Patent Description:
A gas turbine engine typically includes a fan assembly and a turbomachine. The turbomachine generally includes an inlet, one or more compressors, a combustor, and at least one turbine. The compressors compress air which is channeled to the combustor where it is mixed with fuel. The mixture is then ignited for generating hot combustion gases. The combustion gases are channeled to the turbine(s) which extracts energy from the combustion gases for powering the compressor(s), as well as for producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator. In a turbofan engine, the fan assembly generally includes a fan having a plurality of airfoils or fan blades extending radially outwardly from a central hub and/or a disk. During certain operations, the fan blades provide an airflow into the turbomachine and over the turbomachine to generate thrust. <CIT> discloses a lightweight, impact-resistant gas turbine blade, such as an aircraft engine fan blade, which has a metal solid section, composite or structural/syntactic foam segments, and metal solid spars all attached together to define an airfoil portion. The solid section includes the leading edge, blade tip, and trailing edge. The segments together are bounded in part by the solid section near the leading edge, blade tip, and trailing edge. The solid spars separate and are attached to the segments.

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.

As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. The terms "includes" and "including" are intended to be inclusive in a manner similar to the term "comprising. " Similarly, the term "or" is generally intended to be inclusive (i.e., "A or B" is intended to mean "A or B or both"). The term "at least one of" in the context of, e.g., "at least one of A, B, and C" refers to only A, only B, only C, or any combination of A, B, and C. In addition, here and throughout the specification and claims, range limitations may be combined and/or interchanged. Such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

Accordingly, a value modified by a term or terms, such as "generally," "about," "approximately," and "substantially," are not to be limited to the precise value specified. For example, the approximating language may refer to being within a <NUM> percent margin, i.e., including values within ten percent greater or less than the stated value. In this regard, for example, when used in the context of an angle or direction, such terms include within ten degrees greater or less than the stated angle or direction.

" In addition, references to "an embodiment" or "one embodiment" does not necessarily refer to the same embodiment, although it may. Any implementation described herein as "exemplary" or "an embodiment" is not necessarily to be construed as preferred or advantageous over other implementations. Moreover, each example is provided by way of explanation of the disclosure, not limitation of the disclosure. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present disclosure without departing from the scope of the disclosure. Thus, it is intended that the present disclosure covers such modifications and variations as come within the scope of the appended claims.

As used herein, the term "first stream" or "free stream" refers to a stream that flows outside of the engine inlet and over a fan, which is unducted. Furthermore, the first stream is a stream of air that is free stream air. As used herein, the term "second stream" or "core stream" refers to a stream that flows through the engine inlet and the ducted fan and also travels through the core inlet and the core duct. As used herein, the term "third stream" or "mid-fan stream" refers to a stream that flows through an engine inlet and a ducted fan but does not travel through a core inlet and a core duct. Furthermore, the third stream is a stream of air that takes inlet air as opposed to free stream air. The third stream goes through at least one stage of the turbomachine, e.g., the ducted fan.

Thus, a third stream means a non-primary air stream capable of increasing fluid energy to produce a minority of total propulsion system thrust. A pressure ratio of the third stream is higher than that of the primary propulsion stream (e.g., a bypass or propeller driven propulsion stream). The thrust may be produced through a dedicated nozzle or through mixing of an airflow through the third stream with a primary propulsion stream or a core air stream, e.g., into a common nozzle.

In certain exemplary embodiments an operating temperature of the airflow through the third stream may be less than a maximum compressor discharge temperature for the engine, and more specifically may be less than <NUM> degrees Celsius (<NUM> degrees Fahrenheit) (such as less than <NUM> degrees Celsius (<NUM> degrees Fahrenheit), such as less than <NUM> degrees Celsius (<NUM> degrees Fahrenheit), such as less than <NUM> degrees Celsius (<NUM> degrees Fahrenheit), and at least as great as an ambient temperature). In certain exemplary embodiments, these operating temperatures may facilitate heat transfer to or from the airflow through the third stream and a separate fluid stream. Further, in certain exemplary embodiments, the airflow through the third stream may contribute less than <NUM>% of the total engine thrust (and at least, e.g., <NUM>% of the total engine thrust) at a takeoff condition, or more particularly while operating at a rated takeoff power at sea level, static flight speed, <NUM> degrees Fahrenheit ambient temperature operating conditions. In other exemplary embodiments, it is contemplated that the airflow through the third stream may contribute greater than <NUM>% of the total engine thrust (and at least, e.g., <NUM>% of the total engine thrust) at an engine operating condition. In other exemplary embodiments, it is contemplated that the airflow through the third stream may contribute approximately <NUM>% of the total engine thrust (and at least, e.g., <NUM>% of the total engine thrust) at an engine operating condition.

