ceramic materials without a reinforcing material), are particularly suitable for higher temperature applications. Additionally, these ceramic materials are lightweight compared to superalloys, yet can still provide strength and durability to the component made therefrom. Therefore, such materials are currently being considered for many gas turbine components used in higher temperature sections of gas turbine engines, such as airfoils (e.g., turbines, and vanes), combustors, shrouds and other like components, that would benefit from the lighter-weight and higher temperature capability these materials can offer.

The term “metallic” as used herein is indicative of a material that includes metal such as, but not limited to, titanium, iron, aluminum, stainless steel, and nickel alloys. A metallic material or alloy can be a combination of at least two or more elements or materials, where at least one is a metal.

The inventors' practice has proceeded in the foregoing manner of designing a core used in the manufacture of a component such as an airfoil, designing the airfoil to have improved stiffness transition between the core and an exterior skin, decreased weight, identifying whether or not the component was manufactured as designed and satisfies component objectives, and modifying the engine component with new geometric characteristics of the cast engine. This process is repeated during the design of several different types of components, such as those shown inFIG.1.

FIG.1is a schematic cross-sectional diagram of a turbine engine10for an aircraft. The turbine engine10has a generally longitudinally extending axis or engine centerline12extending forward14to aft16. The turbine engine10includes, in downstream serial flow relationship, a fan section18including a fan20, a compressor section22including a booster or low pressure (LP) compressor24and a high pressure (HP) compressor26, a combustion section28including a combustor30, a turbine section32including a HP turbine34, and a LP turbine36, and an exhaust section38.

The fan section18includes a fan casing40surrounding the fan20. The fan20includes a plurality of fan blades42disposed radially about the engine centerline12. The HP compressor26, the combustor30, and the HP turbine34form an engine core44of the turbine engine10, which generates combustion gases. The engine core44is surrounded by a core casing46, which can be coupled with the fan casing40.

An HP shaft or spool48disposed coaxially about the engine centerline12of the turbine engine10drivingly connects the HP turbine34to the HP compressor26. An LP shaft or spool50, which is disposed coaxially about the engine centerline12of the turbine engine10within the greater diameter annular HP spool48, drivingly connects the LP turbine36to the LP compressor24and fan20. The spools48,50are rotatable about the engine centerline12and couple to a plurality of rotatable elements, which can collectively define a rotor51.

The LP compressor24and the HP compressor26respectively include a plurality of compressor stages52,54, in which a set of compressor blades56,58rotate relative to a corresponding set of static compressor vanes60,62to compress or pressurize the stream of fluid passing through the stage. In a single compressor stage52,54, multiple compressor blades56,58can be provided in a ring and can extend radially outwardly relative to the engine centerline12, from a blade platform to a blade tip, while the corresponding static compressor vanes60,62are positioned upstream of and adjacent to the rotating compressor blades56,58. It is noted that the number of blades, vanes, and compressor stages shown inFIG.1were selected for illustrative purposes only, and that other numbers are possible.

The compressor blades56,58for a stage of the compressor can be mounted to (or integral to) a disk61, which is mounted to the corresponding one of the HP and LP spools48,50. The static compressor vanes60,62for a stage of the compressor can be mounted to the core casing46in a circumferential arrangement.

The HP turbine34and the LP turbine36respectively include a plurality of turbine stages64,66, in which a set of turbine blades68,70are rotated relative to a corresponding set of static turbine vanes72,74, also referred to as a nozzle, to extract energy from the stream of fluid passing through the stage. In a single turbine stage64,66, multiple turbine blades68,70can be provided in a ring and can extend radially outwardly relative to the engine centerline12while the corresponding static turbine vanes72,74are positioned upstream of and adjacent to the rotating turbine blades68,70. It is noted that the number of blades, vanes, and turbine stages shown inFIG.1were selected for illustrative purposes only, and that other numbers are possible.

