Airfoil assembly with trunnion and spar

Abstract

An airfoil assembly and a method of forming an airfoil assembly having a trunnion, a spar, a skin, and an insert. The trunnion has a flared socket with an open top. The spar extends from the flared socket and through the open top. The spar has a first end located within the flared socket. The skin has a woven or braided fiber.

Description

TECHNICAL FIELD

[0001] The present disclosure relates generally to airfoil assemblies, and more particularly, to airfoil assemblies having trunnions and spars. BACKGROUND

[0002] Turbine engines (and particularly gas or combustion turbine engines) are rotary engines that extract energy from a flow of gas passing through a fan with multiple fan blades, then through a series of compressor stages that include pairs of rotating blades and stationary vanes, through a combustor, and then through a series of turbine stages that include pairs of rotating blades and stationary vanes into the engine. The blades are mounted to a rotating disk, while the vanes are mounted to a stator disk.

[0003] During operation, air enters the compressor section through the fan section, is pressurized in the compressor, and is mixed with fuel in the combustor to generate hot combustion gases that flow downstream through the turbine stages where the air expands and exhausts the exhaust section. The expansion of the air in the turbine section is used to drive the rotating sections of the fan section and the compressor section. The intake of air, the pressurization of air, and the expansion of air are accomplished in part by the rotation of various rotating blades that are mounted to respective disks of the entire fan section, compressor section, and turbine section, respectively. The rotation of the rotating blades exerts mechanical stresses along various portions of the blades; particularly along the locations where the blades are mounted to the disks.

[0004] In some turbine engines, variable pitch airfoils can be included that can be selectively rotated to adjust or otherwise modulate fluid flow over the variable pitch airfoils. The variable pitch airfoils can be moved by using a trunnion and a spar. The trunnion can be rotated about a rotation axis that in turn rotates the spar and the variable pitch airfoils. The trunnion is coupled to or otherwise formed with the spar, and the airfoils are coupled to or otherwise formed with the spar. BRIEF DESCRIPTION OF DRAWINGS

[0005] A complete and enabling disclosure of the present disclosure, directed to one of ordinary skill in the art, is set forth in the specification, which is to be construed in connection with the accompanying drawings, wherein:

[0006] Figure 1 is a schematic cross-sectional view of a turbine engine.

[0007] Figure 2 is a schematic view of an airfoil assembly suitable for use within the turbine engine of Figure 1 , the airfoil assembly including an airfoil, a trunnion, and a spar.

[0008] Figure 3 is a schematic view of an airfoil assembly suitable for use within the turbine engine of Figure 2a schematic cross-sectional view of the airfoil assembly as seen from section line III-III, further showing the insert and a skin covering the insert and a portion of the spar.

[0009] Figure 4 is a schematic top-down view of the airfoil assembly as seen from section line IV-IV. Figure 3

[0010] Figure 5 is an exemplary method of forming Figure 2 the airfoil assembly. DETAILED DESCRIPTION

[0011] Aspects disclosed herein relate to airfoil assemblies of a turbine engine. The airfoil assembly includes an airfoil, a spar, and a trunnion. The spar couples the airfoil to the trunnion. The airfoil assembly further includes an insert facing a portion of the spar. A composite fiber weave is arranged on at least a portion of the insert and the spar.

[0012] For purposes of illustration, the present disclosure will be described with respect to airfoil assemblies of a turbine engine, particularly fan blades of a turbine engine. However, it will be understood that aspects of the present disclosure described herein are not so limited and can have general applicability within other engines or within other portions of a turbine engine. For example, the present disclosure can be applicable to airfoil assemblies in other engines or vehicles and can be used to provide benefits in industrial, commercial, and residential applications.

[0013] As used herein, the term "upstream" refers to a direction opposite to that of fluid flow, while the term "downstream" refers to a direction the same as that of fluid flow. The term "forward" or "front" denotes in front of something, while "aft" or "rear" denotes behind something. For example, when used in relation to fluid flow, forward / aft can denote upstream / downstream.

[0014] Further, as used herein, the term "radial" or "radially" refers to a direction away from a common center. For example, in the overall context of a turbine engine, radial refers to a direction along a ray extending between a central longitudinal axis of the engine and an outer periphery of the engine. Further, as used herein, the term "set" or a "set" of elements can be any number of elements, including only one.

[0015] ​Further, as used herein, the term“fluid” or iterations thereof can refer to any suitable fluid within a gas turbine engine, at least a portion of which is exposed to, for example, but not limited to, combustion gases, ambient air, pressurized air flow, working gas flow, or any combination thereof. It is further contemplated that the gas turbine engine can be other suitable turbine engines such as, but not limited to, a steam turbine engine or a supercritical carbon dioxide turbine engine. As a non-limiting example, the term“fluid” can refer to steam in a steam turbine engine or to carbon dioxide in a supercritical carbon dioxide turbine engine.