Furthermore in certain exemplary embodiments, aspects of the airflow through the third stream (e.g., airstream, mixing, or exhaust properties), and thereby the aforementioned exemplary percent contribution to total thrust, may passively adjust during engine operation or be modified purposefully through use of engine control features (such as fuel flow, electric machine power, variable stators, variable inlet guide vanes, valves, variable exhaust geometry, or fluidic features) to adjust or optimize overall system performance across a broad range of potential operating conditions.

Certain modem fan blades are formed of composite material(s) to reduce a weight of the fan blades. However, aircraft engine components, such as fan blades, nacelles, guide vanes, etc., used in jet engine applications are susceptible to foreign object impact damage or ingestion events, such as an ice ingestion or bird strike. Moreover, fan blades formed from composite material(s) may be more susceptible to damage in such events, e.g., by blade fracture, component delamination, bending or deformation damage, or other forms of blade damage. Accordingly, improved airfoil designs for addressing one or more of the above-mentioned problems would be useful. More specifically, an airfoil assembly with a lightweight and structurally sound design that can withstand foreign object ingestion events would be particularly beneficial.

As explained herein, composite fan blades that use internal foam support may be used in gas turbine engines. However, the foam may have a high risk of debonding from other portions of the blade. More specifically, under certain operational loads or during an ingestion event (e.g., ice ingestion or bird strike), the foam within a composite blade may shear or otherwise lose its bond with the spar, the outer blade skin, etc. Accordingly, aspects of the present subject matter are generally directed to a structurally reinforced foam positioned within the composite blade, e.g., within a void defined between a spar and an outer blade skin.

According to exemplary embodiments, the foam may be segmented and placed at different regions of the blade to meet the strength and stiffness requirements of the blade. The structural reinforcement may include a frame, grid, cross members, elongated supports, etc. that are formed from chopped fiber polymer matrix composites ("PMCs"), continuous PMCs, glass, sheet metal, etc. The structurally reinforced foam may be co-cured to generate a bond line between the structural spar and blade flowpath skins.

Such a composite blade construction may facilitate improved foam durability, thus enabling fan blade weight reduction while minimizing the potential for blade deformation, debonding, failure, or other operational degradation. In addition, local blade stiffnesses may be modified and tailored by selectively designing and positioning structural reinforcements within the foam. Moreover, such constructions may improve fan blade stability to meet aeromechanical requirements, may result in an improvement in dissipation of shock wave energy due to impact loads, may provided better control of blade untwist behavior to improve the operability margins, may improve fan blade durability, etc..

Referring now to <FIG>, a schematic cross-sectional view of a gas turbine engine <NUM> is provided according to an example embodiment of the present disclosure. Particularly, <FIG> provides an engine having a rotor assembly with a single stage of unducted rotor blades. In such a manner, the rotor assembly may be referred to herein as an "unducted fan," or the entire gas turbine engine <NUM> may be referred to as an "unducted engine," or an engine having an open rotor propulsion system <NUM>. In addition, the engine of <FIG> includes a mid-fan stream extending from the compressor section to a rotor assembly flowpath over the turbomachine, as will be explained in more detail below. It is also contemplated that, in other exemplary embodiments, the present disclosure is compatible with an engine having a duct around the unducted fan. It is also contemplated that, in other exemplary embodiments, the present disclosure is compatible with a turbofan engine having a third stream as described herein.

For reference, the gas turbine engine <NUM> defines an axial direction A, a radial direction R, and a circumferential direction C. Moreover, the gas turbine engine <NUM> defines an axial centerline or longitudinal axis <NUM> that extends along the axial direction A. In general, the axial direction A extends parallel to the longitudinal axis <NUM>, the radial direction R extends outward from and inward to the longitudinal axis <NUM> in a direction orthogonal to the axial direction A, and the circumferential direction extends three hundred sixty degrees (<NUM>°) around the longitudinal axis <NUM>. The gas turbine engine <NUM> extends between a forward end <NUM> and an aft end <NUM>, e.g., along the axial direction A.

