Patent Description:
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.

Traditionally, airfoils include a metallic spar that is formed with or coupled to a hub of an airfoil.

Aspects of the disclosure herein are directed to a component, i.e. a blade or vane, for a turbine engine, the component having an airfoil with a metallic spar, a composite spar, and a stiffener together defining a spar assembly. The metallic spar can be shaped to receive a portion of the composite spar. The stiffener bonded to at least one of the airfoil, metallic spar, or the composite spar. An example of related prior art is given in <CIT>, which according to its abstract discloses a structural spar of an aerodynamic blade, preferably made of a composite material, connected to the rotor hub through a primary connecting cuff to which the spar is bonded, and which normally constitutes the sole load path, and redundantly through a secondary cuff which is also bonded to the spar and which is normally unloaded and becomes loaded only upon failure of the bond between the spar and the primary cuff; the cuffs are metallic.

The present disclosure is directed at a component according to independent claim <NUM>, with preferred embodiments being defined in dependent claims.

Reference will now be made in detail to present embodiments of the disclosure, one or more examples of which are illustrated in the accompanying drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the disclosure.

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 metal, non-metallic, or a combination of metal 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 fiber, polymeric resin, thermoplastic, bismaleimide (BMI), polyimide materials, epoxy resin, glass fiber, 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 preimpregnated 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 the 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 provides 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.

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 (Al<NUM>O<NUM>), silicon dioxide (SiO<NUM>), 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 (Al<NUM>O<NUM>), silicon dioxide (SiO<NUM>), 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 (Al<NUM>O<NUM>), silicon dioxide (SiO<NUM>), aluminosilicates, and mixtures thereof. Aluminosilicates can include crystalline materials such as mullite (3Al<NUM>O<NUM> 2SiO<NUM>), as well as glassy aluminosilicates.

In certain non-limiting examples, the reinforcing fibers can be bundled and/or coated prior to inclusion within the ceramic matrix. For example, bundles of the fibers can be formed as a reinforced tape, such as a unidirectional reinforced tape. A plurality of the tapes can be laid up together to form a preform component. The bundles of fibers can be impregnated with a slurry composition prior to forming the preform or after formation of the preform. The preform can then undergo thermal processing, such as a cure or burn-out to yield a high char residue in the preform, and subsequent chemical processing, such as melt-infiltration with silicon, to arrive at a component formed of a CMC material having a desired chemical composition.

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.

As used herein, the terms "first" and "second" may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

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 term "fore" or "forward" means in front of something and "aft" or "rearward" means behind something.

The term "fluid" may be a gas or a liquid, or multi-phase. The term "fluid communication" means that a fluid is capable of making the connection between the areas specified.

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.

All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, aft, etc.) are only used for identification purposes to aid the reader's understanding of the present disclosure, and do not create limitations, particularly as to the position, orientation, or use of aspects of the disclosure described herein. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and can include intermediate structural elements between a collection of elements and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and in fixed relation to one another. The exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the drawings attached hereto can vary.

Furthermore, as used herein, the term "set" or a "set" of elements can be any number of elements, including only one.

In certain aspects of the present disclosure, an unducted or open rotor turbine engine includes a set of circumferentially spaced fan blades, which extend, exteriorly, beyond a nacelle encasing an engine core.

<FIG> is a schematic cross-sectional diagram of a turbine engine, specifically an open rotor or unducted turbine engine <NUM> for an aircraft. The unducted turbine engine <NUM> has a generally longitudinally extending axis or engine centerline <NUM> extending from a forward end <NUM> to an aft end <NUM>. The unducted turbine engine <NUM> includes, in downstream serial flow relationship, a set of circumferentially spaced blades or propellers defining a fan section <NUM> including a fan <NUM>, a compressor section <NUM> including a booster or low pressure (LP) compressor <NUM> and a high pressure (HP) compressor <NUM>, a combustion section <NUM> including a combustor <NUM>, a turbine section <NUM> including a HP turbine <NUM>, and a LP turbine <NUM>, and an exhaust section <NUM>. The unducted turbine engine <NUM> as described herein is meant as a non-limiting example, and other architectures are possible, such as, but not limited to, a steam turbine engine, a supercritical carbon dioxide turbine engine, or any other suitable turbine engine.

