Parts having composite laminates with co-cured chopped fibers

Abstract

Composite components, such as a gas turbine engine airfoil, and methods of forming composite components are provided. For example, a composite component of a gas turbine engine includes a plurality of continuous fiber layers, each continuous fiber layer formed of a plurality of continuous fibers disposed in a matrix material; and a chopped fiber layer including a plurality of chopped fibers formed into a thin film sheet. The chopped fiber layer is laminated with the plurality of continuous fiber layers to form the composite component. A method for forming a composite component includes laying up a plurality of continuous fiber layers, each continuous fiber layer including a plurality of continuous fibers disposed in a matrix material; laminating a chopped fiber layer with the plurality of continuous fiber layers, the continuous fiber layers and the chopped fiber layer forming a reinforcement laminate; and curing the reinforcement laminate.

Description

Technical Field

[0001] This subject matter generally relates to composite components. More specifically, this subject matter relates to composite components (e.g., airfoils or other components) formed from composite laminates comprising chopped fibers co-cured with a composite material. Background Technology

[0002] Aircraft gas turbine engines operate under various conditions, and under certain load conditions, fan blade tips may rub against the fan casing. The fan casing can be designed, for example, to accommodate blade friction through abrasive material sections, groove fillers, etc., to mitigate the effects of blade friction on composite fan blades. Furthermore, composite fan blades may include metal tip caps, metal leading and / or trailing edge guards, and / or other external metal structural reinforcements to mitigate damage from delamination and fiber breakage, as well as debris impact and / or blade friction. However, reinforcing the fan casing and / or fan blades increases the weight and complexity of the fan module, which increases the cost of the fan module and affects engine performance, without unavoidable rotor imbalance, such as during FBO events. Therefore, improved composite airfoils for gas turbine engines with features that mitigate airfoil tip friction, delamination, fiber breakage, and / or other airfoil damage are desirable. Summary of the Invention

[0003] The aspects and advantages of this disclosure will be set forth in part in the description which follows, or may be apparent from the description, or may be learned by practice of this disclosure.

[0004] In one embodiment of this subject matter, a composite airfoil for a gas turbine engine is provided. The composite airfoil includes opposing pressure and suction sides extending radially from a root to a tip along the wingspan. The root defines a first radial end of the composite airfoil, and the tip defines a second radial end of the composite airfoil. The composite airfoil further includes opposing leading and trailing edges extending radially along the wingspan. The pressure and suction sides extend axially between the leading and trailing edges. The composite airfoil also includes a plurality of continuous fiber layers, each containing a plurality of continuous fibers; and a chopped fiber layer containing a plurality of chopped fibers formed as a thin film sheet. The chopped fiber layer is stacked with the plurality of continuous fiber layers to form the composite airfoil.

[0005] In another embodiment of this subject matter, a method for forming a composite component of a gas turbine engine is provided. The method includes laying a plurality of continuous fiber layers, each continuous fiber layer comprising a plurality of continuous fibers disposed in a first matrix material; stacking chopped fiber layers with the plurality of continuous fiber layers, the stack of the plurality of continuous fiber layers and the chopped fiber layers forming a reinforcing stack; and curing the reinforcing stack.

[0006] In another embodiment of this subject matter, a composite component for a gas turbine engine is provided. The composite component includes a plurality of continuous fiber layers, each formed of a plurality of continuous fibers disposed in a matrix material; and a chopped fiber layer comprising a plurality of chopped fibers formed as a thin film sheet. The chopped fiber layer is stacked with the plurality of continuous fiber layers to form the composite component.

[0007] These and other features, aspects, and advantages of this disclosure will be better understood by referring to the following description and the appended claims. The accompanying drawings, which are incorporated in and form part of this specification, illustrate embodiments of the disclosure and, together with the description, serve to explain the principles of the disclosure. Attached Figure Description

[0008] The specification with reference to the accompanying drawings sets forth a complete and feasible disclosure for those skilled in the art, including its best mode, wherein:

[0009] Figure 1 Schematic cross-sectional views of gas turbine engines according to various embodiments of this subject are provided.