[0016] All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, aft, etc.) are only used for identification purposes to aid the reader’s understanding of the present disclosure, and do not create limitations, particularly as to the position, orientation, or use of the disclosure described herein. Connection references (e.g., attached, coupled, connected, joined, fixed, fastened, and the like) are to be construed in an open and inclusive way, and can encompass intermediate members between the elements connected to one another, and relative movement between elements unless otherwise indicated. Thus, connection references do not necessarily mean that two elements are directly connected and in fixed relation to one another. The exemplary drawings are for purposes of example only and the dimensions, positions, order and relative sizes reflected in the attached drawings attached hereto can vary.

[0017] As used herein, the term“composite” denotes a component having two or more materials. The composite can be a combination of at least two or more metals, non-metals, or a combination of metallic and non-metallic elements or materials. Examples of composites can be, but are not limited to, polymer matrix composites (PMC), ceramic matrix composites (CMC), metal matrix composites (MMC), carbon fibers, polymeric resins, thermoplastic resins, bismaleimide (BMI) materials, polyimide materials, epoxy resins, glass fibers, and silicon matrix materials.

[0018] 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, can include several layers or several plies of composite material. The stiffness, material, and size of the layers or plies can vary to achieve a desired composite component or composite portion of a component having a predetermined weight, size, stiffness, and strength.

[0019] One or more layers of adhesive can be used to form or couple composite components. The adhesive can include resins and phenolic resins, where the adhesive can need to be cured at elevated temperatures or other hardening techniques.

[0020] As used herein, PMC refers to a class of materials. By way of example, PMC materials are defined in part by prepreg, which is a reinforcing material pre-impregnated with a polymeric matrix material, such as a thermoplastic resin. Non-limiting examples of processes used to produce thermoplastic prepreg include: hot-melt prepreg, in which a fibrous reinforcement is pulled through a bath of molten resin; and powder prepreg, in which resin is deposited onto a fibrous reinforcement, such as by electrostatically depositing onto a fibrous reinforcement, which is then adhered to the fibers, such as in an oven or with the aid of heated rollers. The prepreg can be in the form of unidirectional tape or woven fabric, which are then stacked on top of one another to form the desired stack-up of composite plies for the part.

[0021] The multi-layer prepreg is stacked to the appropriate thickness and orientation for the composite part, and then the resin is cured and solidified to provide a fiber-reinforced composite part. Resins for PMC matrix materials can generally be classified as either thermoset or thermoplastic resins. Thermoplastic resins are generally classified as polymers that can repeatedly soften and flow upon heating and harden upon sufficient cooling due to a physical change rather than a chemical change. A notable example class of thermoplastic resins includes nylon, thermoplastic polyester, polyaryletherketone, and polycarbonate resins. Specific examples of high-performance thermoplastic resins that have been contemplated for aerospace applications include polyether ether ketone (PEEK), polyether ketone ketone (PEKK), polyetherimide (PEI), polyaryletherketone (PAEK), and polyphenylene sulfide (PPS). In contrast, thermoset resins do not undergo appreciable softening upon heating, but rather thermally decompose upon sufficient heating, once fully cured into a hard, rigid solid. Notable examples of thermoset resins include epoxy, bismaleimide (BMI), and polyimide resins.

[0022] Instead of using prepreg, in another non-limiting example, a woven fabric can be utilized by using a thermoplastic polymer. The woven fabric can include, but is not limited to, dry carbon fibers woven together with thermoplastic polymer fibers or filaments. A non-prepreg braided architecture can be fabricated in a similar manner. By this method, the fiber volume of the part can be tailored by specifying the relative concentrations of the thermoplastic and reinforcing fibers that have been woven or braided together. Further, different types of reinforcing fibers can be braided or woven together at different concentrations to tailor the properties of the part. For example, glass fibers, carbon fibers, and thermoplastic fibers can all be woven together at different concentrations to tailor the properties of the part. Carbon fibers provide the strength of the system, glass fibers can be incorporated to enhance impact properties, which is a design feature of parts located near the engine inlet, and thermoplastic fibers provide bonding for the reinforcing fibers.

[0023] In yet another non-limiting example, resin transfer molding (RTM) can be used to form at least a portion of the composite part. Generally, RTM includes applying dry fibers to a mold or cavity. The dry fibers can include a prepreg, a braided material, a woven material, or any combination thereof.

[0024] Resin can be pumped or otherwise provided to the mold or cavity to impregnate the dry fibers. The combination of impregnated fibers and resin is then cured and removed from the mold. The composite part can require post-cure treatment when removed from the mold.

[0025] It is contemplated that the RTM can be a vacuum assisted process. That is, air can be removed from the cavity or mold and replaced with resin prior to heating or curing. It is further contemplated that the placement of the dry fibers can be manual or automated.

[0026] The dry fibers can be shaped to shape the composite part or to direct the resin. Optionally, additional layers or reinforcement layers of material different from the dry fibers can also be included or added prior to heating or curing.