The gas turbine engine <NUM> includes a turbomachine <NUM>, also referred to as a core of the gas turbine engine <NUM>, and a rotor assembly, also referred to as a fan section <NUM>, positioned upstream thereof. Generally, the turbomachine <NUM> includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. Particularly, as shown in <FIG>, the turbomachine <NUM> includes a core cowl <NUM> that defines an annular core inlet <NUM>. The core cowl <NUM> further encloses at least in part a low pressure system and a high pressure system. For example, the core cowl <NUM> depicted encloses and supports at least in part a booster or low pressure ("LP") compressor <NUM> for pressurizing the air that enters the turbomachine <NUM> through core inlet <NUM>. A high pressure ("HP"), multi-stage, axial-flow compressor <NUM> receives pressurized air from the LP compressor <NUM> and further increases the pressure of the air. The pressurized air stream flows downstream to a combustor <NUM> of the combustion section where fuel is injected into the pressurized air stream and ignited to raise the temperature and energy level of the pressurized air and produce high energy combustion products.

It will be appreciated that as used herein, the terms "high/low speed" and "high/low pressure" are used with respect to the high pressure/high speed system and low pressure/low speed system interchangeably. Further, it will be appreciated that the terms "high" and "low" are used in this same context to distinguish the two systems, and are not meant to imply any absolute speed and/or pressure values.

The high energy combustion products flow from the combustor <NUM> downstream to a high pressure turbine <NUM>. The high pressure turbine <NUM> drives the high pressure compressor <NUM> through a high pressure shaft <NUM>. In this regard, the high pressure turbine <NUM> is drivingly coupled with the high pressure compressor <NUM>. The high energy combustion products then flow to a low pressure turbine <NUM>. The low pressure turbine <NUM> drives the low pressure compressor <NUM> and components of the fan section <NUM> through a low pressure shaft <NUM>. In this regard, the low pressure turbine <NUM> is drivingly coupled with the low pressure compressor <NUM> and components of the fan section <NUM>. The LP shaft <NUM> is coaxial with the HP shaft <NUM> in this example embodiment. After driving each of the turbines <NUM>, <NUM>, the combustion products exit the turbomachine <NUM> through a core or turbomachine exhaust nozzle <NUM>.

Accordingly, the turbomachine <NUM> defines a working gas flowpath or core duct <NUM> that extends between the core inlet <NUM> and the turbomachine exhaust nozzle <NUM>. The core duct <NUM> is an annular duct positioned generally inward of the core cowl <NUM> along the radial direction R. The core duct <NUM> (e.g., the working gas flowpath through the turbomachine <NUM>) may be referred to as a second stream.

The fan section <NUM> includes a fan <NUM>, which is the primary fan in this example embodiment. For the depicted embodiment of <FIG>, the fan <NUM> is an open rotor or unducted fan <NUM>. As depicted, the fan <NUM> includes an array of fan blades <NUM> (only one shown in <FIG>). The fan blades <NUM> are rotatable, e.g., about the longitudinal axis <NUM>. As noted above, the fan <NUM> is drivingly coupled with the low pressure turbine <NUM> via the LP shaft <NUM>. The fan <NUM> can be directly coupled with the LP shaft <NUM>, e.g., in a direct-drive configuration. However, for the embodiments shown in <FIG>, the fan <NUM> is coupled with the LP shaft <NUM> via a speed reduction gearbox <NUM>, e.g., in an indirect-drive or geared-drive configuration.

Moreover, the fan blades <NUM> can be arranged in equal spacing around the longitudinal axis <NUM>. Each fan blade <NUM> has a root and a tip and a span defined therebetween. Each fan blade <NUM> defines a central blade axis <NUM>. For this embodiment, each fan blade <NUM> of the fan <NUM> is rotatable about their respective central blade axis <NUM>, e.g., in unison with one another. One or more actuators <NUM> are provided to facilitate such rotation and therefore may be used to change a pitch the fan blades <NUM> about their respective central blade axis <NUM>.

The fan section <NUM> further includes a fan guide vane array <NUM> that includes fan guide vanes <NUM> (only one shown in <FIG>) disposed around the longitudinal axis <NUM>. For this embodiment, the fan guide vanes <NUM> are not rotatable about the longitudinal axis <NUM>. Each fan guide vane <NUM> has a root and a tip and a span defined therebetween. The fan guide vanes <NUM> may be unshrouded as shown in <FIG> or, alternatively, may be shrouded, e.g., by an annular shroud spaced outward from the tips of the fan guide vanes <NUM> along the radial direction R or attached to the fan guide vanes <NUM>.

Each fan guide vane <NUM> defines a central blade axis <NUM>. For this embodiment, each fan guide vane <NUM> of the fan guide vane array <NUM> is rotatable about their respective central blade axis <NUM>, e.g., in unison with one another. One or more actuators <NUM> are provided to facilitate such rotation and therefore may be used to change a pitch of the fan guide vane <NUM> about their respective central blade axis <NUM>. However, in other embodiments, each fan guide vane <NUM> may be fixed or unable to be pitched about its central blade axis <NUM>. The fan guide vanes <NUM> are mounted to a fan cowl <NUM>.