The turbine blades68,70for a stage of the turbine can be mounted to a disk71, which is mounted to the corresponding one of the HP and LP spools48,50. The static turbine vanes72,74for a stage of the compressor can be mounted to the core casing46in a circumferential arrangement.

Complementary to the rotor portion, the stationary portions of the turbine engine10, such as the static vanes60,62,72,74among the compressor and turbine sections22,32are also referred to individually or collectively as a stator63. As such, the stator63can refer to the combination of non-rotating elements throughout the turbine engine10.

In operation, the airflow exiting the fan section18is split such that a portion of the airflow is channeled into the LP compressor24, which then supplies a pressurized airflow76to the HP compressor26, which further pressurizes the air. The pressurized airflow76from the HP compressor26is mixed with fuel in the combustor30and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine34, which drives the HP compressor26. The combustion gases are discharged into the LP turbine36, which extracts additional work to drive the LP compressor24, and an exhaust gas is ultimately discharged from the turbine engine10via the exhaust section38. The driving of the LP turbine36drives the LP spool50to rotate the fan20and the LP compressor24.

A portion of the pressurized airflow76can be drawn from the compressor section22as bleed air77. The bleed air77can be drawn from the pressurized airflow76and provided to engine components requiring cooling. The temperature of pressurized airflow76entering the combustor30is significantly increased above the bleed air temperature. The bleed air77may be used to reduce the temperature of the core components downstream of the combustor30.

A remaining portion of the airflow bypasses the LP compressor24and engine core44as a bypass airflow78, and exits the turbine engine10through a stationary vane row, and more particularly an outlet guide vane assembly80, comprising a plurality of airfoil guide vanes82, at a fan exhaust side84. More specifically, a circumferential row of radially extending airfoil guide vanes82are utilized adjacent the fan section18to exert some directional control of the airflow78.

Some of the air supplied by the fan20can bypass the engine core44and be used for cooling of portions, especially hot portions, of the turbine engine10, and/or used to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are normally downstream of the combustor30, especially the turbine section32, with the HP turbine34being the hottest portion as it is directly downstream of the combustion section28. Other sources of cooling fluid can be, but are not limited to, fluid discharged from the LP compressor24or the HP compressor26.

FIG.2is a schematic perspective view of a composite airfoil assembly104and a disk assembly102suitable for use within the turbine engine10ofFIG.1. The disk assembly102is suitable for use as the disk61,71(FIG.1) or any other disk such as, but not limited to, a disk within the fan section18, the compressor section22, or the turbine section32of the turbine engine10. The composite airfoil assembly104can be rotating or non-rotating such that the composite airfoil assembly104can include at least one of the static compressor vanes60,62(FIG.1), the set of compressor blades56,58(FIG.1), the static turbine vanes72,74(FIG.1), the set of turbine blades68,70(FIG.1), or the plurality of fan blades42(FIG.1). As a non-limiting example, the composite airfoil assembly104can be a composite fan blade assembly.

The disk assembly102can be rotatable or stationary about a rotational axis106. The rotational axis106can coincide with or be offset from the engine centerline (e.g., the engine centerline12ofFIG.1). The disk assembly102includes a plurality of slots108extending axially through a radially outer portion of the disk assembly102and being circumferentially spaced about the disk assembly102, with respect to the rotational axis106.

The composite airfoil assembly104includes an airfoil portion110and a dovetail portion112extending from the airfoil portion110. The airfoil portion110extends between a leading edge114and a trailing edge116to define a chord-wise direction. The airfoil portion110extends between a root118and a tip120to define a span-wise direction. The airfoil portion110includes a pressure side122and a suction side124. The dovetail portion112extends between a first end126and a second end128in the span-wise direction. The first end126is radially spaced inwardly from the second end128, with respect to the rotational axis106. The first end126defines a radial inner surface of the dovetail portion112. The second end128denotes a transition between the dovetail portion112and the airfoil portion110. As a non-limiting example, the second end128coincides with the root118of the airfoil portion110. The dovetail portion112and the airfoil portion110can be integrally or non-integrally formed with each other.

