Component for a turbine engine having a spar assembly

Abstract

A component and blade assembly for a turbine engine, the component having a wall defining an interior and defining an outer surface, the outer surface extending radially between a leading edge and a trailing edge to define a chordwise direction, and extending between a hub and a tip to define a spanwise direction. A metal spar extends from the hub into the interior in the spanwise direction, the metal spar defining a socket. A composite spar extends in the spanwise direction between a spar root and a spar tip, the spar root being received in the socket. The component also has a reinforcement.

Description

Technical Field

[0001] The present disclosure relates generally to components for turbine engines and, more particularly, to a blade assembly having a spar assembly. Background Art

[0002] Composite materials generally 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 used in various applications, such as turbine engines or aircraft. Composite materials can, for example, be mounted on a fuselage and / or a wing, a rudder, a manifold, an airfoil, or other components of an aircraft or a turbine engine or define a portion of a fuselage and / or a wing, a rudder, a manifold, an airfoil, or other components of an aircraft or a turbine engine. Extreme loads or sudden forces may be applied to composite components of an aircraft or a turbine engine. For example, during the ingestion of various materials by a turbine engine, extreme loads may occur to one or more airfoils. BRIEF DESCRIPTION OF THE DRAWINGS

[0003] A full and enabling disclosure of the present disclosure, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings, in which:

[0004] Figure 1 is a schematic cross-sectional view of a non-ducted or open rotor turbine engine.

[0005] Figure 2 Yes Figure 1 A schematic perspective view of an aircraft having a non-ducted or open rotor turbine engine.

[0006] Figure 3 is a schematic diagram of an engine component in the form of a blade assembly having a spar assembly.

[0007] Figure 4 According to the aspects disclosed herein, Figure 3 A perspective view of a variation of a wing spar assembly.

[0008] Figure 5 According to another aspect disclosed herein, Figure 3 A perspective view of another variation of a wing spar assembly.

[0009] Figure 6 According to another aspect disclosed herein, Figure 3 A perspective view of yet another variation of a wing spar assembly.

[0010] Figure 7 According to another aspect disclosed herein, Figure 3 A perspective view of another variation of a wing spar assembly. DETAILED DESCRIPTION

[0011] Traditionally, an airfoil includes a metal spar formed with or coupled to a hub of the airfoil.

[0012] Aspects disclosed herein relate to a component for a turbine engine having an airfoil having a metal spar, a composite spar, and a reinforcement that together define a spar assembly. The metal spar may be shaped to receive a portion of the composite spar. The reinforcement is bonded to at least one of the airfoil, the metal spar, or the composite spar.

[0013] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter references to refer to features in the drawings. Like or similar reference numbers in the drawings and description have been used to refer to like or similar parts of the present disclosure.

[0014] As used herein, the term "composite material" refers to a component having two or more materials. A composite material may be a combination of at least two or more metals, non-metals, or a combination of metal and non-metal elements or materials. Examples of composite materials may be, but are not limited to, polymer matrix composites (PMC), ceramic matrix composites (CMC), metal matrix composites (MMC), carbon fibers, polymer resins, thermoplastics, bismaleimide (BMI), polyimide materials, epoxy resins, glass fibers, and silicon matrix materials.

[0015] As used herein, a "composite" component refers to a structure or component that includes any suitable composite material. A composite component, such as a composite airfoil, may include multiple layers or plies of composite material. The stiffness, material, and dimensions of the layers or plies may be varied to achieve a desired composite component or composite portion of a component having a predetermined weight, size, stiffness, and strength.

[0016] One or more layers of adhesive may be used to form or join the composite parts. The adhesive may include resins and phenolics, where the adhesive may require curing at elevated temperatures or other hardening techniques.

[0017] As used herein, PMC refers to a class of materials. As an example, PMC materials are defined in part by prepregs, which are reinforcing materials pre-impregnated with polymer matrix materials (e.g., thermoplastic resins). Non-limiting examples of processes for producing thermoplastic prepregs include: hot melt prepregs, in which fiber reinforcements are pulled through a molten bath of resin; and powder prepregs, in which resins are electrostatically deposited onto fiber reinforcements by non-limiting examples, and then adhered to fibers in an oven or with the help of heated rollers by non-limiting examples. Prepregs can be in the form of unidirectional tapes or woven fabrics, which are then stacked on top of each other to form the desired number of stacked layers for a part.

[0018] Multilayer prepreg is stacked to a thickness and orientation suitable for composite parts, and then the resin is cured and solidified to provide fiber reinforced composite parts. Resins used for PMC matrix materials can generally be divided into thermosetting plastics or thermoplastics. Thermoplastic resins are generally classified as polymers that can be repeatedly softened and flowed when heated and hardened due to physical changes rather than chemical changes when fully cooled. Famous examples of thermoplastic resins include nylon, thermoplastic polyesters, polyaryletherketones and polycarbonate resins. Specific examples of high-performance thermoplastic resins that have been envisioned for 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, thermosetting resins do not undergo significant softening when heated, but instead undergo thermal decomposition when fully heated. Famous examples of thermosetting resins include epoxy resins, bismaleimide (BMI) and polyimide resins.