As shown in <FIG>, in addition to the fan <NUM>, which is unducted, a ducted fan <NUM> is included aft of the fan <NUM>, such that the gas turbine engine <NUM> includes both a ducted and an unducted fan which both serve to generate thrust through the movement of air without passage through at least a portion of the turbomachine <NUM> (e.g., the HP compressor <NUM> and combustion section for the embodiment depicted). The ducted fan is shown at about the same axial location as the fan blade <NUM>, and radially inward of the fan blade <NUM>. The ducted fan <NUM>, for the embodiment depicted, is driven by the low pressure turbine <NUM> (e.g., coupled to the LP shaft <NUM>).

The fan cowl <NUM> annularly encases at least a portion of the core cowl <NUM> and is generally positioned outward of at least a portion of the core cowl <NUM> along the radial direction R. Particularly, a downstream section of the fan cowl <NUM> extends over a forward portion of the core cowl <NUM> to define a fan flowpath or fan duct <NUM>. The fan flowpath or fan duct <NUM> may be referred to as a third stream of the gas turbine engine <NUM>.

Incoming air may enter through the fan duct <NUM> through a fan duct inlet <NUM> and may exit through a fan exhaust nozzle <NUM> to produce propulsive thrust. The fan duct <NUM> is an annular duct positioned generally outward of the core duct <NUM> along the radial direction R. The fan cowl <NUM> and the core cowl <NUM> are connected together and supported by a plurality of substantially radially-extending, circumferentially-spaced stationary struts <NUM> (only one shown in <FIG>). The stationary struts <NUM> may each be aerodynamically contoured to direct air flowing thereby. Other struts in addition to the stationary struts <NUM> may be used to connect and support the fan cowl <NUM> and/or core cowl <NUM>. In many embodiments, the fan duct <NUM> and the core duct <NUM> may at least partially co-extend (generally axially) on opposite sides (e.g., opposite radial sides) of the core cowl <NUM>. For example, the fan duct <NUM> and the core duct <NUM> may each extend directly from a leading edge <NUM> of the core cowl <NUM> and may partially co-extend generally axially on opposite radial sides of the core cowl.

The gas turbine engine <NUM> also defines or includes an inlet duct <NUM>. The inlet duct <NUM> extends between an engine inlet <NUM> and the core inlet <NUM>/fan duct inlet <NUM>. The engine inlet <NUM> is defined generally at the forward end of the fan cowl <NUM> and is positioned between the fan <NUM> and the fan guide vane array <NUM> along the axial direction A. The inlet duct <NUM> is an annular duct that is positioned inward of the fan cowl <NUM> along the radial direction R. Air flowing downstream along the inlet duct <NUM> is split, not necessarily evenly, into the core duct <NUM> and the fan duct <NUM> by a splitter or leading edge <NUM> of the core cowl <NUM>. The inlet duct <NUM> is wider than the core duct <NUM> along the radial direction R. The inlet duct <NUM> is also wider than the fan duct <NUM> along the radial direction R.

Referring now generally to <FIG>, airfoil assemblies that may be used in a gas turbine engine will be described according to exemplary embodiments of the present subject matter. Specifically, <FIG> provide schematic illustrations of an airfoil assembly <NUM> that may be used in gas turbine engine <NUM>, e.g., as fan blade <NUM> or as fan guide vanes <NUM>. In addition, <FIG> provides another exemplary configuration of an airfoil assembly <NUM>, e.g., similar to that which may be used in a ducted turbofan engine. <FIG> provide exemplary schematic cross-sections of airfoil assemblies in accordance with exemplary embodiments of the present subject matter.

Notably, due to the similarity between embodiments described herein, like reference numerals may be used to refer to the same or similar features among various embodiments. Although airfoil assemblies <NUM> are described herein as being used with gas turbine engine <NUM>, it should be appreciated that aspects of the present subject matter may be applicable to any suitable blades for any suitable gas turbine engine. Indeed, the exemplary blade constructions and features described herein may be interchangeable among embodiments to generate additional exemplary embodiments. The specific structures illustrated and described herein are only exemplary and are not intended to limit the scope of the present subject matter in any manner.