An exterior surface, defined by a housing or nacelle <NUM>, of the unducted turbine engine <NUM> extends from the forward end <NUM> of the unducted turbine engine <NUM> toward the aft end <NUM> of the unducted turbine engine <NUM> and covers at least a portion of the compressor section <NUM>, the combustion section <NUM>, the turbine section <NUM>, and the exhaust section <NUM>. The fan section <NUM> can be positioned at a forward portion of the nacelle <NUM> and extend radially outward from the nacelle <NUM> of the unducted turbine engine <NUM>, specifically, the fan section <NUM> extends radially outward from the nacelle <NUM>. The fan section <NUM> includes a set of fan blades <NUM>, and a set of stationary fan vanes <NUM> downstream the set of fan blades <NUM>, both disposed radially about the engine centerline <NUM>. The unducted turbine engine <NUM> includes any number of one or more sets of rotating blades or propellers (e.g., the set of fan blades <NUM>) disposed upstream of the set of stationary fan vanes <NUM>. As a non-limiting example, the unducted turbine engine <NUM> can include multiple sets of fan blades <NUM> or fan vanes <NUM>. As such, the unducted turbine engine <NUM> is further defined as an unducted single-fan turbine engine. The unducted turbine engine <NUM> is further defined by the location of the fan section <NUM> with respect to the combustion section <NUM>. The fan section <NUM> can be upstream, downstream, or in-line with the axial positioning of the combustion section <NUM>.

The compressor section <NUM>, the combustion section <NUM>, and the turbine section <NUM> are collectively referred to as an engine core <NUM>, which generates combustion gases. The engine core <NUM> is surrounded by an engine casing <NUM>, which is operatively coupled with a portion of the nacelle <NUM> of the unducted turbine engine <NUM>.

A HP shaft or spool <NUM> disposed coaxially about the engine centerline <NUM> of the unducted turbine engine <NUM> drivingly connects the HP turbine <NUM> to the HP compressor <NUM>. A LP shaft or spool <NUM>, which is disposed coaxially about the engine centerline <NUM> of the unducted turbine engine <NUM> within the larger diameter annular HP spool <NUM>, drivingly connects the LP turbine <NUM> to the LP compressor <NUM> and fan <NUM>. The spools <NUM>, <NUM> are rotatable about the engine centerline <NUM> and coupled to a set of rotatable elements, which collectively define a rotor <NUM>.

It will be appreciated that the unducted turbine engine <NUM> is either a direct drive or integral drive engine utilizing a reduction gearbox coupling the LP shaft or spool <NUM> to the fan <NUM>.

The LP compressor <NUM> and the HP compressor <NUM>, respectively, include a set of compressor stages <NUM>, <NUM>, in which a set of compressor blades <NUM>, <NUM> rotate relative to a corresponding set of static compressor vanes <NUM>, <NUM> (also called a nozzle) to compress or pressurize the stream of fluid passing through the stage. In a single compressor stage <NUM>, <NUM>, multiple compressor blades <NUM>, <NUM> are provided in a ring and extend radially outwardly relative to the engine centerline <NUM>, from a blade platform to a blade tip, while the corresponding static compressor vanes <NUM>, <NUM> are positioned upstream of and adjacent to the compressor blades <NUM>, <NUM>. It is noted that the number of blades, vanes, and compressor stages shown in <FIG> were selected for illustrative purposes only, and that other numbers are possible.

The compressor blades <NUM>, <NUM> for a stage of the compressor are mounted to a disk <NUM>, which is mounted to the corresponding one of the HP and LP spools <NUM>, <NUM>, with each stage having its own disk <NUM>. The static compressor vanes <NUM>, <NUM> for a stage of the compressor are mounted to the engine casing <NUM> in a circumferential arrangement.