The composite airfoil assembly104is coupled to the disk assembly102by inserting at least a portion of the dovetail portion112into a respective slot of the plurality of slots108. The composite airfoil assembly104is held in place by frictional contact with the slot108or can be coupled to the slot108via any suitable coupling method such as, but not limited to, welding, adhesion, fastening, or the like. While only a single composite airfoil assembly104is illustrated, it will be appreciated that there can be any number of composite airfoils assemblies104coupled to the disk assembly102. As a non-limiting example, there can be a plurality of composite airfoil assemblies104corresponding to a total number of slots of the plurality of slots108.

For the sake of reference, a set of relative reference directions, along with a coordinate system can be applied to the composite airfoil assembly104. An axial direction (Ad), can extend from forward to aft and is shown extending at least partially into the page. The axial direction (Ad) and can be arranged parallel to the rotational axis106. A radial direction (Rd) extends perpendicular to the axial direction (Ad), and can extend perpendicular to the engine centerline12. A circumferential direction (Cd) can be defined perpendicular to the radial direction (Rd), and can be defined along the circumference of the turbine engine10(FIG.1) relative to the engine centerline12(FIG.1).

FIG.3shows a cross-sectional view of the composite airfoil assembly104ofFIG.2, taken along section III-III, illustrating an interior140of the composite airfoil assembly104. The composite airfoil assembly104includes a woven core130, a laminate skin132provided over the woven core130, and a coating134provided over the laminate skin132. The woven core130can include a first stiffness and a first elasticity.

The woven core130can be dry, with no additional materials, or alternatively, be impregnated with a resin and cured in one non-limiting example. The woven core130can made of a woven structure. Such a woven structure can be a three-dimensional woven structure. More specifically, the woven structure can be woven in a combination of the axial direction Ad, the radial direction Rd, and the circumferential direction Cd (FIG.2), while it should be appreciated that the weave pattern can be formed and defined separate from the turbine engine10(FIG.1), such that the weave pattern is woven in any three, mutually-orthogonal planes in order to define a three-dimensional object relative to said planes. In one non-limiting example, the woven structure can include a three-dimensional weaving including a plurality of warp fibers136and weft fibers138which can be woven in three directions to form a three-dimensional structure for the woven core130. The three directions for the warp fibers136and weft fibers138can be defined along or angled relative to the axial direction Ad, the radial direction Rd, and the circumferential direction Cd. In one non-limiting example, a Jacquard loom, or 3D weaving machine can be used to create complex three-dimensional woven structures, which can include interweaving one or more composites to form the woven core130. The woven core130can be comprised of composite materials, such as carbon or carbon fibers, glass or glass fibers, nylon, rayon, or other aramid fibers, while other materials such as nickel, titanium, or ceramic composites are contemplated in non-limiting examples.

It is further contemplated that the woven core130can be formed as a three-dimensional woven structure, having a braided or a plaited geometry or pattern. A braided or a plaited geometry or pattern can include a weave pattern that includes three or more interlaced fibers that are woven in a repeating pattern, for example. In another non-limiting example, the braided geometry can include a set of fibers or strands that are sequentially laid over one another to define the braided geometry. The woven or braided geometry or pattern can repeat for the entirety of the woven core130, or only a portion thereof. Such additional braided geometries can be similar, where the arrangement of the fibers is the same, but the orientation is different, or where the arrangement of the fibers is different, and the orientation can be similar or dissimilar. The braided geometry or pattern can be formed with a Jacquard loom or 3D weaving machine with composite materials. A three-dimensional braided structure can include a braided pattern that extends in three dimensions, such as a combination of the axial direction Ad, the radial direction Rd, and the circumferential direction Cd.