[0010] Figure 2 Provided Figure 1 A side perspective view of a fan blade of a gas turbine engine having a composite airfoil according to one embodiment of the subject matter.

[0011] Figure 3 A schematic diagram of a stack of multiple continuous fiber layers and multiple chopped fiber layers according to one embodiment of the subject matter is provided.

[0012] Figure 4 A schematic diagram of a stack of multiple continuous fiber layers and multiple chopped fiber layers according to another embodiment of this subject matter is provided.

[0013] Figure 5A A schematic diagram of a thin film sheet comprising a plurality of chopped fibers in a matrix material is provided according to one embodiment of the subject matter.

[0014] Figure 5B A schematic diagram of two short-cut fibers with different lengths and diameters, according to one embodiment of this subject matter, is provided.

[0015] Figures 6 to 8 All provided Figure 1 A side perspective view of a fan blade of a gas turbine engine having a composite airfoil according to various embodiments of the subject matter.

[0016] Figure 9 A flowchart illustrating a method for forming a composite airfoil according to an embodiment of this subject is provided. Detailed Implementation

[0017] Reference will now be made in detail to embodiments of this disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerals and letter reference numerals to denote features in the drawings. Similar or analogous reference numerals in the drawings and description have been used to denote similar or analogous portions of the disclosed embodiments.

[0018] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components.

[0019] The terms "front" and "rear" refer to relative positions within a gas turbine engine or vehicle and to the normal operating posture of the gas turbine engine or vehicle. For example, for a gas turbine engine, "front" refers to the position closer to the engine inlet, while "rear" refers to the position closer to the engine nozzle or exhaust port.

[0020] The terms "upstream" and "downstream" refer to the relative directions of fluid flow within a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction in which the fluid flows.

[0021] Unless otherwise specified herein, the terms “connection,” “fixed,” “attached to,” etc., refer to both direct connection, fixation, or attachment, and indirect connection, fixation, or attachment through one or more intermediate components or features.

[0022] Unless the context clearly indicates otherwise, the singular forms “a,” “one,” and “the” include plural references.

[0023] Approximate language, as used throughout the specification and claims, is applied to modify any quantitative expression that allows for variation without altering its underlying function. Therefore, values ​​modified by one or more terms such as “about,” “approximately,” and “substantially” are not limited to specified exact values. In at least some cases, approximate language may correspond to the precision of the instrument used to measure the value, or the precision of the method or machine used to construct or manufacture the component and / or system. Approximate language may refer to a margin of + / - 1, 2, 4, 10, 15, or 20% at the endpoints of a single value, a range of values, and / or a defined range of values.

[0024] Throughout this specification and claims, scope limitations are combined and interchanged, and such scopes are identified and include all subscopes contained therein, unless the context or language otherwise indicates. For example, all scopes disclosed herein include endpoints, and endpoints may be combined independently of each other.

[0025] Typically, this subject matter provides composite components comprising multiple continuous fiber layers and at least one chopped fiber layer that reinforces or strengthens the composite component against delamination, fiber breakage, or damage from friction or debris impact. As described herein, each continuous fiber layer comprises multiple continuous fibers disposed in a matrix material, and each chopped fiber layer comprises multiple chopped fibers formed as a thin film sheet. For example, the thin film sheet can be formed by injecting a matrix material into the multiple chopped fibers, which may be the same as or different from the matrix material in which the continuous fibers of the continuous fiber layers are disposed. One or more chopped fiber layers are stacked with multiple continuous fiber layers to form the composite component. In some embodiments, the chopped fibers may be metallic, but other non-metallic materials may also be used for the chopped fibers. Furthermore, the composite component may be a gas turbine engine component, such as an airfoil, more specifically, a fan blade, but the composite component can be used for any other suitable application.