The HP turbine <NUM> and the LP turbine <NUM>, respectively, include a set of turbine stages <NUM>, <NUM>, in which a set of turbine blades <NUM>, <NUM> are rotated relative to a corresponding set of static turbine vanes <NUM>, <NUM> (also called a nozzle) to extract energy from the stream of fluid passing through the stage. In a single turbine stage <NUM>, <NUM>, multiple turbine blades <NUM>, <NUM> are provided in a ring and extends radially outwardly relative to the engine centerline <NUM>, from a blade platform to a blade tip, while the corresponding static turbine vanes <NUM>, <NUM> are positioned upstream of and adjacent to the turbine blades <NUM>, <NUM>. It is noted that the number of blades, vanes, and turbine stages shown in <FIG> were selected for illustrative purposes only, and that other numbers are possible.

The turbine blades <NUM>, <NUM> for a stage of the turbine section <NUM> are mounted to a disk <NUM>, which is mounted to the corresponding one of the HP and LP spools <NUM>, <NUM>, with each stage having a dedicated disk <NUM>. The static turbine vanes <NUM>, <NUM> for a stage of the turbine section <NUM> are mounted to the engine casing <NUM> in a circumferential arrangement.

Rotary portions of the unducted turbine engine <NUM>, such as the blades <NUM>, <NUM><NUM>, <NUM> among the compressor section <NUM> and the turbine section <NUM> are also referred to individually or collectively as the rotor <NUM>. As such, the rotor refers to the combination of rotating elements throughout the unducted turbine engine <NUM>.

Complementary to the rotor portion, the stationary portions of the unducted turbine engine <NUM>, such as the static vanes <NUM>, <NUM>, <NUM>, <NUM> among the compressor section <NUM> and the turbine section <NUM> are also referred to individually or collectively as a stator <NUM>. As such, the stator <NUM> refers to the combination of non-rotating elements throughout the unducted turbine engine <NUM>.

The nacelle <NUM> is operatively coupled to the unducted turbine engine <NUM> and covers at least a portion of the engine core <NUM>, the engine casing <NUM>, or the exhaust section <NUM>. At least a portion of the nacelle <NUM> extends axially forward or upstream the illustrated position. For example, the nacelle <NUM> extends axially forward such that a portion of the nacelle <NUM> overlays or covers a portion of the fan section <NUM> or a booster section (not illustrated) of the unducted turbine engine <NUM>.

During operation of the unducted turbine engine <NUM>, a freestream airflow <NUM> flows against a forward portion of the unducted turbine engine <NUM>. A portion of the freestream airflow <NUM> enters an annular area <NUM> defined by a swept area between an outer surface of the nacelle and the tip of the blade, with this air flow being an inlet airflow <NUM>. A portion of the inlet airflow <NUM> enters the engine core <NUM> and is described as a working airflow <NUM>, which is used for combustion within the engine core <NUM>.

More specifically, the working airflow <NUM> flows into the LP compressor <NUM>, which then pressurizes the working airflow <NUM> thus defining a pressurized airflow that is supplied to the HP compressor <NUM>, which further pressurizes the air. The working airflow <NUM>, or the pressurized airflow, from the HP compressor <NUM> is mixed with fuel in the combustor <NUM> and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine <NUM>, which drives the HP compressor <NUM>. The combustion gases are discharged into the LP turbine <NUM>, which extracts additional work to drive the LP compressor <NUM>, and the working airflow <NUM>, or exhaust gas, is ultimately discharged from the unducted turbine engine <NUM> via the exhaust section <NUM>. The driving of the LP turbine <NUM> drives the LP spool <NUM> to rotate the fan <NUM> and the LP compressor <NUM>. The working airflow <NUM>, including the pressurized airflow and the combustion gases, defines a working airflow that flows through the compressor section <NUM>, the combustion section <NUM>, and the turbine section <NUM> of the unducted turbine engine <NUM>.