Referring now specifically to <FIG>, airfoil assembly <NUM> includes a central spar <NUM> that extends outward along a radial direction R, e.g., which corresponds to radial direction R when airfoil assembly <NUM> is installed in gas turbine engine <NUM>. More specifically, as illustrated, central spar <NUM> may include a blade attachment structure <NUM>, e.g., illustrated as a dovetail, for securing airfoil assembly <NUM> to a rotating central hub (e.g., or mechanically coupling airfoil assemblies <NUM> to actuators <NUM>). Central spar <NUM> may generally define a root <NUM> of airfoil assembly <NUM> and may extend outward from root <NUM> along the radial direction R toward a tip <NUM> of airfoil assembly <NUM>. In general, central spar <NUM> may be formed from any suitably rigid material(s) that can withstand the forces exerted on airfoil assembly <NUM> during operation of the gas turbine engine <NUM>.

In addition, airfoil assembly <NUM> includes a blade skin <NUM> that is positioned or wrapped around central spar <NUM> to define an airfoil <NUM>. Blade skin <NUM> may be a polymer matrix composite (PMC), epoxy resin, carbon fiber, glass fiber, thermoplastics material, etc. As used herein, the terms "airfoil" and the like may generally refer to the shape or geometry of an outer surface of airfoil assembly <NUM>, e.g., the surface that interacts with the stream of air passing over airfoil assembly <NUM>. Airfoil <NUM> has a pressure side <NUM> and a suction side <NUM> extending in the axial direction A between a leading edge <NUM> (e.g., a forward end of airfoil <NUM>) and a trailing edge <NUM> (e.g., an aft end of airfoil <NUM>). In addition, a chord line <NUM> may be generally defined as a line extending between leading edge <NUM> and trailing edge <NUM>, and the term "chordwise direction" may generally refer to the relative position along chord line <NUM>. In addition, a span <NUM> of airfoil assembly <NUM> may be generally defined as the distance between root <NUM> and tip <NUM> of airfoil assembly <NUM> as measured along the radial direction R, and the term "spanwise direction" may generally refer to relative position along span <NUM>.

As illustrated, at least one cavity <NUM> is defined between blade skin <NUM> and central spar <NUM>. Airfoil assembly <NUM> further includes a support structure <NUM> that is positioned at least partially within the cavities <NUM>, e.g., to provide additional structural support and rigidity to airfoil assembly <NUM> without unnecessarily increasing a weight of airfoil assembly <NUM>. Notably, as explained above, filling cavities within an airfoil with only foam may provide a lightweight solution for improving the rigidity of the airfoil. However, foam filler often suffers from deformation, damage, or debonding during blade impact events, such as ice ingestion or bird strike. Accordingly, aspects of the present subject matter are directed toward an improved support structure <NUM> that is lightweight, provides improved structural rigidity to airfoil assembly <NUM>, and can withstand the forces associated high operational loads or ingestion events.

Specifically, referring again to the figures, support structure <NUM> includes a foam <NUM> and a foam reinforcement structure <NUM> that is embedded within foam <NUM>. In general, the foam reinforcement structure <NUM> embedded within the foam comprises a plurality of structural support members <NUM> for providing improved rigidity to foam <NUM>, support structure <NUM>, and airfoil assembly <NUM>. Although exemplary foam reinforcement structures <NUM> are described herein according to exemplary embodiments of the present subject matter, it should be appreciated that these are examples and are not intended to limit the scope of the present subject matter in any manner, which is defined by the appended claims.

According to exemplary embodiments, foam <NUM> may generally include at least one of polymethacrylimide (PMI) foam or a urethane foam. In addition, or alternatively, foam <NUM> may also include cast syntactic or expanding syntactic foams, e.g., glass, carbon, or phenolic micro balloons cast in resin. Other suitable foams are possible and within the scope of the present subject matter. Foam reinforcement structure <NUM> may generally be formed from any material suitable for improving the rigidity or durability of support structure <NUM>. For example, according to exemplary embodiments, foam reinforcement structure <NUM> may include at least one of a polymer matrix composite material (PMC), metallic reinforcements, carbon reinforcements, thermoplastics, or glass. In addition, as described in more detail below, foam reinforcement structure <NUM> may have a variety of geometries, such as the linear support members, honeycomb structures, cellular matrix structures, etc..

Foam reinforcement structure <NUM> may include any of the aforementioned materials in a unidirectional pre-preg, braided, and/or woven construction. By way of example, according to one embodiment, the PMC material is defined in part by prepreg, which is a reinforcement material pre-impregnated with a matrix material, such as thermoplastic resin desired for the matrix material. 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 (for example electrostatically) and then adhered to the fiber (for 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.

According to an alternative option, instead of using a prepreg, with the use of thermoplastic polymers it is possible to have a woven fabric as the foam reinforcement structure <NUM> that has, for example, dry carbon fiber 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 fiber, carbon fiber, and thermoplastic fiber could all be woven together in various concentrations to tailor the properties of the part. The carbon fiber provides the strength of the system, the glass may 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 are the matrix that will be flowed to bind the reinforcement fibers.