[0019] Instead of using prepreg, in another non-limiting example, by using thermoplastic polymers, woven fabrics can be utilized. Woven fabrics may include, but are not limited to, dry carbon fibers woven together with thermoplastic polymer fibers or filaments. Non-prepreg weaving architectures can be made in a similar manner. With this approach, the fiber volume of a part can be customized by specifying the relative concentrations of thermoplastic fibers and reinforcing fibers that are woven or woven together. In addition, different types of reinforcing fibers can be woven or woven together at different concentrations to customize the properties of a part. For example, glass fibers, carbon fibers, and thermoplastic fibers can all be woven together at different concentrations to customize the properties of a part. Carbon fibers provide the strength of the system, glass fibers can be incorporated to enhance impact properties, which is a design feature of a part located near the engine inlet, and thermoplastic fibers provide bonding for the reinforcing fibers.

[0020] As used herein, CMC refers to a class of materials having reinforcing fibers in a ceramic matrix. Typically, the reinforcing fibers provide structural integrity to the ceramic matrix. Some examples of reinforcing fibers may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), non-oxide carbon-based materials (e.g., carbon), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (Al 2 O 3 ), silicon dioxide (SiO 2 ), aluminosilicates (such as mullite), or mixtures thereof), or mixtures thereof.

[0021] Some examples of ceramic matrix materials may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (Al 2 O 3), silicon dioxide (SiO 2 ), aluminosilicates or mixtures thereof), or mixtures thereof. Optionally, ceramic particles (e.g., oxides of Si, Al, Zr, Y and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite and montmorillonite) may also be included in the ceramic matrix.

[0022] In general, specific CMCs can be referred to as their fiber type / matrix type combination. For example, C / SiC is carbon fiber reinforced silicon carbide, SiC / SiC is silicon carbide fiber reinforced silicon carbide, SiC / SiN is silicon carbide fiber reinforced silicon nitride, SiC / SiC-SiN is silicon carbide fiber reinforced silicon carbide / silicon nitride matrix mixture, etc. In other examples, CMCs can be made of materials containing oxide-based materials (such as aluminum oxide (Al2O3)). 2 O 3 ), silicon dioxide (SiO 2 ), aluminosilicates and mixtures thereof) and a matrix and reinforcing fibers. Aluminosilicates may include crystalline materials (e.g., mullite (3Al 2 O 3 2SiO 2 )), and glassy aluminosilicates.

[0023] In certain non-limiting examples, the reinforcing fibers may be bound and / or coated prior to inclusion in the ceramic matrix. For example, the fiber bundles may be formed into reinforcing tapes, such as unidirectional reinforcing tapes. Multiple tapes may be stacked together to form a preform component. The fiber bundles may be impregnated with the slurry composition prior to forming the preform or after forming the preform. The preform may then be thermally treated (e.g., cured or burned out) to produce a high coke residue in the preform, and subsequently chemically treated (e.g., infiltrated with silicon melt) to obtain a component formed of a CMC material having a desired chemical composition.

[0024] Such materials, along with certain monolithic ceramics (i.e., ceramic materials without reinforcement materials), are particularly suitable for higher temperature applications. In addition, these ceramic materials are lightweight compared to superalloys, but still provide strength and durability to the parts made from them. As a result, such materials are currently being considered for many gas turbine components used in the higher temperature sections of gas turbine engines, such as airfoils (e.g., turbines and buckets), combustors, shrouds, and other components that would benefit from the lighter weight and higher temperature capabilities that these materials can provide.

[0025] The term "metallic" as used herein refers to materials including metals such as, but not limited to, titanium, iron, aluminum, stainless steel, and nickel alloys. A metallic material or alloy may be a combination of at least two or more elements or materials, at least one of which is a metal.

[0026] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. In addition, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

[0027] As used herein, the terms “first” and “second” may be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0028] As used herein, the term "upstream" refers to the direction opposite to the direction of fluid flow, while the term "downstream" refers to the direction in the same direction as the direction of fluid flow. The terms "front" or "front" mean in front of something, and "rear" or "rearward" mean behind something. For example, for a gas turbine engine, front refers to a location closer to the engine inlet, while rear refers to a location closer to the engine nozzle or exhaust.

[0029] The term "fluid" may be a gas or a liquid, or multiple phases. The term "fluid communication" means that a fluid is able to establish a connection between specified areas.

[0030] Furthermore, as used herein, the term "radial" or "radially" refers to directions away from a common center. For example, in the overall context of a turbine engine, radial refers to directions along a ray extending between a central longitudinal axis of the engine and an outer engine circumference.

[0031] All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, rear, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, backward, etc.) are only used for identification purposes to help the reader understand the present disclosure, and do not create limitations, especially limitations on the position, orientation or use of the aspects of the present disclosure described herein. Unless otherwise stated, connection references (e.g., attachment, connection, connection and engagement) will be interpreted broadly and may include intermediate structural elements between element sets and relative movement between elements. Therefore, connection references do not necessarily infer that two elements are directly connected and have a fixed relationship to each other. The exemplary drawings are only for illustrative purposes, and the dimensions, positions, orders and relative sizes reflected in the attached drawings may vary.