The laminate skin132can be formed as a set of laminate layers, provided around or about the woven core130. The laminate skin132can be pre-impregnated, fiber placed, or dry fiber laminate layers, in non-limiting examples. Such laminate layers forming the laminate skin132can be formed by resin transfer molding (RTM), partial RTM, same qualified resin transfer molding (SQRTM), or out-of-autoclave in non-limiting examples. The laminate skin132can include a second stiffness and a second elasticity. In one example, the second stiffness and the second elasticity can be greater than that of the first stiffness or the first elasticity of the woven core130.

The coating134can be applied directly onto the laminate skin132, while it is contemplated that an intermediate adhesive layer is provided between the laminate skin132and the coating134. The coating134can cover the entirety of the laminate skin132, or only a portion thereof. It is further contemplated that the coating134can be provided on the woven core130where portions of the woven core130are uncovered by the intermediate laminate skin132. The coating134can include a third stiffness and third elasticity. The third stiffness and third elasticity can be greater than that of the second stiffness and the second elasticity of the laminate skin132.

Additionally, the coating134can be an environmental barrier coating, for example, which can be used to resist oxidization or corrosion. In another example, the coating134can be a thermal barrier coating, at least partially thermally insulating the woven core130and laminate skin132. Additional non-limiting examples coatings134can include an anti-ice coating such as polyurethane, ice-phobic materials, an ultraviolet radiation coating, or an oil barrier coating such as polyethylene or polypropylene. In one additional non-limiting example, the coating134an be formed as a polypropylene base layer and a polyurethane layer provided on the polypropylene base layer. It is further contemplated that an exterior paint layer (not shown) may be provided on the exterior of the coating134, where such a paint layer may provide radiation protection, such as ultraviolet radiation.

During manufacture, the woven core130can be formed defining a specific woven structure. The specific woven structure can be a preform, specified to have a predetermined geometry, or can be cut or otherwise sized and shaped after manufacture of the woven structure, such as by cutting or grinding the woven core130. The laminate skin132can be applied directly onto the woven core130, or alternatively, it is contemplated that an adhesive material or other layer is provided therebetween. The woven structure of the woven core130provides for greater adhesion to the laminate skin132, as opposed to the adhesion between the laminate skin132and a non-woven core, and can provide for improved stiffness transition or elasticity transition between the woven core130and the coating134, whereby the difference in stiffness or elasticity between adjacent materials is less than that compared to non-adjacent materials, or an airfoil having a non-woven core.

Furthermore, the woven core130, or other woven feature, can aid in facilitating handling of a dry preform, before being injected with an interior resin or other material, which would otherwise require careful handling, thereby decreasing cost and complexity of the manufacture and formation process.

FIG.4shows a portion of a sectional view of another airfoil150. The airfoil150includes an outer wall152defining an interior154, and including a pressure side156and a suction side158. The interior154includes a core160, which can be a woven core, a woven braided core, or a foam core in non-limiting examples. A laminate skin162can be provided over the core160. The laminate skin162can surround the core160, and can partially or wholly cover the core160. A coating164can be provided on the laminate skin162. The coating164can be resistant to water or ice, such as a hydrophobic or ice-phobic coating. Additionally, it is contemplated that the coating164can be textured, having a non-flat or non-planar texture166defined in three dimensions. Such a non-planar texture166can provide for increased resistance to ice formation, for example. Furthermore, the textured structure can be used to increase or otherwise vary the stiffness defined for the coating164, which can be used to create a stronger exterior surface for the airfoil150while providing for improved stiffness transition utilizing the laminate skin between the coating164and the core160.

FIG.5shows a portion of a sectional view of another airfoil200having an outer wall202defining an interior204, and including a suction side208. While shown as generally parallel, it should be understood that the outer wall202at the suction side208can be curved, similar to that shown inFIG.3, and the section shown is schematic.