The inlet airflow <NUM> flows through the set of fan blades <NUM> and over the nacelle <NUM> of the unducted turbine engine <NUM>. Subsequently, the inlet airflow <NUM> flows over at least a portion of the set of stationary fan vanes <NUM>, which directs the inlet airflow <NUM> such that it is transverse toward the engine centerline <NUM>. The inlet airflow <NUM> then flows past the set of stationary fan vanes <NUM>, following the curvature of the nacelle <NUM> and toward the exhaust section <NUM>. A pylon <NUM> mounts the unducted turbine engine <NUM> to an exterior structure (e.g., a fuselage of an aircraft, a wing, a tail wing, etc.).

The working airflow <NUM> and at least some of the inlet airflow <NUM> merge downstream of the exhaust section <NUM> of the unducted turbine engine <NUM>. The working airflow <NUM> and the inlet airflow <NUM>, together, form an overall thrust of the unducted turbine engine <NUM>.

It is contemplated that a portion of the working airflow <NUM> is drawn as bleed air <NUM> (e.g., from the compressor section <NUM>). The bleed air <NUM> provides an airflow to engine components requiring cooling. The temperature of the working airflow <NUM> exiting the combustor <NUM> is significantly increased with respect to the working airflow <NUM> within the compressor section <NUM>. As such, cooling provided by the bleed air <NUM> is necessary for operating of such engine components in heightened temperature environments or a hot portion of the unducted turbine engine <NUM>. In the context of a turbine engine, the hot portions of the engine are normally downstream of the combustor <NUM>, especially the turbine section <NUM>, with the HP turbine <NUM> being the hottest portion as it is directly downstream of the combustion section <NUM>. Other sources of cooling fluid are, but are not limited to, fluid discharged from the LP compressor <NUM> or the HP compressor <NUM>.

<FIG> is a schematic perspective view of an aircraft <NUM> including a generic unducted turbine engine <NUM> suitable for use as the unducted turbine engine <NUM> of <FIG>. The aircraft <NUM> includes a fuselage <NUM> with an exterior surface. At least one wing <NUM> and a tail wing <NUM> extend from the fuselage <NUM>. The tail wing <NUM> is operably coupled to and spaced from the fuselage <NUM> via a tail wing pylon <NUM>. The unducted turbine engine <NUM> is operably coupled to the exterior surface of the fuselage <NUM> via a pylon <NUM>. The unducted turbine engine <NUM> includes a set of circumferentially spaced fan blades <NUM>. A set of stationary fan vanes <NUM> is provided downstream of the set of circumferentially spaced fan blades <NUM>. The fuselage <NUM> extends between a nose <NUM> and a tail <NUM> and includes a fuselage centerline <NUM> extending therebetween.

Additionally, while the tail wing <NUM> is a T-wing tail wing (e.g., the tail wing <NUM> as illustrated), other conventional tail wings are contemplated such as, a cruciform tail wing, an H-tail, a triple tail, a V-tail, an inverted tail, a Y-tail, a twin-tail, a boom-mounted tail, or a ring tail, all of which are referred to herein as the tail wing <NUM>.

<FIG> is schematic illustration of an engine component in the form of, by way of non-limiting example, a blade assembly <NUM>. The blade assembly <NUM> includes an airfoil <NUM> illustrated, by way of example, as a composite blade. The airfoil <NUM> can be, by way of non-limiting example, a blade of the set of fan blades <NUM>, <NUM> or a blade from the compressor blades <NUM>, <NUM> or the turbine blades <NUM>, <NUM>. Further, the engine component can be a vane assembly, where the airfoil <NUM> is a vane of the set of stationary fan vanes <NUM>, <NUM>, or a vane of the static vanes <NUM>, <NUM>, <NUM>, <NUM>. It is contemplated that the airfoil <NUM> can be a blade, vane, airfoil, or other component of any turbine engine, such as, but not limited to, a gas turbine engine, a turboprop engine, a turboshaft engine, or a turbofan engine.