[0032] Unless the context clearly dictates otherwise, the singular forms "a," "an," and "the" include plural referents. Also, as used herein, the term "set" or a "set of" elements may include any number of elements, including only one.

[0033] Here and throughout the specification and claims, range limitations are combined and interchanged, and unless context or language indicates otherwise, such ranges are identified and include all sub-ranges contained therein. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

[0034] In certain aspects of the present disclosure, an unducted or open rotor turbine engine includes a set of circumferentially spaced fan blades that extend outwardly beyond a nacelle surrounding an engine core.

[0035] Figure 1 is a schematic cross-sectional view of a turbine engine, particularly an open rotor or non-ducted turbine engine 10 for an aircraft. The non-ducted turbine engine 10 has a generally longitudinally extending axis or engine centerline 12 extending from a forward end 14 to an aft end 16. The non-ducted turbine engine 10 includes, in downstream serial flow relationship, a set of circumferentially spaced blades or propellers defining: a fan section 18 including a fan 20; a compressor section 22 including a supercharger or low pressure (LP) compressor 24 and a high pressure (HP) compressor 26; a combustion section 28 including a combustor 30; a turbine section 32 including an HP turbine 34 and an LP turbine 36; and an exhaust section 38. The non-ducted turbine engine 10 as described herein is by way of 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.

[0036] The outer surface defined by the casing or nacelle 40 of the non-ducted turbine engine 10 extends from the front end 14 of the non-ducted turbine engine 10 toward the rear end 16 of the non-ducted turbine engine 10 and covers at least a portion of the compressor section 22, the combustion section 28, the turbine section 32, and the exhaust section 38. The fan section 18 may be positioned at the front of the nacelle 40 and extends radially outward from the nacelle 40 of the non-ducted turbine engine 10. Specifically, the fan section 18 extends radially outward from the nacelle 40. The fan section 18 includes a set of fan blades 42 and a set of stationary fan blades 82 downstream of the set of fan blades 42, both of which are radially arranged around the engine centerline 12. The non-ducted turbine engine 10 includes any number of one or more sets of rotating blades or propellers (e.g., the set of fan blades 42) disposed upstream of the set of stationary fan blades 82. As a non-limiting example, the non-ducted turbine engine 10 may include multiple sets of fan blades 42 or fan blades 82. Therefore, the non-ducted turbine engine 10 is further defined as a non-ducted single fan turbine engine. The non-ducted turbine engine 10 is further defined by the position of the fan section 18 relative to the combustion section 28. The fan section 18 may be upstream, downstream, or in line with the axial location of the combustion section 28.

[0037] The compressor section 22 , the combustion section 28 , and the turbine section 32 are collectively referred to as an engine core 44 , which generates combustion gases. The engine core 44 is surrounded by an engine casing 46 operably coupled to a portion of a nacelle 40 of the non-ducted turbine engine 10 .

[0038] An HP shaft or spool 48 disposed coaxially about the engine centerline 12 of the non-ducted turbine engine 10 drivingly connects the HP turbine 34 to the HP compressor 26. An LP shaft or spool 50 disposed coaxially about the engine centerline 12 of the non-ducted turbine engine 10 within the larger diameter annular HP spool 48 drivingly connects the LP turbine 36 to the LP compressor 24 and the fan 20. The spools 48, 50 are rotatable about the engine centerline 12 and are coupled to a set of rotatable elements that collectively define a rotor 51.

[0039] It should be appreciated that the non-ducted turbine engine 10 is a direct drive or integral drive engine utilizing a reduction gearbox coupling the LP shaft or spool 50 to the fan 20 .

[0040] The LP compressor 24 and the HP compressor 26 each include a set of compressor stages 52, 54 in which a set of compressor blades 56, 58 rotate relative to a corresponding set of static compressor vanes 60, 62 (also referred to as nozzles) to compress or pressurize the fluid flow through the stage. In a single compressor stage 52, 54, a plurality of compressor blades 56, 58 are arranged in a ring and extend radially outward from a blade platform to a blade tip relative to the engine centerline 12, while the corresponding static compressor vanes 60, 62 are positioned upstream of and adjacent to the compressor blades 56, 58. Notably, Figure 1 The numbers of blades, buckets, and compressor stages shown in FIG. 5 are selected for illustration purposes only, and other numbers are possible.

[0041] Compressor blades 56, 58 for a compressor stage are mounted to disks 61 mounted to respective ones of the HP spool 48 and the LP spool 50, with each stage having its own disk 61. Static compressor vanes 60, 62 for a compressor stage are mounted to the engine case 46 in a circumferential arrangement.

[0042] The HP turbine 34 and the LP turbine 36 each include a set of turbine stages 64, 66 in which a set of turbine blades 68, 70 rotate relative to a corresponding set of static turbine vanes 72, 74 (also referred to as nozzles) to extract energy from the fluid flow passing through the stage. In a single turbine stage 64, 66, a plurality of turbine blades 68, 70 are arranged in a ring and extend radially outward from the blade platform to the blade tip relative to the engine centerline 12, while the corresponding static turbine vanes 72, 74 are positioned upstream of and adjacent to the turbine blades 68, 70. Notably, Figure 1 The number of blades, buckets, and turbine stages shown in FIG. 5 are selected for illustration purposes only, and other numbers are possible.