Many PMC materials are fabricated with the use of prepreg, which is a fabric or unidirectional tape that is impregnated with resin. Multiple layers of prepreg may be layered as needed to form the desired geometry for the part, e.g., one of the structural support members <NUM> or the foam reinforcement structure <NUM>, 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.

According to the exemplary embodiment illustrated in <FIG>, foam reinforcement structure <NUM> includes a plurality of structural support members <NUM> that extend through foam <NUM> and which have any suitable geometry, angle, and spacing for improving the rigidity are structural integrity of airfoil assembly <NUM>. According to the illustrated embodiment, each of the plurality of structural support members <NUM> is substantially straight or planar and extends between leading edge <NUM> of airfoil <NUM> to an upstream edge <NUM> of central spar <NUM>. Alternatively, structural support members <NUM> may additionally extend between a downstream edge <NUM> of central spar <NUM> and trailing edge <NUM> of airfoil <NUM>.

Referring still to <FIG>, cavity <NUM> may generally define a plurality of regions (e.g., identified herein generally by reference numeral <NUM>). It should be appreciated that the design of foam reinforcement structure <NUM> may vary from one region <NUM> to another, e.g., based on the anticipated forces experienced at those particular regions <NUM>. For example, the foam reinforcement structure <NUM> may vary in geometry, may be constructed of a different material, may have a different thickness or orientation, etc., depending on the location within airfoil assembly <NUM>.

For example, <FIG> illustrates airfoil assembly <NUM> with three regions <NUM> of foam reinforcement structure <NUM> (e.g., with other regions being omitted for clarity). According to the illustrated embodiment, each of the plurality of structural support members <NUM> may extend at a first angle <NUM> measured relative to the axial direction A. In this regard, as shown schematically in <FIG>, ingested objects (e.g., as identified by arrow <NUM>) are commonly moving along the axial direction A relative to airfoil assembly <NUM>, e.g., due to the air speed of gas turbine engine <NUM> upon ingestion. Notably, due to the rotational speed of airfoil <NUM>, the shape of airfoil <NUM>, and other factors, the forces exerted on airfoil assembly <NUM> may be known for each region <NUM> of airfoil assembly <NUM>. First angle <NUM> may be tailored at a specific position within airfoil assembly <NUM> for absorbing the most commonly experienced forces at that location.

According to exemplary embodiments of the present subject matter, first angle <NUM> may be <NUM>° (e.g., parallel to the axial direction A), may be normal to leading edge <NUM> of airfoil <NUM>, or may be any other suitable angle. Specifically, according to the illustrated embodiment, first angle <NUM> may be between about <NUM>° and <NUM>°, between about <NUM>° and <NUM>°, between about <NUM>° and <NUM>°, or about <NUM>°. In addition, although first angle <NUM> is illustrated as being substantially constant within each region <NUM>, it should be appreciated that first angle <NUM> may vary for each structural support member <NUM> (e.g., within each region <NUM>) while remaining within the scope of the present subject matter. In addition, it should be appreciated that according to alternative embodiments structural support members <NUM> may be non-parallel within a given region <NUM>, may have different cross-sectional profiles, etc..

Referring now specifically to <FIG>, structural support members <NUM> may additionally extend between pressure side <NUM> and suction side <NUM> of airfoil <NUM>. As illustrated, these structural support members <NUM> may define a second angle <NUM> is measured relative to pressure side <NUM> of airfoil <NUM>. According to an exemplary embodiment of the present subject matter (e.g., as illustrated for example in <FIG>), structural support member <NUM> may extend substantially normal to pressure side <NUM>, such that second angle <NUM> is substantially <NUM>°. By contrast, according to alternative embodiments, structural support members <NUM> may extend at any other suitable second angle <NUM>, such as between about between about <NUM>° and <NUM>°, between about <NUM>° and <NUM>°, between about <NUM>° and <NUM>°, or about <NUM>°. Other suitable second angles <NUM> may be used, e.g., depending on the anticipated forces that will be exerted on the pressure side <NUM> and suction side <NUM>, respectively.