The airfoil <NUM> includes a wall <NUM> bounding an interior <NUM>. The wall <NUM> defines an exterior surface <NUM> extending radially between a leading edge <NUM> and a trailing edge <NUM> to define a chordwise direction (denoted "C"). The exterior surface <NUM> can further extend between a root <NUM> and a tip <NUM> to define a spanwise direction (denoted "S"). The wall <NUM> can be a composite wall made of one or more layers of material. The one or more layers of material can be applied during the same stage or different stages of the manufacturing of the airfoil <NUM>.

By way of non-limiting example, wall <NUM> can include at least a polymer matrix composite (PMC) portion or a polymeric portion. The polymer matrix composite can include, but is not limited to, a matrix of thermoset (epoxies, phenolics) or thermoplastic (polycarbonate, polyvinylchloride, nylon, acrylics) and embedded glass, carbon, steel, or Kevlar fibers.

The blade assembly <NUM> further includes a spar assembly <NUM>. The spar assembly <NUM> including, but not limited to, a metallic spar <NUM>, a composite spar <NUM>, and a stiffener <NUM>. The stiffener <NUM> can be formed from, at least in part, by a metal. It is further contemplated that the stiffener <NUM> is metallic. The stiffener <NUM> defines a wing portion <NUM> of the spar assembly <NUM>. The spar assembly <NUM> defines an axis A extending radially from and perpendicular to the engine centerline <NUM> (<FIG>). The airfoil <NUM> is mounted to the spar assembly <NUM> near the root <NUM> via a metallic trunnion <NUM> defining a hub <NUM>. The airfoil <NUM> has a span length (denoted "L") measured along the spanwise direction S from the hub <NUM> at <NUM>% the span length L to the tip <NUM> at <NUM>% the span length L. The span length L can run parallel to the axis A and be defined as the maximum distance between the root <NUM> and the tip <NUM> of the airfoil <NUM>. An entirety of the spar assembly <NUM> can be located below <NUM>% of the span length L. The wall <NUM> can circumscribe and/or surround at least a portion of the spar assembly <NUM>. At least a portion of the spar assembly <NUM> can be bonded to the wall <NUM>.

The metallic spar <NUM> can be formed from metals such as, but not limited to, titanium, iron, aluminum, stainless steel and nickel alloys. At least a portion of the metallic spar <NUM> can be located within the interior <NUM>. The metallic spar <NUM> can be integral with the metallic trunnion <NUM>. The metallic spar <NUM> can be located above the hub <NUM> and within the interior <NUM>. Remaining portions of the metallic trunnion <NUM> can be located below the airfoil <NUM>.

The composite spar <NUM> can be formed from a polymeric material or other non-metal materials. At least a portion of the composite spar <NUM> can be located within the interior <NUM>. The composite spar <NUM> can extend in the spanwise direction S between a spar root <NUM> and a spar tip <NUM>. A majority of the composite spar <NUM> including the spar tip <NUM> can be located within the interior <NUM>.

The stiffener <NUM> can be formed from metals such as, but not limited to, titanium, iron, aluminum, stainless steel, and nickel alloys. In one aspect, the stiffener <NUM> can be located within the interior <NUM>. In another aspect the stiffener <NUM> can be located on the exterior surface <NUM> of the wall <NUM>. It is further contemplated that a portion of the stiffener <NUM> can be located within the interior <NUM> while other portions are located on the exterior surface <NUM> of the wall <NUM>.

It is also contemplated that one or more layers of adhesive (not shown) can be applied between the wall <NUM> and any portion of the spar assembly <NUM>. Further, it is contemplated that the adhesive can be absorbed by the wall <NUM>, and/or one or more portions of the spar assembly <NUM>. The adhesive can include resin and phenolics, wherein the adhesive can require curing at elevated temperatures or other hardening technique. It should be understood that any part of the spar assembly <NUM> can be located in the interior <NUM> and/or on the exterior surface <NUM> and bonded at those positions accordingly.