[0043] Turbine blades 68, 70 for the stages of turbine section 32 are mounted to disks 71 mounted to respective ones of HP spool 48 and LP spool 50, with each stage having a dedicated disk 71. Static turbine buckets 72, 74 for the stages of turbine section 32 are mounted to engine case 46 in a circumferential arrangement.

[0044] The rotating portions of the non-ducted turbine engine 10, such as the blades 56, 58, 68, 70 in the compressor section 22 and the turbine section 32, are also individually or collectively referred to as rotors 51. Thus, a rotor refers to the combination of rotating elements throughout the non-ducted turbine engine 10.

[0045] In addition to the rotor portion, the stationary portion of the non-ducted turbine engine 10 (such as the static blades 60, 62, 72, 74 in the compressor section 22 and the turbine section 32) are also individually or collectively referred to as stators 63. Therefore, the stator 63 refers to the combination of non-rotating elements throughout the non-ducted turbine engine 10.

[0046] The nacelle 40 is operably coupled to the non-ducted turbine engine 10 and covers at least a portion of the engine core 44, the engine case 46, or the exhaust section 38. At least a portion of the nacelle 40 extends axially forward or upstream of the position shown. For example, the nacelle 40 extends axially forward so that a portion of the nacelle 40 covers or covers a portion of the fan section 18 or the supercharger section (not shown) of the non-ducted turbine engine 10.

[0047] During operation of the non-ducted turbine engine 10, a free stream airflow 79 flows toward the front of the non-ducted turbine engine 10. A portion of the free stream airflow 79 enters the annular region 25 defined by the outer surface of the nacelle and the swept area between the tips of the blades, where the airflow is the inlet airflow 78. A portion of the inlet airflow 78 enters the engine core 44 and is described as the working airflow 76, which is used for combustion within the engine core 44.

[0048] More specifically, the working airflow 76 flows into the LP compressor 24, which then pressurizes the working airflow 76, thereby defining a pressurized airflow that is supplied to the HP compressor 26, which further pressurizes the air. The working airflow 76 or pressurized airflow from the HP compressor 26 is mixed with the fuel in the combustor 30 and ignited, thereby generating combustion gases. The HP turbine 34 extracts some work from these gases, which drives the HP compressor 26. The combustion gases are discharged into the LP turbine 36, which extracts additional work to drive the LP compressor 24, and the working airflow 76 or exhaust gases are ultimately discharged from the non-ducted turbine engine 10 via the exhaust section 38. The drive of the LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP compressor 24. The working airflow 76, including the pressurized airflow and the combustion gases, defines the working airflow flowing through the compressor section 22, the combustion section 28, and the turbine section 32 of the non-ducted turbine engine 10.

[0049] Inlet airflow 78 flows through the set of fan blades 42 and flows over the nacelle 40 of the non-ducted turbine engine 10. The inlet airflow 78 then flows over at least a portion of the set of stationary fan blades 82, which direct the inlet airflow 78 so that it is transverse toward the engine centerline 12. The inlet airflow 78 then follows the curve of the nacelle 40 and flows through the set of stationary fan blades 82 toward the exhaust section 38. The pylon 84 mounts the non-ducted turbine engine 10 to an external structure (e.g., a fuselage, wing, tail, etc. of an aircraft).

[0050] Working airflow 76 and at least some inlet airflow 78 merge downstream of exhaust section 38 of non-ducted turbine engine 10. Working airflow 76 and inlet airflow 78 together form the overall thrust of non-ducted turbine engine 10.

[0051] It is envisioned that a portion of the working airflow 76 is extracted as bleed air 77 (e.g., from the compressor section 22). The bleed air 77 provides airflow to engine components that require cooling. The temperature of the working airflow 76 leaving the combustor 30 is significantly increased relative to the working airflow 76 within the compressor section 22. Therefore, the cooling provided by the bleed air 77 is necessary for operating these engine components in an elevated temperature environment or in the hot section of the non-ducted turbine engine 10. In the context of a turbine engine, the hot section of the engine is typically downstream of the combustor 30, particularly downstream of the turbine section 32, with the HP turbine 34 being the hottest section because it is directly downstream of the combustion section 28. Other sources of cooling fluid are, but are not limited to, fluid discharged from the LP compressor 24 or the HP compressor 26.

[0052] Figure 2 is suitable for use as Figure 1A schematic perspective view of an aircraft 100 of a generic non-ducted turbine engine 102 of a non-ducted turbine engine 10 is shown. Aircraft 100 includes a fuselage 104 having an outer surface. At least one wing 106 and an empennage 108 extend from fuselage 104. Emitter 108 is operably coupled to and spaced apart from fuselage 104 via an empennage pylon 110. Non-ducted turbine engine 102 is operably coupled to the outer surface of fuselage 104 via pylon 112. Non-ducted turbine engine 102 includes a set of circumferentially spaced fan blades 116. A set of stationary fan blades 120 are disposed downstream of the set of circumferentially spaced fan blades 116. Fuselage 104 extends between a nose 122 and a tail 124 and includes a fuselage centerline 126 extending therebetween.