[0027] As used herein, CMC refers to a class of materials having reinforcing fibers in a ceramic matrix. Generally, the reinforcing fibers provide structural integrity to the ceramic matrix. Some examples of reinforcing fibers can include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), non-oxide carbon-based materials (e.g., carbon), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (AI2O3), silicon dioxide (SiO2), aluminosilicates (such as mullite), or mixtures thereof), or mixtures thereof.

[0028] Some examples of ceramic matrix materials can include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, aluminum oxide (AI2O3), silicon dioxide (SiO2), aluminosilicates, or mixtures thereof), or mixtures thereof. Optionally, ceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite) can also be included within the ceramic matrix.

[0029] Generally, specific CMCs can be referred to by their combination of fiber type / matrix type. 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 / nitride matrix hybrid, etc. In other examples, CMCs can be composed of a matrix comprising oxide-based materials such as alumina (AI2O3), silica (SiO2), aluminosilicates, and mixtures thereof, and reinforcing fibers. Aluminosilicates can include crystalline materials (e.g., mullite (3AI2O3-2SiO2)), as well as glassy aluminosilicates.

[0030] In certain non-limiting examples, the reinforcing fibers can be bundled and / or coated prior to being included in the matrix. For example, fiber tows can be formed into reinforcing tapes, such as unidirectional reinforcing tapes. Multiple tapes can be laid up together to form a preform component. The fiber tows can be impregnated with a slurry composition prior to forming the preform or after forming the preform. The preform can then be subjected to heat treatment and subsequent chemical treatment to yield a component formed from a CMC material having a desired chemical composition. For example, the preform can be subjected to cure or burnout to produce a high char residue in the preform and subsequently melt infiltration with silicon, or cure or pyrolysis to produce a silicon carbide matrix in the preform and subsequent chemical vapor infiltration with silicon carbide. Additional steps can be taken to improve the densification of the preform, either before or after chemical vapor infiltration, by infusing the preform with a liquid resin or polymer, followed by a heat treatment step to fill the voids with silicon carbide. The CMC materials as used herein can be formed using any known or later developed method, including but not limited to melt infiltration, chemical vapor infiltration, polymer impregnation pyrolysis (PIP), or any combination thereof.

[0031] Such materials, along with certain monolithic ceramics (i.e., ceramic materials without reinforcing materials), are particularly suitable for use in higher temperature applications. Further, these ceramic materials are lighter in weight compared to superalloys, yet are able to provide strength and durability to components made therefrom. As such, there is currently consideration being given to using such materials for many gas turbine components used in the higher temperature sections of gas turbine engines, such as airfoils (e.g., turbine and vane), combustors, shrouds, and the like, which would benefit from the lighter weight and higher temperature capability that these materials can provide.

[0032] The term "metal" as used herein means materials that include metals (e.g., but not limited to titanium, iron, aluminum, stainless steel, and nickel alloys). The metallic material or alloy can be a combination of at least two or more elements or materials, at least one of which is a metal.

[0033] Figure 1is a schematic cross-sectional view of a turbine engine 10 for an aircraft. The turbine engine 10 has a generally longitudinally extending axis or centerline 12 extending from a forward end 14 to an aft end 16. The turbine engine 10 includes, in downstream serial flow relationship: a fan section 18 including a fan 20; a compressor section 22 including a booster 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 a HP turbine 34 and a LP turbine 36; and an exhaust section 38.

[0034] The fan section 18 includes a fan casing 40 surrounding the fan 20. The fan 20 includes a plurality of fan blades 42 disposed radially about the engine centerline 12. The HP compressor 26, the combustor 30, and the HP turbine 34 form an engine core 44 of the turbine engine 10 that generates combustion gases. The engine core 44 is surrounded by a core casing 46, which can be coupled with the fan casing 40.

[0035] A HP shaft or spool 48 coaxially disposed about the engine centerline 12 of the turbine engine 10 drivingly connects the HP turbine 34 to the HP compressor 26. A LP shaft or spool 50 coaxially disposed about the engine centerline 12 of the turbine engine 10 within a 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 and are coupled to a plurality of rotatable elements, which can collectively define a rotor 51.

[0036] The LP compressor 24 and the HP compressor 26 each include a plurality of compressor stages 52, 54 with a set of compressor blades 56, 58 rotating relative to a corresponding set of stationary compressor vanes 60, 62 to compress or pressurize a fluid flow passing through the stage. In a single compressor stage 52, 54, the plurality of compressor blades 56, 58 can be disposed in a ring and can extend radially outward from a blade platform to a blade tip relative to the engine centerline 12, while the corresponding stationary compressor vanes 60, 62 are positioned upstream of and adjacent to the rotating compressor blades 56, 58. Notably, Figure 1 The number of blades, vanes, and compressor stages shown in FIG. 1 is selected for purposes of illustration only, and other numbers are possible.