Turning to <FIG>, a perspective view of a spar assembly <NUM> according to an aspect of the disclosure herein is illustrated. The spar assembly <NUM> is similar to the spar assembly <NUM>, therefore, like parts of the spar assembly <NUM> will be identified with like numerals increased by <NUM>, with it being understood that the description of the like parts of the spar assembly <NUM> applies to the spar assembly <NUM>, except where noted.

A metallic spar <NUM> can extend from a base <NUM> of a metallic trunnion <NUM> to define a hub <NUM> of the spar assembly <NUM>. The metallic spar <NUM> includes a set of walls <NUM>, each wall of the set of walls <NUM> defining an interior surface spaced from each other a first distance (denoted "D1"), to define a socket <NUM> illustrated in dashed line. The socket <NUM> is closed by the set of walls <NUM> on at least two sides <NUM>. The socket <NUM> is open on another two opposing sides <NUM> to define side openings <NUM>. The socket <NUM> can have a socket length SL where the socket <NUM> has a cross-sectional area (denoted "CA") equal to the socket length SL multiplied by the first distance D1 (CA = SL x D1). The cross-sectional area CA can range between <NUM> in<NUM> and <NUM> in<NUM> (<NUM><NUM> and <NUM><NUM>).

A composite spar <NUM> is received in the socket <NUM>. The composite spar <NUM> can have a first thickness (denoted "T1") that is less than the first distance D1. It is further contemplated that the first thickness T1 is almost equal to the first distance D1 to provide a snug fit of the composite spar <NUM> in the socket <NUM>. The composite spar <NUM> can include a body portion <NUM> and a wing portion <NUM>. The wing portion <NUM> and the body portion <NUM> forming an upside down "T" shape. The body portion <NUM> can extend in the spanwise direction S within the socket <NUM> from a spar root <NUM>. The wing portion <NUM> can extend in a direction substantially perpendicular to the spanwise direction S from the body portion <NUM> out of the side openings <NUM>.

A stiffener <NUM> is bonded with at least a portion of the composite spar <NUM>. The stiffener <NUM> can be bonded to an outer surface of the wing portion <NUM> on one or both sides of the composite spar <NUM>. In one non-limiting example the stiffener <NUM> is bonded to the wing portion <NUM> of the composite spar <NUM>. A single stiffener 243c can extend along the composite spar <NUM> through the socket <NUM> and overlap with the metallic spar <NUM>. It is further contemplated that the stiffener <NUM> can be multiple parts 243a, 243b located on the wing portion <NUM> outside of the socket <NUM>. In this example the stiffener <NUM> does not extend through the socket <NUM> or overlap with the metallic spar <NUM>. The stiffener <NUM> can have a second thickness (denoted "T2"). Further, the stiffener <NUM> includes at least one tapered edge <NUM> that tapers from the second thickness T2 toward the composite spar <NUM>. In one non-limiting example the stiffener <NUM> has all tapered edges <NUM>. In the aspect where the stiffener is a single stiffener 243c, the first thickness T1 and the second thickness T2 together are almost equal to the first distance D1 (D1 ≈ T1+T2) to provide the snug fit previously described herein. While illustrated as having uniform thicknesses T1, T2, it is contemplated that the thicknesses as described herein can change.

In one aspect, one or more layers of adhesive (not shown) can be applied between the stiffener <NUM> and the wing portion <NUM> of the composite spar <NUM>. Additionally, or alternatively, one or more layers of adhesive (not shown) can be applied between the metallic spar <NUM> and the stiffener <NUM> or the composite spar <NUM>, or between both the metallic spar <NUM> and the stiffener <NUM> and the composite spar <NUM>.