[0053] Additionally, while the empennage 108 is a T-tail empennage (e.g., as shown in the figure), other conventional empennages are contemplated, such as a cruciform empennage, an H-tail empennage, a triple empennage, a V-tail empennage, an inverted empennage, a Y-tail empennage, a twin empennage, a cantilever empennage, or a ring empennage, all of which are referred to herein as the empennage 108 .

[0054] Figure 3 1 is a schematic diagram of an engine component in the form of a blade assembly 130 as a non-limiting example. The blade assembly 130 includes an airfoil 131, which is shown as a composite blade as an example. As non-limiting examples, the airfoil 131 may be a blade in a set of fan blades 42, 116, or a blade from a compressor blade 56, 58 or a turbine blade 68, 70. In addition, the engine component may be a bucket assembly, wherein the airfoil 131 is a blade in the set of stationary fan blades 82, 120, or a blade in a static blade 60, 62, 72, 74. It is contemplated that the airfoil 131 may be a blade, bucket, 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.

[0055] Airfoil 131 includes a wall 132 defining an interior 133. Wall 132 defines an exterior surface 134 that extends radially between a leading edge 135 and a trailing edge 136 to define a chordwise direction (denoted as "C"). Exterior surface 134 may further extend between a root 137 and a tip 138 to define a spanwise direction (denoted as "S"). Wall 132 may be a composite wall made of one or more layers of material. The one or more layers of material may be applied at the same stage or at different stages of manufacturing airfoil 131.

[0056] As a non-limiting example, the wall 132 may include at least a polymer matrix composite (PMC) portion or a polymer portion. The polymer matrix composite may include, but is not limited to, a matrix of thermosetting (epoxy, phenolic) or thermoplastic (polycarbonate, polyvinyl chloride, nylon, acrylic) and embedded with glass, carbon, steel or Kevlar fibers.

[0057] The blade assembly 130 further includes a spar assembly 140. The spar assembly 140 includes, but is not limited to, a metal spar 141, a composite spar 142, and a reinforcement 143. The reinforcement 143 may be at least partially formed of metal. It is further contemplated that the reinforcement 143 is metal. The reinforcement 143 defines a wing portion 161 of the spar assembly 140. The spar assembly 140 defines a portion of the wing 161 extending from the engine centerline 12 ( Figure 1 ) extends radially and perpendicular to the engine centerline 12 ( Figure 1 ). The airfoil 131 is mounted to the spar assembly 140 near the root 137 via a metal trunnion 144 defining a hub 145. The airfoil 131 has a span length (denoted as "L") measured along the span direction S from the hub 145 at 0% of the span length L to the tip 138 at 100% of the span length L. The span length L may be parallel to the axis A and defined as the maximum distance between the root 137 and the tip 138 of the airfoil 131. The entire spar assembly 140 may be located less than 20% of the span length L. The wall 132 may surround and / or surround at least a portion of the spar assembly 140. At least a portion of the spar assembly 140 may be bonded to the wall 132.

[0058] The metal spar 141 may be formed of a metal such as, but not limited to, titanium, iron, aluminum, stainless steel, and nickel alloys. At least a portion of the metal spar 141 may be located within the interior 133. The metal spar 141 may be integral with the metal trunnion 144. The metal spar 141 may be located above the hub 145 and within the interior 133. The remainder of the metal trunnion 144 may be located below the airfoil 131.

[0059] The composite spar 142 may be formed of a polymeric material or other non-metallic material. At least a portion of the composite spar 142 may be located within the interior 133. The composite spar 142 may extend between a spar root 146 and a spar tip 147 in the span direction S. Most of the composite spar 142 including the spar tip 147 may be located within the interior 133.

[0060] The reinforcement 143 can be formed of a metal such as, but not limited to, titanium, iron, aluminum, stainless steel, and nickel alloys. In one aspect, the reinforcement 143 can be located within the interior 133. In another aspect, the reinforcement 143 can be located on the exterior surface 134 of the wall 132. It is further contemplated that a portion of the reinforcement 143 can be located within the interior 133 and other portions can be located on the exterior surface 134 of the wall 132.

[0061] It is also contemplated that one or more layers of adhesive (not shown) may be applied between the wall 132 and any portion of the spar assembly 140. Further, it is contemplated that the adhesive may be absorbed by one or more portions of the wall 132 and / or the spar assembly 140. The adhesive may include resins and phenolics, where the adhesive may need to be cured at elevated temperatures or other hardening techniques. It should be understood that any portion of the spar assembly 140 may be located within the interior 133 and / or on the exterior surface 134 and bonded accordingly at these locations.

[0062] Steering Figure 4 , a perspective view of a spar assembly 240 according to aspects disclosed herein is shown. Spar assembly 240 is similar to spar assembly 140, and thus, like parts of spar assembly 240 will be identified with like numbers increased by 100, it being understood that the description of like parts of spar assembly 140 applies to spar assembly 240 unless otherwise noted.