[0037] The compressor blades 56, 58 for a stage of the compressor can be mounted to (or integrated with) a disk 61 that is mounted to a corresponding one of the HP and LP spools 48, 50. The stationary compressor vanes 60, 62 for a stage of the compressor can be mounted to the core casing 46 in a circumferential arrangement.

[0038] The HP turbine 34 and LP turbine 36 each include a plurality of turbine stages 64, 66 with a set of turbine blades 68, 70 rotating relative to a corresponding set of static turbine vanes 72, 74 (also referred to as nozzles) to extract energy from a fluid flow passing through the stage. In a single turbine stage 64, 66, the plurality of turbine blades 68, 70 can be arranged in a ring and can extend radially outward relative to the engine centerline 12, while the corresponding static turbine vanes 72, 74 are positioned upstream of and adjacent to the rotating turbine blades 68, 70. Notably, Figure 1 The number of blades, vanes, and turbine stages shown in FIG. 1 is chosen for purposes of illustration only, as other numbers are possible.

[0039] The turbine blades 68, 70 for a stage of the turbine can be mounted to a disk 71 mounted to a corresponding one of the HP and LP spools 48, 50. The turbine vanes 72, 74 for a stage of the compressor can be mounted to the core casing 46 in a circumferential arrangement.

[0040] Complementary to the rotor portion, the stationary portion of the turbine engine 10, such as the static vanes 60, 62, 72, 74 in the compressor section 22 and turbine section 32, are also referred to individually or collectively as a stator 63. Thus, the stator 63 can refer to the combination of non-rotating elements throughout the turbine engine 10.

[0041] In operation, the airflow exiting the fan section 18 is split so that a portion of the airflow is directed into the LP compressor 24, which then supplies a pressurized airflow 76 to the HP compressor 26, which further pressurizes the air. The pressurized airflow 76 from the HP compressor 26 is mixed with fuel in the combustor 30 and ignited, 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 exhaust is ultimately discharged from the turbine engine 10 via the exhaust section 38. The driving of the LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP compressor 24.

[0042] A portion of the pressurized airflow 76 can be extracted from the compressor section 22 as bleed air 77. The bleed air 77 can be extracted from the pressurized airflow 76 and provided to engine components that require cooling. The temperature of the pressurized airflow 76 entering the combustor 30 is significantly elevated above the bleed air temperature. The bleed air 77 can be used to reduce the temperature of core components downstream of the combustor 30. The bleed air 77 can also be utilized by other systems.

[0043] The remaining portion of the airflow, referred to as the bypass airflow 78, bypasses the LP compressor 24 and engine core 44 and exits the turbine engine 10 through stationary vanes at the fan exhaust side 84 and, more specifically, through an outlet guide vane assembly 80 (including a plurality of airfoil guide vanes 82) at the fan exhaust side 84. More specifically, a circumferential row of radially extending airfoil guide vanes 82 adjacent the fan section 18 are used to impart some directional control to the bypass airflow 78.

[0044] Some of the air supplied by the fan 20 can bypass the engine core 44 and be used to cool portions of the turbine engine 10, particularly hot portions, and / or to cool or power other aspects of the aircraft. In the case of a turbine engine, the hot portions of the engine are generally downstream of the combustor 30, particularly the turbine section 32, with the HP turbine 34 being the hottest portion as it is directly downstream of the combustion section 28. Other sources of cooling fluid can be, but are not limited to, fluid discharged from the LP compressor 24 or the HP compressor 26.

[0045] Figure 2 is a schematic view of an airfoil assembly 130 suitable for use within a turbine engine 10 of Figure 1 is a schematic view of an airfoil assembly 130 suitable for use within a turbine engine 10 of

[0046] The airfoil 132 can include a wall 138 that bounds an interior 148. The wall 138 can extend between a leading edge 144 and a trailing edge 146 to define a chordwise direction (C). The wall 138 can further extend between a root 140 and a tip 142 to define a spanwise direction (S). The wall 138 can be a composite wall made from one or more layers of composite material. The one or more layers of material can be applied during the same stage or different stages of the manufacture of the airfoil 132.

[0047] As a non-limiting example, the wall 138 can include at least a polymer matrix composite (PMC) portion or a polymer portion. The polymer matrix composite can include, but is not limited to, a thermoset matrix (epoxy, phenolic) or a thermoplastic matrix (polycarbonate, polyvinyl chloride, nylon, acrylic), as well as embedded glass, carbon, steel, or Kevlar fibers.