[0026] Referring now to the accompanying drawings, where the same numerals throughout the drawings indicate the same elements. Figure 1 This is a schematic cross-sectional view of a gas turbine engine according to an embodiment of the present disclosure. More specifically, for Figure 1 In one embodiment, the gas turbine engine is a high-bypass turbofan jet engine 10, referred to herein as "turbofan engine 10". Figure 1 As shown, the turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 for reference) and a radial direction R. Typically, the turbofan engine 10 includes a fan section 14 and a core turbine engine 16 disposed downstream of the fan section 14.

[0027] The core turbine engine 16 shown typically includes a basic tubular housing 18 defining an annular inlet 20. The housing 18 surrounds a compressor section in series flow relationship, which includes a turbocharger or low-pressure (LP) compressor 22 and a high-pressure (HP) compressor 24; a combustion section 26; a turbine section including a high-pressure (HP) turbine 28 and a low-pressure (LP) turbine 30; and an exhaust nozzle section 32. A high-pressure (HP) shaft or spool 34 drives the HP turbine 28 to the HP compressor 24. A low-pressure (LP) shaft or spool 36 drives the LP turbine 30 to the LP compressor 22.

[0028] In the illustrated embodiment, fan section 14 includes a fan 38 having a plurality of fan blades 40 spaced apart and coupled to disk 42. As shown, the fan blades 40 extend generally radially outward from disk 42. The fan blades 40 and disk 42 together can be rotated about longitudinal axis 12 via LP shaft 36. In some embodiments, a power gearbox with a plurality of gears may be included to reduce the rotational speed of LP shaft 36 to a more efficient fan rotation speed.

[0029] Still referencing Figure 1 In one embodiment, the disk 42 is covered by a rotatable forward nacelle 48, which has an aerodynamic profile to facilitate airflow through multiple fan blades 40. Furthermore, the fan section 14 includes an annular fan housing or outer nacelle 50 circumferentially surrounding at least a portion of the fan 38 and / or the core turbine engine 16. It should be understood that the nacelle 50 may be configured to be supported relative to the core turbine engine 16 by a plurality of circumferentially spaced outlet guide vanes 52. Additionally, a downstream section 54 of the nacelle 50 may extend over the outer portion of the core turbine engine 16 to define a bypass airflow passage 56 therebetween.

[0030] During operation of the turbofan engine 10, a volume of air 58 enters the turbofan engine 10 through the nacelle 50 and / or the associated inlet 60 of the fan section 14. As the volume of air 58 passes through the fan blades 40, a first portion of the air 58, as indicated by arrow 62, is directed or guided into the bypass airflow passage 56, and a second portion of the air 58, as indicated by arrow 64, is directed or guided into the LP compressor 22. The ratio between the first portion of air 62 and the second portion of air 64 is commonly referred to as the bypass ratio. The pressure of the second portion of air 64 then increases as it is delivered through the high-pressure (HP) compressor 24 and into the combustion section 26, where it mixes with fuel and is burned to provide combustion gases 66.

[0031] Combustion gas 66 is guided through HP turbine 28, where a portion of the thermal and / or kinetic energy from combustion gas 66 is extracted via a sequential stage of HP turbine stator blades 68 connected to housing 18 and HP turbine rotor blades 70 connected to HP shaft or spool 34, thus causing HP shaft or spool 34 to rotate, thereby supporting the operation of HP compressor 24. Combustion gas 66 is then guided through LP turbine 30, where a second portion of the thermal and kinetic energy from combustion gas 66 is extracted via a sequential stage of LP turbine stator blades 72 connected to housing 18 and LP turbine rotor blades 74 connected to LP shaft or spool 36, thus causing LP shaft or spool 36 to rotate, thereby supporting the operation of LP compressor 22 and / or the rotation of fan 38.