[0063] The metal spar 241 may extend from a base 250 of the metal trunnion 244 to define a hub 245 of the spar assembly 240. The metal spar 241 includes a set of walls 251, each of which defines an inner surface that is spaced a first distance (indicated as "D1") from each other, thereby defining a socket 252 shown in phantom. The socket 252 is closed on at least two sides 257 by the set of walls 251. The socket 252 is open on two other opposite sides 255 to define side openings 254. The socket 252 may have a socket length S L , wherein the socket 252 has a socket length S L multiplied by the cross-sectional area (denoted as "CA") of the first distance D1 (CA = S L x D1). The cross-sectional area CA can range from 1.0 in. 2 and 15in 2 (6.4516cm 2 and 96.774cm 2 )between.

[0064] The composite spar 242 is received in the socket 252. The composite spar 242 may have a first thickness (denoted as "T1") that is less than the first distance D1. It is further contemplated that the first thickness T1 is nearly equal to the first distance D1 to provide a tight fit of the composite spar 242 in the socket 252. The composite spar 242 may include a body portion 260 and a wing portion 261. The wing portion 261 and the body portion 260 form an inverted "T" shape. The body portion 260 may extend from the spar root 246 within the socket 252 in the span direction S. The wing portion 261 may extend from the body portion 260 out of the side opening 254 in a direction substantially perpendicular to the span direction S.

[0065] The reinforcement 243 is bonded to at least a portion of the composite spar 242. The reinforcement 243 may be bonded to the outer surface of the wing portion 262 on one or both sides of the composite spar 242. In one non-limiting example, the reinforcement 243 is bonded to the wing portion 261 of the composite spar 242. A single reinforcement 243c may extend through the socket 252 along the composite spar 242 and overlap the metal spar 241. It is further contemplated that the reinforcement 243 may be a plurality of pieces 243a, 243b on the wing portion 261 outside the socket 252. In this example, the reinforcement 243 does not extend through the socket 252 or overlap the metal spar 241. The reinforcement 243 may have a second thickness (denoted as "T2"). Further, the reinforcement 243 includes at least one tapered edge 263 that tapers from the second thickness T2 toward the composite spar 242. In one non-limiting example, the reinforcement 243 has all tapered edges 263. In aspects where the reinforcement is a single reinforcement 243c, the first thickness T1 and the second thickness T2 together are nearly equal to the first distance D1 (D1≈T1+T2) to provide the tight fit previously described herein. Although shown as having uniform thicknesses T1, T2, it is contemplated that the thicknesses described herein may vary.

[0066] In one aspect, one or more layers of adhesive (not shown) can be applied between the stiffener 243 and the wing portion 261 of the composite spar 242. Additionally or alternatively, one or more layers of adhesive (not shown) can be applied between the metal spar 241 and the stiffener 243 or the composite spar 242, or between the metal spar 241 and the stiffener 243 and the composite spar 242.

[0067] Steering Figure 5 , a spar assembly 340 according to another aspect disclosed herein is shown. Spar assembly 340 is similar to spar assembly 240, and thus, like parts of spar assembly 340 will be identified with like numbers increased by 100, it being understood that the description of like parts of spar assembly 240 applies to spar assembly 340 unless otherwise noted.

[0068] The metal spar 341 may extend from a base 350 of the metal trunnion 344 to define a hub 345 of the spar assembly 340. The metal spar 341 may include a set of walls 351 spaced apart from one another to define a socket 352. The socket 352 is closed on three sides defined by the set of walls 351. The socket 352 is open on the remaining sides to define a side opening 354. In other words, the socket 352 has a lateral "U" shape 353. A through hole 356 is positioned opposite the side opening 354 in the set of walls 351 at the hub 345. At least a portion of the through hole 356 is formed in the base 350.

[0069] The composite spar 342 is received in the socket 352. The composite spar 342 includes a body portion 360 (shown in dotted lines) and a wing portion 361 (shown in solid lines). The wing portion 361 and the body portion 360 form an inverted "T" shape. The body portion 360 may extend from the spar root 346 in the socket 352 in the span direction S. The wing portion 361 may extend a side opening 354 from the body portion 360 on one side of the body portion 360 in a direction substantially perpendicular to the span direction S, and extend a through hole 356 from the body portion 360 on the other side of the body portion 360.

[0070] The reinforcement 343 is bonded to at least a portion of the composite spar 342. The reinforcement 343 may be bonded to the outer surface of the wing portion 262 on one or both sides of the composite spar 342. In one non-limiting example, the reinforcement 343 is bonded to the wing portion 361 of the composite spar 342. A single reinforcement 343c may extend along the composite spar 342 through the socket 352 and overlap the metal spar 341. It is further contemplated that the reinforcement 343 may be multiple pieces 343a, 343b located on the wing portion 361 outside the socket 352.

[0071] Steering Figure 6 , a spar assembly 440 is shown according to another aspect disclosed herein. Spar assembly 440 is similar to spar assembly 340, and thus, like parts of spar assembly 440 will be identified with like numbers increased by 100, it being understood that the description of like parts of spar assembly 340 applies to spar assembly 440 unless otherwise noted.

[0072] The metal spar 441 may extend from the base 450 of the metal trunnion 444 to define a hub 445 of the spar assembly 440. The metal spar 441 includes a set of walls 451 spaced apart from one another to define a socket 452. The metal spar 441 includes a middle truss 462 that divides the socket 452 into two openings 454. In other words, the socket 452 has a lateral "H" shape 453. Although shown as being in the middle, the middle truss 462 may be located closer to either opening 454 and need not be located directly in the middle.