Turning to <FIG>, a spar assembly <NUM> according to another aspect of the disclosure herein is illustrated. The spar assembly <NUM> is similar to the spar assembly <NUM>, therefore, like parts of the spar assembly <NUM> will be identified with like numerals increased by <NUM>, with it being understood that the description of the like parts of the spar assembly <NUM> applies to the spar assembly <NUM>, except where noted.

A metallic spar <NUM> can extend from a base <NUM> of a metallic trunnion <NUM> to define a hub <NUM> of the spar assembly <NUM>. The metallic spar <NUM> can include a set of walls <NUM> spaced from each other to define a socket <NUM>. The socket <NUM> is closed on three sides, defined by the set of walls <NUM>. The socket <NUM> is open on a remaining side to define a side opening <NUM>. In other words, the socket <NUM> has a sideways "U" shape <NUM>. A thru hole <NUM> is located opposite the side opening <NUM> in the set of walls <NUM> at the hub <NUM>. At least a portion of the thru hole <NUM> is formed in the base <NUM>.

A composite spar <NUM> is received in the socket <NUM>. The composite spar <NUM> includes a body portion <NUM> (illustrated in dashed line) and a wing portion <NUM> (illustrated in solid line). The wing portion <NUM> and the body portion <NUM> forming an upside down "T" shape. The body portion <NUM> can extend in the spanwise direction S within the socket <NUM> from a spar root <NUM>. The wing portion <NUM> can extend in a direction substantially perpendicular to the spanwise direction S from the body portion <NUM> out of the side opening <NUM> on one side and out of the thru hole <NUM> on the other side of the body portion <NUM>.

A stiffener <NUM> is bonded with at least a portion of the composite spar <NUM>. The stiffener <NUM> can be bonded to an outer surface of the wing portion <NUM> on one or both sides of the composite spar <NUM>. In one non-limiting example the stiffener <NUM> is bonded to the wing portion <NUM> of the composite spar <NUM>. A single stiffener 343c can extend along the composite spar <NUM> through the socket <NUM> and overlap with the metallic spar <NUM>. It is further contemplated that the stiffener <NUM> can be multiple parts 343a, 343b located on the wing portion <NUM> outside of the socket <NUM>.

A metallic spar <NUM> can extend from a base <NUM> of a metallic trunnion <NUM> to define a hub <NUM> of the spar assembly <NUM>. The metallic spar <NUM> includes a set of walls <NUM> spaced from each other to define a socket <NUM>. The metallic spar <NUM> includes a middle truss <NUM> separating the socket <NUM> into two openings <NUM>. In other words, the socket <NUM> has a sideways "H" shape <NUM>. While illustrated as in the middle, the middle truss <NUM> can be located nearer to either opening <NUM> and does not necessarily have to be located directly in the middle.

A composite spar <NUM> is received in the socket <NUM>. The composite spar <NUM> includes a body portion <NUM> (illustrated in dashed line) and a wing portion <NUM> (illustrated in solid line). The wing portion <NUM> and the body portion <NUM> forming an upside down split "T" shape, where a slot <NUM> is formed in the composite spar <NUM> to accommodate the middle truss <NUM> when assembled. The body portion <NUM> can extend in the spanwise direction S within the socket <NUM> from a spar root <NUM>. The wing portion <NUM> can extend in a direction substantially perpendicular to the spanwise direction S from the body portion <NUM> out of the openings <NUM>.

At least one stiffener <NUM> is bonded with at least a portion of the composite spar <NUM>. In one non-limiting example the stiffener <NUM> is bonded to the wing portion <NUM> of the composite spar <NUM> on both sides as illustrated. The at least one stiffener <NUM> can also be multiple parts 443a, 443b located on the wing portion <NUM> outside of the socket <NUM>. It is further contemplated that the at least one stiffener <NUM> is an extended stiffener 443c extending at least partially into the socket <NUM>.