The airfoil200includes a woven core210. A laminate skin212can be provided over the woven core210. A first coating214is provided over the laminate skin212and a second coating216is provided over the first coating214. In one example, it is contemplated that the airfoil200can include a plurality of coating layers or a set of coating layers, having any number of coatings provided exterior of the laminate skin212.

The first coating214can be an environmental barrier coating, for example, such as polyurethane, polypropylene, or polyethylene. The second coating216can be another environmental barrier coating, which can be different than that of the first coating214, thereby providing greater environmental shielding than one coating alone. Additional exterior layers are contemplated, such as an ultraviolet (UV) radiation paint layer, or adhesive layers between the first and second coatings214,216.

Utilizing the woven core210with the laminate skin212and the first and second coatings214,216can provide for improved stiffness transition or elasticity transition between layers, where the stiffness transition or elasticity transition between the first coating214and the laminate skin212is less than the stiffness transition or elasticity transition between the second coating216and the laminate skin212, or wherein the overall stiffness transition or elasticity transition, either among the woven core210and the second coating216, or measured as an average change in stiffness among the woven core210and the layers212,214,216, that is less than an average stiffness among a similar system without using the woven core210and the layers212,214,216.

FIG.6shows a portion of a sectional view of another airfoil250having an outer wall252defining an interior254, and including a suction side258. While shown as generally parallel, it should be understood that the outer wall252at the suction side258can be curved, similar to that shown inFIG.3, and the section shown is schematic.

The airfoil250includes a woven core260. A laminate skin262can be provided over the woven core260. A coating264is provided over the laminate skin262, and an intermediate adhesive layer266is provided between the coating264and the laminate skin262to facilitate adhesion of the coating264to the laminate skin262.

Utilizing the woven core260can provide for improved stiffness or elasticity transition between the woven core260of the airfoil, the laminate skin262, and the coating264. Utilizing the intermediate adhesive layer266can provide for improved adhesion for the coating264to the laminate skin262. Furthermore, the intermediate adhesive layer266can include a stiffness or elasticity that is between the coating264and the laminate skin262, further facilitating the overall stiffness or elasticity transition for the airfoil250. Therefore, the airfoil250can provide for an improved core structure having an improved stiffness or elasticity transition between the woven core260and exterior layers forming the outer wall252of the airfoil250.

FIG.7shows a portion of a sectional view of another airfoil300having an outer wall302defining an interior304, and including a suction side308. While shown as generally flat, it should be understood that the outer wall302at the suction side308can be curved, similar to that shown inFIG.3, and the section as shown is schematic. The airfoil300includes a woven core310. A laminate skin312can be provided over the woven core310and a first coating314is provided over the laminate skin312. A second coating316is provided over the first coating314. An intermediate adhesive layer318can be provided between the first coating314and the second coating316.

The first and second coatings314,316can be environmental barrier coatings, such as an oil barrier coating, an anti-ice or ice-phobic coating, or a UV radiation coating in non-limiting examples. Additionally, the first and second coatings314,316can be a paint coating, while including environmental barrier properties. The stiffness or elasticity for the first and second coatings314,316can be arranged such that the stiffness or elasticity transition among adjacent layers is minimized. For example, a stiffness for the first coating314can be nearer to the stiffness of the laminate skin312, while the second coating316can include a stiffness nearer to that of the first coating314than that of the laminate skin312. Furthermore, the intermediate adhesive layer318can include a stiffness that is between the stiffness of the first coating314and the second coating316, providing for an improved stiffness or elasticity transition between the exterior layers. Such an improved stiffness or elasticity transition can increase durability of the airfoil300, extending lifetime and reducing maintenance costs.

In alternative examples, it is contemplated that the airfoil300can include a plurality of coatings or a set of coatings, having an adhesive layer provided between at least some or each of the coatings.

FIG.8shows a portion of a sectional view of another airfoil350having an outer wall352defining an interior354, and including a pressure side356and a suction side358. While shown as flat and generally parallel to one another, it should be understood that the outer wall352at the pressure side356and suction side358can be curved, similar to that shown inFIG.3, and the section shown is schematic.