[0048] The airfoil assembly 130 can further include a spar 136 and a trunnion 134. The spar 136 can extend into the interior 148. The spar 136 can extend from the root 140. The spar 136 can be operably coupled to the trunnion 134. The spar 136 can be of any suitable material, such as, but not limited to, a composite material. The spar 136 can be a metal composite. The trunnion 134 can include any suitable material, such as, but not limited to, a metallic material or a composite material. It should be understood that the term composite material can further include a metal, but have a composite architecture (e.g., a metal matrix composite). In the case of a composite material, the spar 136 and / or the trunnion 134 can be of any suitable composite material, such as a 2D or 3D composite material, a laminated skin, a woven or braided composite material, or any other suitable composite material. The spar 136 can further include a composite material having a binder impregnated into the composite material. The spar 136 can then be cured to bind the composite material together. Thus, the spar 136 can be made of a pre-cured composite material.

[0049] The airfoil 132 has a span length (L) measured along the spanwise direction S from the root 140 at 0% of the span length (L) to the tip 142 at 100% of the span length (L). The entire spar 136 can be located below 20% of the span length (L). Alternatively, the spar 136 can extend beyond 20% of the span length (L).

[0050] During operation of the airfoil assembly 130, the trunnion 134 can rotate in the rotational direction (Rd) about the pitch axis (Pax). As the spar 136 couples the trunnion 134 to the airfoil 132, rotation of the trunnion 134 in the rotational direction (Rd) causes the airfoil 132 to rotate about the pitch axis (Pax). This rotation can be used to control the pitch of the airfoil assembly 130, such that the airfoil assembly 130 is defined as a variable pitch airfoil assembly. The pitch of the airfoil assembly 130 can vary based on the operation or intended operation of the turbine engine (e.g., the turbine engine 10 of Figure 1 .

[0051] Figure 3 is a schematic cross-sectional view of the airfoil assembly 130 as seen from the cross-sectional line III-III of Figure 2 . For purposes of illustration, the airfoil 132 Figure 2 has been removed from the airfoil assembly 130.

[0052] The trunnion 134 includes a wall 163 having an interior surface 162 that at least partially defines a flared socket 164 of the trunnion 134. The flared socket 164 extends between an open top 170 and a bottom 168. The spar 136 extends through the open top 170 and into the interior 148. The bottom 168 can be an open bottom or a sealed / closed bottom. The flared socket 164 can take any suitable shape having at least one flared cross-section, as shown.

[0053] The spar 136 extends along the centerline axis 150 and terminates within the flared socket 164 at a first end 160. The spar 136 can be symmetric or asymmetric about the centerline axis 150.

[0054] An insert 172 is received within the flared socket 164 abutting the first end 160 of the spar 136. The insert 172 extends between a first distal end 176 and a second distal end 178, with the second distal end 178 contacting the first end 160. The insert 172 includes opposing side walls 174 extending between the first distal end 176 and the second distal end 178.

[0055] A skin 180 is laid up or woven over at least a portion of the spar 136 and the insert 172. The skin 180 can be laid up or woven over at least a portion of the spar 136 as a dry fiber that is subsequently cured to define a composite layer covering the respective portion of the spar 136. As a non-limiting example, the skin 180 can be formed over the entire opposing side walls 174 or a portion of the opposing side walls 174 without being formed over the first distal end 176 or the second distal end 178. Alternatively, the skin 180 can be formed over at least a portion of the first distal end 176.

[0056] The spar 136 extends along the centerline axis 150 a first axial distance (Laxl). The skin 180 can extend along the spar 136 a second axial distance (Lax2). The first axial distance (Laxl) can be greater than the second axial distance (Lax2). As a non-limiting example, the second axial distance can be greater than 0% and less than or equal to 10% of the first axial distance (Laxl).

[0057] The insert 172 can include any suitable material that is the same or different than the spar 136 or the skin 180. As a non-limiting example, the insert 172 can include a metallic material or a composite material.

[0058] Extending the skin 180 by a second axial distance (Lax2) ensures that the skin 180 can at least partially engage the insert 172 to the spar 136. Furthermore, the extension of the skin 180 strengthens the engagement between the spar 136 and the insert 172 (e.g., the interface between the second distal end 178 and the first end 160). As a non-limiting example, if a bending moment is applied to the spar 136, the skin 180 serves to limit the effect of the bending moment. In other words, without the skin 180, the bending moment could cause the first end 160 and the second distal end 178 to separate.

[0059] Insert 172 is used to help retain the spar 136 within the flared socket 164. This is accomplished by attaching insert 172 to spar 136 and subsequently using insert 172 to radially push the skin 180 outward, such that the skin 180, and thus insert 172 and spar 136, maintain frictional contact with the trunnion 134. The attachment of insert 172 to spar 136 can be accomplished by any suitable method (e.g., but not limited to co-curing at least insert 172 and spar 136 together). It will be further understood that spar 136 and insert 172 can be retained within the trunnion 134 by any other suitable method (e.g., but not limited to co-curing, adhesion, welding, bonding, fastening, etc.).