[0032] Combustion gas 66 is then directed through the injection exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. Simultaneously, the pressure of the first portion of air 62 increases significantly as it is directed through the bypass airflow passage 56 before exiting the fan nozzle exhaust section 76 of the turbofan 10, also providing propulsive thrust. The HP turbine 28, LP turbine 30, and injection exhaust nozzle section 32 at least partially define the hot gas path 78 for directing combustion gas 66 through the core turbine engine 16.

[0033] In some embodiments, components of the turbofan engine 10 may include composite materials with high-temperature capabilities, such as polymer-based composites (PMCs) or ceramic-based composites (CMCs). Composite materials typically contain fiber reinforcements, such as polymer or ceramic-based materials, embedded in a matrix material. The reinforcements serve as the load-bearing component of the composite material, while the matrix of the composite material serves to bond the fibers together and act as a medium to transfer and distribute externally applied stress to the fibers. As used herein, the term "composite" should be understood to include, but is not limited to, PMCs, CMCs, and hybrid composites, such as combinations of PMCs or CMCs with one or more metallic materials, or combinations of PMCs or CMCs with more than one type of PMC or CMC.

[0034] PMC materials are typically manufactured by impregnating a fabric or unidirectional tape with a resin (prepreg) and then curing it. Prior to impregnation, the fabric may be referred to as a "dry" fabric and typically comprises a stack of two or more fiber layers (layers). The fiber layers can be formed from a variety of materials, non-limiting examples of which include carbon (e.g., graphite), glass (e.g., glass fiber), and polymers (e.g., polymers). Fibers and metal fibers. Fiber-reinforced materials can be used in the form of relatively short chopped fibers, typically less than two inches in length, more preferably less than one inch, or long continuous fibers, the latter typically used to produce woven fabrics or unidirectional tapes. PMC materials can be produced by dispersing dry fibers into a mold and then allowing the matrix material to flow around the reinforcing fibers, or by using prepregs. For example, multiple layers of prepreg can be stacked to the appropriate thickness and orientation for the part, and then the resin can be cured and solidified to provide fiber-reinforced composite parts. Resins used for PMC matrix materials are generally classified as thermosetting or thermoplastic. Thermoplastic resins are generally classified as polymers that can repeatedly soften and flow upon heating due to physical rather than chemical changes and can harden upon adequate cooling. Notable examples of thermoplastic resin categories include nylon, thermoplastic polyesters, polyaryletherketones (PEEKs), and polycarbonate resins. Specific examples of high-performance thermoplastic resins considered for aerospace applications include polyetheretherketone (PEEK), polyetherketoneketone (PEKK), polyetherimide (PEI), and polyphenylene sulfide (PPS). Conversely, once fully cured into a rigid solid, thermosetting resins do not undergo significant softening upon heating, but rather thermal decomposition upon sufficient heating. Notable examples of thermosetting resins include epoxy resins, bismaleimide (BMI), and polyimide resins. Therefore, PMC materials typically comprise a thermosetting or thermoplastic matrix and reinforcing materials, including but not limited to glass, graphite, aramids, or organic fibers of any length, size, or orientation, or combinations of these reinforcing materials, and are further understood to include, but not limited to, manufacturing by injection molding, resin transfer molding, prepreg tape lamination (manual or automated), pultrusion, or any other suitable method for manufacturing reinforced polymer matrix composite structures, or combinations of these manufacturing methods.

[0035] Exemplary CMC materials may include silicon carbide (SiC), silicon, silica, carbon or alumina matrix materials, and combinations thereof. Ceramic fibers may be embedded within the matrix, such as oxide-stabilized reinforcing fibers, including monofilaments like sapphire and silicon carbide (e.g., Textron's SCS-6), as well as rovings and yarns, including silicon carbide (e.g., Nippon Carbon's). Ube Industries And Dao Corning Aluminosilicate (e.g., 3M's Nextel 440 and 480) and chopped whiskers and fibers (e.g., 3M's Nextel 440 and 480) The material may include, for example, ceramic particles (e.g., oxides of silicon, aluminum, zirconium, yttrium, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite). For instance, in some embodiments, fiber bundles comprising a ceramic refractory coating may be formed as reinforcing strips, such as unidirectional reinforcing strips. Multiple strips may be stacked together (e.g., as layers) to form a preform component. The fiber bundles may be impregnated with a slurry composition before or after forming the preform (e.g., a prepreg layer). The preform may then be heat-treated, such as cured or burned off, to produce a high coke residue in the preform, followed by chemical treatment, such as silica infiltration, to obtain a component formed from a CMC material having the desired chemical composition. In other embodiments, the CMC material may be formed, for example, carbon fiber cloth instead of strips.