[0073] The composite spar 442 is received in the socket 452. The composite spar 442 includes a body portion 460 (shown in phantom) and a wing portion 461 (shown in solid lines). The wing portion 461 and the body portion 460 form an inverted split "T" shape, wherein a slot 463 is formed in the composite spar 442 to accommodate the intermediate truss 462 when assembled. The body portion 460 may extend from the spar root 446 in the socket 452 in the span direction S. The wing portion 461 may extend from the body portion 460 out of the opening 454 in a direction substantially perpendicular to the span direction S.

[0074] At least one reinforcement member 443 is bonded to at least a portion of the composite spar 442. In one non-limiting example, the reinforcement member 443 is bonded to the wing portion 461 of the composite spar 442 on both sides, as shown. The at least one reinforcement member 443 may also be a plurality of parts 443a, 443b located on the wing portion 461 outside the socket 452. It is further contemplated that the at least one reinforcement member 443 is an extended reinforcement member 443c that extends at least partially into the socket 452.

[0075] Steering Figure 7 , a spar assembly 540 is shown according to another aspect disclosed herein. Spar assembly 540 is similar to spar assembly 440, and thus, like parts of spar assembly 540 will be identified with like numbers increased by 100, it being understood that the description of like parts of spar assembly 440 applies to spar assembly 540 unless otherwise noted.

[0076] The metal spar 541 may extend from the base 550 of the metal trunnion 544 to define a hub 545 of the spar assembly 540. The metal spar 541 includes a first wall 551a and a second wall 551b connected by an intermediate truss 562 separating the first opening 554a from the second opening 554b. Although shown as being in the middle, the intermediate truss 562 may be located closer to either of the openings 554a, 554b and need not be located directly in the middle. The first opening 554a and the second opening 554b together define a socket 552. The first opening 554a is defined by the first wall 551a and the intermediate truss 562 and is open in a first direction 564a. The second opening 554b is defined by the second wall 551b and the intermediate truss 562 and is open in a second direction 564b opposite to the first direction 564a. In other words, the socket 552 has a lightning bolt or a lateral "Z" shape 553.

[0077] The composite spar 542 is received in the socket 552. The composite spar 542 includes a slot 563 formed to receive the intermediate truss 562 when assembled. The composite spar 542 may extend from the spar root 546 within the socket 552 in the spanwise direction S.

[0078] At least one reinforcement 543 is coupled to at least a portion of the metal spar 541. In one non-limiting example, at least one reinforcement 543 is coupled to the metal spar 541 to define a wing portion 561. The at least one reinforcement 543 may be a plurality of reinforcements 543a, 543b each defining a separate wing portion 561a, 561b.

[0079] The disclosure herein is applicable to bonding metal sheets to a composite spar / metal trunnion interface to help spread load transfer and provide chord-wise stiffness. Benefits of the disclosure include spreading load transfer while being versatile in location. These bonded metal sheets can be internal to the spar / trunnion interface or external to the interface area. Additionally, one benefit is the reduction of high composite spar stresses caused by chord-wise bending that occurs during extreme loading conditions.

[0080] It should be understood that application of the disclosed design is not limited to turbine engines having fan and supercharger sections, but is also applicable to turbojets and turbine engines.

[0081] To the extent not yet described, the different features and structures of each aspect can be used in combination or in place of each other as desired. A feature not shown in all examples does not mean that it cannot be shown as such, but is done so for the sake of brevity of description. Therefore, various features of different aspects can be mixed and matched as needed to form new aspects, whether or not the new aspects are explicitly described. All combinations or arrangements of the features described herein are covered by the present disclosure.

[0082] This written description uses examples to describe the aspects of the disclosure described herein, including the best mode, and also enables any person skilled in the art to practice aspects of the disclosure, including making and using any device or system and performing any combined method. The patentable scope of the various aspects of the disclosure is defined by the claims, and may include other examples that occur to one skilled in the art. These other examples are intended to fall within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that do not differ substantially from the literal language of the claims.

[0083] Further aspects of the disclosure are provided by the subject matter of the following clauses:

[0084] A component for a turbine engine, the component comprising: a wall bounding an interior and defining an exterior surface, the exterior surface extending radially between a leading edge and a trailing edge to define a chord-wise direction and extending between a root and a tip to define a spanwise direction; and a spar assembly comprising: a metal spar extending from a hub into the interior in the spanwise direction, the metal spar defining a socket; a composite spar extending between a spar root and a spar tip in the spanwise direction, at least a portion of the spar root being located in the socket; and a reinforcement member bonded to at least one of the wall, the metal spar, or the composite spar.

[0085] A component according to any preceding clause, wherein the wall defines an airfoil having a span length measured from the root to the tip, the root defining 0% of the span length and the tip defining 100% of the span length, and wherein the reinforcement is located at less than 20% of the span length.

[0086] A component as described in any preceding clause, further comprising a trunnion defining said hub and including said metal spar.

[0087] A component as described in any preceding clause, further comprising a wing portion extending at the hub in a direction perpendicular to the spanwise direction.

[0088] A component as claimed in any preceding clause, wherein the reinforcement defines the wing portion.