[0060] Figure 4 It is along the opening at the top 170 of the trunnion 134. Figure 3 The figure shows a schematic top-down view of the airfoil assembly as seen along section line IV-IV. As shown, the spar 136 and skin 180 can be polygonal (e.g., non-circular). As a non-limiting example, when cut along a horizontal plane perpendicular to the centerline axis 150, the spar 136 and skin 180 can each include a rectangular cross-section. The opening top 170 can have a cross-section corresponding to that of the spar 136. As a non-limiting example, when viewed along a horizontal plane, the opening top 170 can have a rectangular cross-section. The rectangular cross-section can be provided along the entire flared socket 164. Alternatively, the cross-section of the opening top 170 can differ from the cross-section of the spar 136. Alternatively, the opening top 170 can have any suitable non-rectangular cross-section, such as a circular cross-section.

[0061] Figure 5 It is formed Figure 2 An exemplary method 182 of the airfoil component assembly 130. This will be achieved by combining method 182 with... Figures 2-4 Refer to the physical aspects of the airfoil assembly outlined in the document. Figures 2-5 .

[0062] Method 182 can begin by positioning the insert 172 abutting against the spar 136 at 184. Specifically, the second distal end 178 can be positioned abutting against the first end 160 of the spar 136. At 186, a pre-woven skin 180 is woven or wrapped around the corresponding portions of the spar 136 and the insert 172. When woven around the corresponding portions of the spar 136 and the insert 172, the skin 180 can be woven directly onto the corresponding portions of the spar 136 and the insert 172. When wrapping the skin 180 around the corresponding portions of the spar 136 and the insert 172, the skin 180 can be woven first and then placed or otherwise wrapped around the corresponding portions of the spar 136 and the insert 172. The insert 172 can then be removed from the interior of the skin 180 and the spar 136 at 188. The skin 180 and spar 136 can then be inserted at 190 through the top opening 170 into the flared socket 164 of the trunnion 134. This may require bending or compressing the skin 180 radially inward toward the centerline axis 150 to ensure that the skin 180 can fit through the top opening 170. Once inserted, the skin 180 can be flared outward and contact the inner surface 162. At 192, the insert 172 can be positioned within the trunnion 134 such that the second distal end 178 faces the first end 160. This can be done by inserting the insert 172 through the bottom 168 of the trunnion 134. Once the insert is inserted, the airfoil assembly 130 can appear as... Figure 2 The airfoil assembly 130 shown. The spar 136, skin 180 and insert 172 can then be fixed together at 194, so that the spar 136, skin 180 and insert 172 form a single body.

[0063] It should be understood that method 182 is flexible and is shown for illustrative purposes only. For example, the order of the depicted steps is for illustrative purposes only and is not intended to limit method 182 in any way, as it should be understood that these steps may be performed in a different logical order without diminishing the embodiments of the invention, or may include additional or intervening steps. As a non-limiting example, the curing step at 194 may further include curing the trunnion 134, spar 136, skin 180, and insert 172 together such that the trunnion 134, spar 136, skin 180, and insert 172 form a single body.

[0064] Method 182 can further include pre-curing the spar 136. In other words, the spar 136 can be cured prior to inserting the spar 136 into the trunnion 134 at 190. The spar 136 can be cured prior to routing or wrapping the skin 180 at 186 or prior to positioning the insert 172 at 184. Pre-curing the spar 136 can be used to ensure that the composite materials of the spar 136 are fully bonded, thereby being as strong as possible. It is further contemplated that method 182 can further include using a first insert formed as the insert 172 at 184 and placing a second insert different from the first insert at 190. It is contemplated that both the first insert and the second insert can be formed as the insert 172, but the first insert is used only during routing or wrapping at 186, while the second insert is formed as the final structural piece of the airfoil assembly 130. The first insert can thus be defined as a mandrel used during manufacturing and can be used to form multiple airfoil assemblies 130, while the second insert can be defined as a structural insert used in a single airfoil assembly 130.

[0065] Method 182 can further include placing the spar 136 into the interior 148 of the airfoil 132. The spar 136 can be placed into the interior 148 of the airfoil 132 prior to or after the steps described in method 182. Method 182 can further include providing the airfoil assembly 130 within a portion of the turbine engine 10 Figure 1 ).

[0066] When compared to conventional airfoil assemblies, the benefits associated with the present disclosure include airfoil assemblies having reduced manufacturing burdens. For example, conventional airfoil assemblies including a spar and a trunnion require additional structure to ensure that the airfoil of the conventional airfoil assembly is coupled to the trunnion. For example, conventional airfoil assemblies utilize a laminated spar that is disposed within a flared socket of the trunnion. However, positioning the laminated spar within the flared socket such that the laminated spar is retained within the flared socket requires a time-intensive and labor-intensive method of isolating certain portions of the laminated spar to couple the laminated spar to the trunnion. Conventional airfoil assemblies can further require that the spar be physically coupled to the trunnion or formed with the trunnion, which in turn can make assembly and disassembly (e.g., during maintenance) difficult. However, the airfoil assembly as described herein includes a spar, a skin, and an insert that are cured together after assembly. This configuration of the airfoil assembly greatly reduces the complexity of the airfoil assembly relative to conventional airfoil assemblies, which in turn greatly reduces manufacturing burdens.