As illustrated in the figures, foam reinforcement structure <NUM> may generally define a spacing <NUM> between adjacent structural support members <NUM>. In this regard, spacing <NUM> may be measured as the distance between adjacent structural support members <NUM> in a direction normal to such structural support members <NUM>. In general, spacing <NUM> may vary depending on the structural loading expected at a particular location on airfoil assembly <NUM>, e.g., with smaller spacing generally supporting higher loads. According to exemplary embodiments, spacing <NUM> within a particular region <NUM> may be between between <NUM> and <NUM> (<NUM> and <NUM> inches), between <NUM> and <NUM> (<NUM> and <NUM> inches), between <NUM> and <NUM> (<NUM> and <NUM> inches), or about <NUM> (<NUM> inch). In addition, each of the plurality of structural support members <NUM> may define a structure thickness <NUM> (see <FIG>). According to exemplary embodiments, structure thickness <NUM> is between <NUM> and <NUM> (<NUM> and <NUM> inches), between <NUM> and <NUM> (<NUM> and <NUM> inches), between <NUM> and <NUM> (<NUM> and <NUM> inches), or about <NUM> (<NUM> inches). Other spacings and structure sizes are possible and within the scope present subject matter.

Notably, it should be appreciated that foam reinforcement structure <NUM> may have any suitable geometry or structure. For example, as described above, foam reinforcement structure <NUM> includes a plurality of linear structural support members <NUM>. However, it should be appreciated that structural support members <NUM> may be curved, serpentine, or may have any other suitable size and/or geometry depending on the application. Moreover, referring now briefly to <FIG>, foam reinforcement structure <NUM> may further include a honeycomb structure (e.g., as identified generally by reference numeral <NUM>), a cellular matrix structure (e.g., as identified generally by reference numeral <NUM>), or any other suitable grid, lattice, or mesh-like structure. In addition, the specific geometry selected for foam reinforcement structure <NUM> may vary depending on the location within airfoil assembly <NUM>.

According to exemplary embodiments, each region <NUM> of airfoil <NUM> may include a foam reinforcement structure <NUM> that is similar to or different that other regions <NUM> of the same airfoil <NUM>. For example, one region <NUM> may include structural support members <NUM>, another may include a honeycomb structure <NUM>, and still another may include a cellular matrix. In addition, the spacing <NUM> and structure thickness <NUM> may vary among regions <NUM>. Indeed, it should be appreciated that any and all of the foam reinforcement structures <NUM> and their variations described herein may not be mutually exclusive and may be utilized in a single airfoil as desired depending on the application.

Referring now briefly to <FIG>, according to an exemplary embodiment of the present subject matter, structural support members <NUM> may include a flared end <NUM> that contacts blade skin <NUM>. In this regard, a thickness of each structural support member <NUM> may increase toward a contact point with blade skin <NUM> to provide improved physical connection or load distribution between the blade skin <NUM> and structural support member <NUM>. Other geometry variations are possible and within the scope of the present subject matter.

In addition, airfoil assembly <NUM> may include an adhesive <NUM> that is positioned on the inside of blade skin <NUM> and/or on central spar <NUM> for improving the structural engagement between portions of airfoil assembly <NUM>. In this regard, adhesive <NUM> may be positioned between support structure <NUM> and at least one of central spar <NUM> or blade skin <NUM>. Adhesives may include epoxy, polyurethane, or any other kind of adhesive known to those of ordinary skill in the art.

In addition, airfoil assembly <NUM> may include additional structural supports for improving the rigidity of airfoil assembly <NUM>. For example, airfoil assembly <NUM> may include a leading edge structural support <NUM> and a trailing edge structural support <NUM>. For example, as illustrated in <FIG>, leading edge structural support <NUM> may be positioned between blade skin <NUM> and support structure <NUM> proximate leading edge <NUM> of airfoil <NUM>. In addition, or alternatively, trailing edge structural support <NUM> may be positioned between blade skin <NUM> and support structure <NUM> proximate trailing edge <NUM> of airfoil <NUM>. Each of the leading edge structural support <NUM> and trailing edge structural support <NUM> may be spaced apart from central spar <NUM> along a chordwise direction to define a plurality of cavities <NUM> filled with support structure <NUM>. In addition, leading edge structural support <NUM> and trailing edge structural support <NUM> may be separate from central spar <NUM> or may be mechanically coupled to central spar <NUM>, e.g., proximate root <NUM> of airfoil <NUM>.

Referring now briefly to <FIG>, according to exemplary embodiments of the present subject matter, foam reinforcement structure <NUM> may further include one or more foam engagement structures <NUM> that are positioned on or extend from blade skin <NUM> and into cavities <NUM> when blade skin <NUM> is wrapped around central spar <NUM> and support structure <NUM>. For example, according to the illustrated embodiment, foam engagement structures <NUM> may include one or more protruding members <NUM> or tapered lugs <NUM>. As illustrated, airfoil <NUM> may generally define a blade thickness <NUM> measured normal to a chordwise direction. According to the illustrated embodiment, foam engagement structures <NUM> extend across only a portion of blade thickness <NUM>. For example, foam engagement structures <NUM> may extend less than half, less than a quarter, or less, through blade thickness <NUM>.