[0036] Turn Figure 2 The composite components of a gas turbine engine (e.g., engine 10) will be described according to embodiments of this subject matter. Figure 2 As shown in the embodiments, the composite component 100 may be a composite airfoil, such as a fan blade 40, which may be referred to herein as "airfoil 100". In other embodiments, the composite component 100 may be another composite airfoil, such as an inlet guide vane (IGV) or an outlet guide vane (OGV) 52, or other composite components, such as a fan housing, shroud, etc.

[0037] Figure 2 The airfoil 100 shown includes a concave pressure side 102 opposite to the convex suction side 104. The opposing pressure side 102 and suction side 104 of the airfoil 100 extend radially along the wingspan S from the root 106 to a tip 108 at the radially outermost portion of the fan blade 40. That is, the root 106 defines a first radial end of the airfoil 100, while the tip 108 defines a second radial end of the airfoil 100. The pressure side 102 and suction side 104 of the airfoil 100 extend axially along a chord length c between the leading edge 110 and the opposing trailing edge 112. The leading edge 110 defines the front end of the airfoil 100, while the trailing edge 112 defines the rear end of the airfoil 100. Furthermore, the pressure side 102 defines an outer pressure surface 114 of the airfoil 100, and the suction side 104 defines an outer suction surface 116 of the airfoil 100.

[0038] Reference Figure 3 and Figure 4 Composite component 100 (e.g., Figure 2The illustrated airfoil 100 comprises a composite material reinforced with chopped fibers (e.g., chopped metal fibers), such as a CMC material or a PMC material. In at least some embodiments, the composite component 100 is formed from a plurality of continuous fiber layers 118 and at least one chopped fiber layer 120 co-cured to produce the composite component 100. Each continuous fiber layer 118 comprises a plurality of continuous fibers 122 disposed in a first matrix material 124. In some embodiments, the continuous fibers 122 are one of the fiber materials described with respect to PMC materials, such as carbon, glass, and / or polymers (e.g., The first matrix material 124 is, for example, the resin described above with respect to PMC materials. In other embodiments, the continuous fiber 122 is ceramic fiber as described with respect to CMC materials, and the first matrix material 124 is, for example, silicon carbide (SiC), silicon, silicon dioxide, carbon, or alumina matrix material, as described above with respect to CMC materials.

[0039] Each chopped fiber layer 120 includes a film sheet 128 ( Figure 5A Multiple chopped fibers 126. For example, such as... Figure 5A As shown, the film sheet 128 is formed from a plurality of randomly oriented chopped fibers 126 injected with a second matrix material 130. The film sheet 128 may be formed as a single chopped fiber layer 120, or the film sheet 128 may be cut to form two or more chopped fiber layers 120.

[0040] The film sheet 128 can be formed in a manner similar to a prepreg layer. For example, in some embodiments, a second matrix material 130 flows around the chopped fibers 126 and is dried. Once the second matrix material 130 is dry, the film sheet 128 can be cut into two or more chopped fiber layers 120 or can be stacked as a single chopped fiber layer 120 with a continuous fiber layer 118.