[0067] For example, different features and structures of various embodiments can be combined to create new embodiments, within the scope of the present disclosure, that were not specifically described in the foregoing. The scope of the disclosure is not to be interpreted merely from the description of various embodiments. Rather, the scope of the disclosure is to be interpreted from the claims.

[0068] This written description uses examples to describe aspects of the disclosure described herein, including the best mode, and also to enable any person skilled in the art to practice aspects of the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of aspects of the disclosure is defined by the claims, and can include other examples that occur to those skilled in the art. Such 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

[0069] Further aspects are provided by the subject matter of the following clauses:

[0070] A airfoil assembly for a turbine engine, the airfoil assembly comprising: an airfoil comprising a wall defining an interior and an exterior surface, the wall extending between a leading edge and a trailing edge to define a chordwise direction and between a root and a tip to define a spanwise direction; a trunnion having a flared socket with an open top; a spar extending from the flared socket, through the open top of the flared socket, and into the interior of the airfoil, and having a first end located in the flared socket; an insert received within the flared socket; and a skin comprising a woven or braided composite, the skin extending along at least a portion of the spar and the insert.

[0071] A method of forming an airfoil assembly having a trunnion, a spar, and an insert, the trunnion having a flared socket with an open top and an open bottom, the spar having a first end, the method comprising: positioning the insert against the first end; weaving or wrapping a skin over a portion of the spar and the insert; removing the insert from the first end and the skin; placing the skin and the spar into the trunnion through the open top of the trunnion; placing the insert through the open bottom such that the insert faces the first end and is at least partially covered by the skin; and curing the spar and the insert such that the spar and the insert form a unitary body.

[0072] A turbine engine including an airfoil assembly having an airfoil including a wall defining an interior and an exterior surface, the wall extending between a leading edge and a trailing edge to define a chordwise direction and between a root and a tip to define a spanwise direction, a trunnion having a flared socket with an open top, a spar extending from the flared socket, through the open top of the flared socket, and into the interior of the airfoil, and having a first end located in the flared socket, an insert received within the flared socket, and a skin including a woven or braided composite material, the skin extending along at least a portion of the spar and the insert.

[0073] The airfoil assembly of any preceding paragraph, wherein the spar includes a composite material.

[0074] The airfoil assembly of any preceding paragraph, wherein the spar includes a pre-cured composite material.

[0075] The airfoil assembly of any preceding paragraph, wherein the insert includes opposing side walls and a distal end interconnecting the opposing side walls, wherein the skin is woven or braided over the opposing side walls and not over the distal end.

[0076] The airfoil assembly of any preceding paragraph, wherein the spar extends a span length in the spanwise direction, and the skin extends along the spar greater than 0% and less than or equal to 10% of the span length.

[0077] The airfoil assembly of any preceding paragraph, wherein the spar, the insert, and the skin form a single body co-cured with one another.

[0078] The airfoil assembly of any preceding paragraph, wherein the trunnion is co-cured with the spar, the insert, and the skin.

[0079] The airfoil assembly of any preceding paragraph, wherein the spar extends along a centerline axis in the spanwise direction, and the flared socket defines a rectangular cross-section when viewed from a horizontal plane perpendicular to the centerline axis.

[0080] The airfoil assembly of any preceding paragraph, wherein the spar defines a rectangular cross-section when viewed along the horizontal plane.

[0081] The airfoil assembly of any preceding paragraph, wherein the insert includes a metallic material or a composite material.

[0082] The airfoil assembly of any preceding paragraph, wherein the turbine engine includes a fan section, a compressor section, a combustion section, and a turbine section in a serial flow arrangement, wherein the airfoil assembly is disposed within one of the fan section, the compressor section, or the turbine section.

[0083] The method of any preceding paragraph, further comprising curing the trunnion, the spar, and the insert such that the spar, the trunnion, and the insert form a unitary body.

[0084] The method of any preceding paragraph, wherein the spar extends along a centerline axis, and the method further comprises draping or wrapping the skin along the centerline axis over greater than 0% and less than or equal to 10% of a total length of the spar.

[0085] The method of any preceding paragraph, further comprising pre-curing the spar.

[0086] The method of any preceding paragraph, further comprising forming the spar from a composite material.

[0087] The method of any preceding paragraph, further comprising forming the insert from a metallic material.

[0088] The method of any preceding paragraph, wherein the insert includes opposing sidewalls and a distal end interconnecting the opposing sidewalls, the method further comprising draping or wrapping the skin over the entire opposing sidewalls of the insert and not over the distal end.