The airfoil350includes a woven core360. A laminate skin362can be provided over the woven core360. A woven sheath or woven layer364is provided over the laminate skin362, and an exterior coating366is provided over the woven layer364. The woven layer364can includes a woven structure, such as that described herein, and can be formed by a Jacquard loom or other three-dimensional weaving machine.

The stiffness or elasticity for the woven layer364can be arranged such that the stiffness or elasticity transition between the laminate skin362and the exterior coating366is decreased among adjacent portions. For example, a stiffness or elasticity for woven layer364can be nearer to the stiffness of the laminate skin362than that of the woven core360. In this way, the woven layer364provides for a stiffness or elasticity transition between the laminate skin362and any exterior layers, such as the exterior coating366. Such an improved transition can increase durability of the airfoil350, extending lifetime and reducing maintenance costs.

In alternative examples, the woven layer364and the exterior coating366can be formed as multiple bi-layers, a set of woven layers and a set of coating layers, in non-limiting examples. In such examples, it is contemplated that one or more intervening adhesive layers can be utilized between layers. Similar increasing or decreasing values for the stiffness or elasticity can be applied to any additional layers, such that the transition among layers can be matched nearer to adjacent layers to improve transition among multiple layers.

FIG.9shows a portion of a sectional view of another airfoil400having an outer wall402defining an interior404, and including a pressure side406and a suction side408. While shown as generally parallel, it should be understood that the outer wall402can be curved, similar to that shown inFIG.3, and the section as shown is schematic.

The airfoil400includes a foam inner core410and a woven outer core412provided over the foam inner core410. A laminate skin414can be provided over the woven outer core412. A woven sheath or woven layer416is provided over the laminate skin414, and a coating layer418is provided over the woven layer416.

The stiffness for the woven outer core412can provide for a stiffness or elasticity transition between the foam inner core410and the laminate skin414. Utilizing the foam inner core410with the woven outer core412can utilize the structural benefits of the foam inner core410, while providing an intermediate transition between the foam inner core410and the laminate skin414. Furthermore, manufacture and handling of the foam inner core410can be facilitated with an exterior woven outer core412prior to resin impregnation or other finishing, facilitating manufacture. Similarly, a stiffness or elasticity for woven layer416can be nearer to the stiffness or elasticity of the laminate skin414, while the coating418can be nearer to that of the woven layer416than that of the laminate skin414. In this way, the woven layer416provides for a smoother transition between the laminate skin414and any exterior layers, such as the coating418. Such an improved stiffness transition can increase durability of the airfoil400, extending lifetime and reducing maintenance costs.

To the extent not already described, the different features and structures of the various embodiments can be used in combination, or in substitution with each other as desired. That one feature is not illustrated in all of the embodiments is not meant to be construed that it cannot be so illustrated, but is done for brevity of description. Thus, the various features of the different embodiments can be mixed and matched as desired to form new embodiments, whether or not the new embodiments are expressly described. All combinations or permutations of features described herein are covered by this disclosure.

A turbine engine comprising: a fan section, a compressor section, combustor section, and turbine section in serial flow arrangement, and defining an engine longitudinal axis; and an airfoil, provided in one of the fan section, the compressor section, or the turbine section, the airfoil including an outer wall having a pressure side and a suction side, and the airfoil rotatable about the engine longitudinal axis, the airfoil comprising: a woven core having a first stiffness, a laminate skin, having a second stiffness, located exteriorly of the woven core, and a first coating, having a third stiffness, located exteriorly of the laminate skin; wherein the second stiffness is between the first stiffness and the third stiffness, providing a stiffness transition between the first stiffness and the third stiffness.

The turbine engine of any preceding clause wherein the woven core includes a three-dimensional weave pattern.