[0041] The chopped fibers 126 may be made of a variety of materials, and the film sheet 128 may comprise chopped fibers 126 formed of one material or a combination of two or more materials. In some embodiments, the chopped fibers 126 are metallic chopped fibers. For example, the metallic chopped fibers 126 are at least one of titanium, aluminum, and nickel fibers. In other embodiments, the chopped fibers 126 are non-metallic fibers, such as ceramic fibers or ceramic alloy fibers. For example, ceramic or ceramic alloy fibers may comprise silicon and, in some embodiments, may be formed of sapphire, silicon carbide, aluminum silicate, and / or carbon or combinations thereof. The second matrix material 130 may be one of the various matrix materials described herein and may be the same as or different from the first matrix material 124 of the continuous fiber layer 118.

[0042] The chopped fibers 126 of the chopped fiber layer 120 are not continuous from one end to the opposite end of the chopped fiber layer 120. Instead, the chopped fibers 126 are typically of relatively short length, such that they do not extend the full length of the chopped fiber layer 120. The chopped fibers 126 are randomly oriented within the film sheet 128. The multiple chopped fibers 126 within the film sheet 128 do not need to have the same length; for example, chopped fibers 126 of various lengths can be used to form the film sheet 128. Therefore, unlike the continuous fibers 122 of the continuous fiber layer 118, the chopped fibers 126 do not extend end-to-end along the chopped fiber layer 120, but rather extend randomly in various or multiple directions relative to the respective chopped fiber layer 120.

[0043] like Figure 3 and Figure 4 As shown, for at least some composite components 100, for example Figure 2 The airfoil 100 depicted in the figure has a plurality of chopped fiber layers 120 stacked with a plurality of continuous fiber layers 118, such that the plurality of chopped fiber layers 120 are dispersed within the plurality of continuous fiber layers 118. In some embodiments, such as Figure 3 As shown, the chopped fiber layer 120 can alternate with the continuous fiber layer 118, such that the continuous fiber layer 118 is disposed between each chopped fiber layer 120. In other embodiments, such as Figure 4 As shown, more than one chopped fiber layer 120 can be laid between continuous fiber layers 118. For example, two chopped fiber layers 120 can be provided between continuous fiber layers 118, such as... Figure 4 As shown, but in other embodiments, three, four or more chopped fiber layers 120 may be provided between the first continuous fiber layer 118 and the second continuous fiber layer 118.

[0044] It should be understood that Figure 3 and Figure 4 The schematic diagram shows only a portion of the laminate of the composite component (e.g., airfoil 100). Additional continuous fiber layers 118 and chopped fiber layers 120 may be used to form the composite component 100. Figure 3 and Figure 4 Showing continuous fiber layer 118 and chopped fiber layer 120 (i.e., Figure 3 Alternating layers 118, 120 and Figure 4Some embodiments illustrate the relative positions of the multilayers 120 (between the middle layers 118) in the stack. In other embodiments, the chopped fiber layers 120 may be disposed on the inner and / or outer surfaces of the composite component, such that the chopped fiber layers 120 are not between two or more continuous fiber layers 118, but rather as an outer or surface layer. In other embodiments, only one chopped fiber layer 120 may be stacked with multiple continuous fiber layers 118 to form a composite component. Other methods of positioning the continuous fiber layers 118 and chopped fiber layers 120 relative to each other may also be used.

[0045] In at least some embodiments, the fiber density varies among the plurality of chopped fiber layers 120 of the composite component 100 and / or within a single chopped fiber layer 120. Fiber density can be understood as the concentration of chopped fibers 126 within a segment of a chopped fiber layer 120 or a chopped fiber layer 120, or the relative proportion of chopped fibers 126 and the second matrix material 130 within a chopped fiber layer 120. For example, a higher fiber density chopped fiber layer 120 has a higher ratio of chopped fibers 126 to the second matrix material 130 compared to a lower fiber density chopped fiber layer 120.