[0089] The method of any preceding paragraph, further comprising disposing the spar within an interior of an airfoil, the airfoil having a wall bounding the interior and defining an exterior surface, the wall extending between a leading edge and a trailing edge to define a chordwise direction and between a root and a tip to define a spanwise direction.

[0090] The method of any preceding paragraph, further comprising disposing the airfoil assembly within one of a fan section, a compressor section, or a turbine section of a turbine engine.

[0091] The airfoil assembly of any preceding paragraph, wherein the airfoil includes a composite material.

Claims

1. An airfoil assembly for a turbine engine, characterized by, The airfoil assembly comprises: an airfoil comprising a wall defining an interior and an exterior surface, the wall extending between a leading edge and a trailing edge to define a chordwise direction and between a root and a tip to define a spanwise direction; a trunnion having a flared socket with an open top; a spar extending from the flared socket, through the open top of the flared socket, and into the interior of the airfoil, and having a first end located in the flared socket; an insert received within the flared socket; and a skin comprising a woven or braided composite material, the skin extending along at least a portion of the spar and the insert, wherein the skin at least partially covers the insert and the spar.

2. The airfoil assembly of claim 1, wherein, wherein, the spar comprises a composite material.

3. The airfoil assembly of claim 2, wherein, wherein, the spar comprises a pre-cured composite material.

4. The airfoil assembly of claim 1, wherein, wherein, the insert comprises opposing side walls and a distal end interconnecting the opposing side walls, wherein the skin is woven or braided over the opposing side walls and not over the distal end.

5. The airfoil assembly of claim 1, wherein, wherein, the spar extends a span length in the spanwise direction, and the skin extends along the spar greater than 0% and less than or equal to 10% of the span length.

6. The airfoil assembly of claim 1, wherein, wherein, the spar, the insert, and the skin form a single body co-cured with one another.

7. The airfoil assembly of claim 6, wherein, wherein, the trunnion is co-cured with the spar, the insert, and the skin.

8. The airfoil assembly of claim 1, wherein, wherein, the spar extends along a centerline axis in the spanwise direction, and the flared socket defines a rectangular cross-section when viewed from a horizontal plane perpendicular to the centerline axis.

9. The airfoil assembly of claim 8, wherein, wherein, the spar defines a rectangular cross-section when viewed along the horizontal plane.

10. The airfoil assembly of claim 1, wherein, wherein, the insert comprises a metallic material or a composite material.

11. The airfoil assembly of Claim 1, wherein, wherein, the turbine engine comprises a fan section, a compressor section, a combustion section, and a turbine section in a serial flow arrangement, wherein the airfoil assembly is disposed within one of the fan section, the compressor section, or the turbine section.

12. A method of forming a wing assembly having a trunnion, a spar and an insert, the trunnion having a flared socket with an open top and an open bottom, the spar having a first end, the method comprising: The method comprises: positioning the insert against the first end; weaving or wrapping a skin over a portion of the spar and the insert; removing the insert from the first end and the skin; placing the skin and the spar into the trunnion through the open top of the trunnion; placing the insert through the open bottom such that the insert faces the first end and is at least partially covered by the skin; and curing the spar and the insert such that the spar and the insert form a single body.

13. The method of claim 12, wherein, further comprising curing the trunnion, the spar, and the insert such that the spar, the trunnion, and the insert form a single body.

14. The method of claim 12, wherein, wherein, the spar extends along a centerline axis, and the method further comprises weaving or wrapping the skin along the centerline axis over greater than 0% and less than or equal to 10% of a total length of the spar.

15. The method of claim 12, wherein, further comprising pre-curing the spar.

16. The method of claim 12, wherein, further comprising forming the spar with a composite material.

17. The method of claim 12, wherein, Further comprising forming the insert with a metallic material.

18. The method of claim 12, wherein, wherein, the insert includes opposing side walls and a distal end interconnecting the opposing side walls, the method further comprising routing or wrapping the skin over the entire opposing side walls of the insert and not over the distal end.

19. The method of claim 12, wherein, Further comprising disposing the spar within an interior of an airfoil, the airfoil having walls bounding the interior and defining an exterior surface, the walls extending between a leading edge and a trailing edge to define a chordwise direction and between a root and a tip to define a spanwise direction.

20. The method of claim 12, wherein, Further comprising disposing the airfoil assembly within one of a fan section, a compressor section, or a turbine section of a turbine engine. Further comprising forming the insert with a metallic material. wherein, the insert includes opposing side walls and a distal end interconnecting the opposing side walls, the method further comprising routing or wrapping the skin over the entire opposing side walls of the insert and not over the distal end. Further comprising disposing the spar within an interior of an airfoil, the airfoil having walls bounding the interior and defining an exterior surface, the walls extending between a leading edge and a trailing edge to define a chordwise direction and between a root and a tip to define a spanwise direction. Further comprising disposing the airfoil assembly within one of a fan section, a compressor section, or a turbine section of a turbine engine.

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