A metallic spar <NUM> can extend from a base <NUM> of a metallic trunnion <NUM> to define a hub <NUM> of the spar assembly <NUM>. The metallic spar <NUM> includes a first wall 551a and a second wall 551b connected by a middle truss <NUM> separating a first opening 554a from a second opening 554b. While illustrated as in the middle, the middle truss <NUM> can be located nearer to either opening 554a, 554b and does not necessarily have to be located directly in the middle. Together the first opening 554a and the second opening 554b defines a socket <NUM>. The first opening 554a is defined by the first wall 551a and the middle truss <NUM> and opens in a first direction 564a. The second opening 554b is defined by the second wall 551b and the middle truss <NUM> and opens in a second direction 564b opposite the first direction 564a. In other words, the socket <NUM> has a lightning bolt or sideways "Z" shape <NUM>.

A composite spar <NUM> is received in the socket <NUM>. The composite spar <NUM> includes a slot <NUM> formed to accommodate the middle truss <NUM> when assembled. The composite spar <NUM> can extend in the spanwise direction S within the socket <NUM> from a spar root <NUM>.

At least one stiffener <NUM> is bonded with at least a portion of the metallic spar <NUM>. In one non-limiting example the at least one stiffener <NUM> is bonded to the metallic spar <NUM> to define a wing portion <NUM>. The at least one stiffener <NUM> can be multiple stiffeners 543a, 543b each defining separate wing portions 561a, 561b.

The disclosure herein applies to bonding metallic metal pieces to a composite spar/metallic trunnion interface to help disperse load transfer and provide chordwise stiffness. Benefits to the disclosure include dispersing the load transfer while being versatile in location. These bonded metallic pieces can be either internal to the spar/trunnion interface or external to the interface region. Additionally, a benefit is to reduce high composite spar stresses that are induced from chordwise bending that occurs during extreme loading conditions.

Claim 1:
A component (<NUM>) for a turbine engine (<NUM>, <NUM>), wherein the turbine engine (<NUM>, <NUM>) is selected from the group consisting of gas turbine engines, unducted turbine engines, turboprop engines, turboshaft engines, turbofan engines and turbojet engines;
wherein the component is a blade or vane, the component (<NUM>) comprising:
a wall (<NUM>) bounding an interior (<NUM>) and defining an exterior surface (<NUM>) extending between a leading edge (<NUM>) and a trailing edge (<NUM>) to define a chordwise direction (C), and radially between a root (<NUM>) and a tip (<NUM>) to define a spanwise direction (S); and
a spar assembly (<NUM>, <NUM>, <NUM>, <NUM>, <NUM>) comprising:
a hub (<NUM>, <NUM>, <NUM>, <NUM>, <NUM>);
a metallic spar (<NUM>, <NUM>, <NUM>, <NUM>, <NUM>) extending from the hub (<NUM>, <NUM>, <NUM>, <NUM>, <NUM>) in the spanwise direction (S) into the interior (<NUM>), the metallic spar (<NUM>, <NUM>, <NUM>, <NUM>, <NUM>) defining a socket (<NUM>, <NUM>, <NUM>, <NUM>) extending in the spanwise direction, the socket having at least one side opening (<NUM>, <NUM>, <NUM>, 554a, 554b);
a composite spar (<NUM>, <NUM>, <NUM>, <NUM>, <NUM>) extending in the spanwise direction (S) between a spar root (<NUM>, <NUM>, <NUM>, <NUM>, <NUM>) and a spar tip (<NUM>), at least a portion of the spar root (<NUM>, <NUM>, <NUM>, <NUM>, <NUM>) located in the socket (<NUM>, <NUM>, <NUM>, <NUM>); and
a stiffener (<NUM>, <NUM>, <NUM>, <NUM>, <NUM>, 543a, 543b) bonded to at least one of the wall (<NUM>), the metallic spar (<NUM>, <NUM>, <NUM>, <NUM>, <NUM>), or the composite spar (<NUM>, <NUM>, <NUM>, <NUM>, <NUM>), at least a portion of the stiffener extending at least in part through the at least one side opening (<NUM>, <NUM>, <NUM>, 554a, 554b).