The turbine engine of any preceding clause wherein the three-dimensional weave pattern includes a braided pattern.

The turbine engine of any preceding clause wherein the first coating is an environmental barrier coating

The turbine engine of any preceding clause wherein the first coating partially covers the laminate skin.

The turbine engine of any preceding clause further comprising an exterior coating located exteriorly of the first coating.

The turbine engine of any preceding clause further comprising an adhesive layer located between the first coating and the exterior coating.

The turbine engine of any preceding clause wherein the adhesive layer includes a stiffness that is intermediate between the third stiffness and a fourth stiffness defined by the exterior coating.

The turbine engine of any preceding clause further comprising an adhesive layer located between the laminate skin and the first coating.

The turbine engine of any preceding clause further comprising a woven layer located between the laminate skin and the first coating.

The turbine engine of any preceding clause wherein the woven layer includes a braided pattern.

The turbine engine of any preceding clause wherein the woven layer includes a different weave pattern than that of the woven core.

The turbine engine of any preceding clause further comprising a foam core located within the woven core.

An airfoil having an outer wall defining an interior, the airfoil comprising: a three-dimensional woven core; and a laminate skin located exteriorly of the three-dimensional woven core; and a three-dimensional woven layer located exteriorly of the laminate skin.

The airfoil of any preceding clause wherein the three-dimensional woven layer includes an elasticity that is greater than an elasticity for the laminate skin.

The airfoil of any preceding clause further comprising a foam core located within the three-dimensional woven core.

The airfoil of any preceding clause wherein the three-dimensional woven core includes a weave pattern that is different than another weave pattern defining the three-dimensional woven layer.

A turbine engine comprising: a fan section, a compressor section, combustor section, and turbine section in serial flow arrangement, and defining an engine longitudinal axis; and an airfoil rotatable about the engine longitudinal axis, the airfoil comprising: a woven core; a laminate skin at least partially covering the woven core; and at least one coating located exteriorly of the laminate skin.

The turbine engine of any preceding clause further comprising a woven layer located exteriorly of the laminate skin.

The turbine engine of any preceding clause wherein the at least one coating is located exteriorly of the woven layer.

The turbine engine of any preceding clause wherein the woven core is dry.

The turbine engine of any preceding clause wherein the woven core is impregnated with a resin and cured.

The turbine engine of any preceding clause wherein the woven core is made from the same material as the laminate skin.

The turbine engine of any preceding clause wherein the woven core is made from a different material than the laminate skin.

The turbine engine of any preceding clause wherein the laminate skin is pre-impregnated.

The turbine engine of any preceding clause wherein the laminate skin is a fiber placed or dry fiber coating.

The turbine engine of any preceding clause wherein the at least one coating includes at least one of an adhesive, a paint coating, polyurethane, polypropylene, or polyethylene.

The turbine engine of any preceding clause wherein the at least one coating is an environmental barrier coating.

The turbine engine of any preceding clause wherein the environmental barrier coating is one of an oil barrier coating, an anti-ice coating, or an ultraviolet (UV) radiation coating.

The turbine engine of any preceding clause wherein the woven core is formed as a preform, and the laminate skin is applied to the preform.

The turbine engine of any preceding clause wherein the at least one coating is provided as a polyurethane base layer and an exterior paint layer provided on the polyurethan base layer.

The turbine engine of any preceding clause wherein the at least one coating is provided as a polypropylene base layer and an exterior polyurethane layer provided on the polypropylene base layer.

The turbine engine of any preceding clause further comprising a paint layer provided on the exterior polyurethan layer.

The turbine engine of any preceding clause wherein the at last one coating further includes multiple layers, with a first layer comprising an adhesive layer, and at least one exterior layer provided on the adhesive layer.

The turbine engine of any preceding clause wherein the at least one exterior layer is a polypropylene layer.

The turbine engine of any preceding clause wherein the woven core and the laminate skin are formed as a composite preform.