[0089] A component as claimed in any preceding clause wherein the composite spar defines the wing portion.

[0090] A component as claimed in any preceding clause wherein the reinforcement is bonded to the wing section.

[0091] A component as claimed in any preceding clause wherein the reinforcement is a plurality of parts.

[0092] A component as claimed in any preceding clause, wherein the reinforcement overlaps the metal spar.

[0093] A component as claimed in any preceding clause wherein the reinforcement is bonded to the outer surface of the wall.

[0094] A component as claimed in any preceding clause, wherein the reinforcement comprises at least one tapered edge.

[0095] A component as claimed in any preceding clause, wherein the socket has an H-shape.

[0096] A component as claimed in any preceding clause, wherein the socket has a Z-shape.

[0097] A component as claimed in any preceding clause, wherein the socket has a U-shape.

[0098] A component as claimed in any preceding clause wherein the component is a blade assembly.

[0099] A blade assembly for a turbine engine, the blade assembly comprising: a composite wall defining an interior and having an exterior surface, the composite wall defining a blade, the blade extending radially between a leading edge and a trailing edge to define a chord-wise direction, and extending between a root and a tip to define a span-wise direction; a spar assembly comprising: a trunnion defining a hub and having a set of walls defining a socket; a composite spar extending between a spar root and a spar tip in the span-wise direction, the spar root being located in the socket; a reinforcement member coupled to at least one of the composite wall, the trunnion, or the composite spar; and a wing portion at least partially defined by the reinforcement member.

[0100] A blade assembly according to any preceding clause, wherein the composite wall has a span length measured from the root to the tip, the root defining 0% of the span length and the tip defining 100% of the span length, and wherein the reinforcement is located at less than 20% of the span length.

[0101] A blade assembly according to any preceding clause, wherein the reinforcement is joined to the composite spar to define the wing section.

[0102] A blade assembly as claimed in any preceding clause, wherein the reinforcement overlaps the trunnion.

[0103] A blade assembly according to any preceding clause, wherein the reinforcement is bonded to the outer surface of the composite wall.

Claims

1. A component for a turbine engine, characterized in that: The components include: a wall bounding the interior and defining an exterior surface extending radially between the leading edge and the trailing edge to define a chord-wise direction and extending between the root and the tip to define a span-wise direction; and A wing spar assembly, the wing spar assembly comprising: a metal spar extending from the hub into the interior in the spanwise direction, the metal spar defining a socket having at least one side opening; a composite spar, the composite spar being different from the metal spar, the composite spar extending in the spanwise direction between a spar root and a spar tip, at least a portion of the spar root being located in the socket; and A reinforcement is bonded to at least a portion of the composite spar, at least a portion of the reinforcement extending at least partially through the at least one side opening.

2. The component according to claim 1, characterized in that in, The wall defines an airfoil having a span length measured from the root to the tip, the root defining 0% of the span length and the tip defining 100% of the span length, and wherein the reinforcement is located below 20% of the span length.

3. The component according to claim 1, characterized in that Further included is a trunnion defining the hub and including the metal spar.

4. The component according to claim 1, characterized in that Further included is a wing portion extending at the hub in a direction perpendicular to the spanwise direction.

5. The component according to claim 4, characterized in that in, The reinforcement defines the wing portion.

6. The component according to claim 4, characterized in that in, The composite spar defines the wing section.

7. The component according to claim 6, characterized in that in, The reinforcement is bonded to the wing section.

8. The component according to claim 1, characterized in that in, The reinforcement is a plurality of parts.

9. The component according to claim 1, characterized in that in, The reinforcement overlaps the metal spar.

10. The component according to claim 1, characterized in that in, The reinforcement member includes at least one tapered edge.

11. The component according to claim 1, characterized in that in, The socket has an H-shape.

12. The component according to claim 1, characterized in that in, The socket has a Z-shape.

13. The component according to claim 1, characterized in that in, The socket has a U-shape.

14. The component according to claim 1, characterized in that in, The component is a blade assembly.

15. A blade assembly for a turbine engine, characterized in that: The blade assembly comprises: a composite wall defining an interior and having an exterior surface, the composite wall defining a blade extending radially between a leading edge and a trailing edge to define a chord-wise direction and extending between a root and a tip to define a span-wise direction; A wing spar assembly, the wing spar assembly comprising: a trunnion defining a hub and having a set of walls defining a socket having at least one side opening, the blade being mounted to the spar assembly near the root via the trunnion, a composite spar extending in the spanwise direction between a spar root and a spar tip, the spar root being located in the socket, a reinforcement member bonded to at least one of the composite wall, the trunnion, or the composite spar, and A wing portion extends at least partially through the at least one side opening and is at least partially defined by the reinforcement.

16. The blade assembly according to claim 15, characterized in that in, The composite wall has a span length measured from the root to the tip, the root defining 0% of the span length and the tip defining 100% of the span length, and wherein the reinforcement is located at less than 20% of the span length.

17. The blade assembly according to claim 15, characterized in that in, The reinforcement is joined to the composite spar to define the wing section.

18. The blade assembly according to claim 15, characterized in that in, The reinforcement overlaps the trunnion.

Images