Referring now to <FIG>, an exemplary method <NUM> for constructing an airfoil assembly will be described according to exemplary embodiments of the present subject matter. For example, method <NUM> may be used to construct airfoil assembly <NUM> as described above. However, it should be appreciated that aspects of method <NUM> may be applied to the construction of any other suitable airfoil. In addition, it should be appreciated that alterations and modifications may be made to method <NUM> while remaining within scope of the present subject matter.

Method <NUM> includes, at step <NUM>, laying of a preform of foam segments separated by one or more foam reinforcement structures. In this regard, for example, a plurality of solid foam blocks may be formed which have the desired dimensions, e.g., such as a rectangular cross-section with a thickness equivalent to a target spacing between adjacent reinforcement structures. The foam reinforcement structures may be separately manufactured or acquired for providing additional structure support within the foam blocks. Specifically, the foam segments may be alternately stacked with the foam reinforcement structure to create a preform.

According to exemplary embodiments, an adhesive or bonding agent may be applied to one or both the foam segments and the foam reinforcement structures for improved engagement between the two. The adhesive may then be allowed to dry to generate the support structure preform. In addition, or alternatively, step <NUM> may include curing the preform to cure the adhesive, improve the bond between the foam segments and the foam reinforcement structure, and create a solid support structure that includes both foam segments and foam reinforcement structures.

Notably, steps <NUM> and <NUM> may result in a support structure preform that does not have a suitable shape to be the base of an airfoil. Accordingly, if the preform does not have the desired airfoil shape, it may be desirable to machine or manipulate the preform to have a profile suitable for forming an airfoil, e.g., upon assembly and skin wrapping. Accordingly, step <NUM> may include machining the preform to create a support structure having a predetermined core profile. In this regard, for example, the resulting support structure may have the shape of the cavities defined between a central spar and blade skin.

Step <NUM> includes laying up the support structure against a central spar. According to exemplary embodiments, an adhesive or bonding agent may be applied between the support structure and the central spar to create the internal structure of airfoil assembly. Step <NUM> includes wrapping the support structure and the central spar with a blade skin to form an airfoil assembly. Once again, an adhesive may be used on the surface of the blade skin that contacts the support structure and/or the central spar. Step <NUM> then includes curing the airfoil assembly to bond all components of airfoil assembly together.

According to alternative embodiments of the present subject matter, airfoil assemblies may be constructed by assembling the central spar, blade skin, and support structures prior to injecting a foam filler. In addition, it should be appreciated that the support structure and other portions of airfoil assemblies may be constructed in any suitable manner, e.g., such as via additive manufacturing or other methods. Other suitable methods for manufacturing airfoil assemblies as described herein are possible and within scope of the present subject matter.

<FIG> depicts steps performed in a particular order for purposes of illustration and discussion. Those of ordinary skill in the art, using the disclosures provided herein, will understand that the steps of any of the methods discussed herein can be adapted, rearranged, expanded, omitted, or modified in various ways without deviating from the scope of the present disclosure. Moreover, although aspects of method <NUM> are explained using airfoil assembly <NUM> as an example, it should be appreciated that this method may be applied to the construction of any other suitable airfoil for any other suitable application.

Claim 1:
An airfoil assembly (<NUM>) comprising:
a central spar (<NUM>) extending along a first direction (R);
a blade skin (<NUM>) positioned around the central spar (<NUM>) to define an airfoil (<NUM>) that has a pressure side (<NUM>) and a suction side (<NUM>) extending in a second direction (A), the second direction (A) being perpendicular to the first direction (R), the pressure side and the suction side extending between a leading edge (<NUM>) and a trailing edge (<NUM>), wherein at least one cavity (<NUM>) is defined between the blade skin (<NUM>) and the central spar (<NUM>); and
a support structure (<NUM>) positioned at least partially within the at least one cavity (<NUM>), the support structure (<NUM>) comprising:
a foam (<NUM>); and
a foam reinforcement structure (<NUM>) embedded within the foam (<NUM>)
characterised in that
the foam reinforcement structure (<NUM>) comprises a plurality of structural support members (<NUM>), and wherein at least one of the plurality of structural support members (<NUM>) extends between the leading edge (<NUM>) of the airfoil (<NUM>) and an upstream edge (<NUM>) of the central spar (<NUM>) or between a downstream edge (<NUM>) of the central spar (<NUM>) and the trailing edge (<NUM>) of the airfoil (<NUM>).