[0046] In some embodiments, the fiber diameter varies among the chopped fibers 126 of a plurality of chopped fiber layers 120 or within a single chopped fiber layer 120 to change the fiber density. That is, as Figure 5B As shown, some chopped fibers 126 have a smaller diameter d than those with smaller diameter d. s Other short-cut fibers have a larger diameter d l A higher fiber density chopped fiber layer 120 can have a relatively small diameter d. s More larger diameters d l Short-cut fibers 126; short-cut fiber layers 120 with lower fiber density can have a relatively large diameter d l More smaller diameters d s The chopped fibers 126. Comparatively, the first chopped fiber layer 120, with its higher fiber density, can have a larger diameter d than the second chopped fiber layer 120. l Short-cut fibers 126.

[0047] In other embodiments, the fiber length varies among the chopped fibers 126 of a plurality of chopped fiber layers 120 or within a single chopped fiber layer 120 to change the fiber density. That is, as... Figure 5B As shown, some chopped fibers 126 have a shorter length than those with a shorter length l s Other short-cut fibers 126 longer length l l A higher fiber density chopped fiber layer 120 can have a relatively short length l s More longer lengths lShort-cut fibers 126; short-cut fiber layers 120 with lower fiber density can have relatively long lengths l l More shorter lengths l s The chopped fibers 126. Comparatively, the first chopped fiber layer 120, with its higher fiber density, can have more longer lengths l than the second chopped fiber layer 120. l Short-cut fibers 126.

[0048] In other embodiments, the areal weight varies among the chopped fibers 126 of the plurality of chopped fiber layers 120 or within a single chopped fiber layer 120 to alter the fiber density. For example, one or more chopped fiber layers 120 may have a higher concentration of chopped fibers 126, such that the weight of chopped fibers 126 over a given area is higher than that of other chopped fiber layers 120 having a lower concentration of chopped fibers over the same area. As another example, one or more chopped fiber layers 120 may be formed from a combination of two or more types of chopped fibers 126, such as chopped fibers 126 made of a first material (referred to as "first chopped fiber 126") and chopped fibers 126 made of a second material (referred to as "second chopped fiber 126"). The first material may have a greater weight than the second material, such that a chopped fiber layer 120 containing a larger percentage of first chopped fibers 126 made of the first material has a greater areal weight than a chopped fiber layer 120 containing a larger percentage of second chopped fibers 126 made of the second material. In other words, the first chopped fiber layer 120 has a larger ratio of first chopped fibers to second chopped fibers than the second chopped fiber layer 120, has a larger area weight, and therefore has a larger fiber density than the second chopped fiber layer 120.

[0049] It should be understood that chopped fiber layers 120 with a greater fiber density can be used, for example, to provide greater reinforcement in areas more susceptible to damage or with a higher probability of delamination. As an example, multiple chopped fiber layers 120 disposed between multiple consecutive fiber layers 118 within the tip section 132 of the illustrated airfoil 100 can increase the fiber density along the wingspan S. Figure 2 In other words, the lower fiber density chopped fiber layer 120 can be positioned closer to the root 106 compared to the higher fiber density chopped fiber layer 120, while the higher fiber density chopped fiber layer 120 can be positioned closer to the tip 108. As another example, the higher fiber density chopped fiber layer 120 can be positioned closer to the outer surface of the composite component 100 (e.g., the outer pressure surface 114 and / or outer suction surface 116 of the airfoil 100) compared to the lower fiber density chopped fiber layer 120. It should be understood that the chopped fibers 126 are reinforcing fibers, such that the higher density of the chopped fibers 126 (e.g., compared to the second matrix material 130) can provide greater reinforcement.

[0050] The location of one or more chopped fiber layers 120 within the composite component 100 can be biased toward one or more segments of the component to provide reinforcement in specific areas, which can be identified through analysis, testing, etc. For example, refer to the reference. Figure 2 In some embodiments of the airfoil 100, one or more chopped fiber layers 120 are disposed adjacent to the tip 108 to reinforce the tip 108 of the airfoil 100. For example, multiple chopped fiber layers 120 may be dispersed within a continuous fiber layer 118 in the tip segment 132, while the remainder of the airfoil 100 is formed solely of the continuous fiber layer 118. For example, as... Figure 2 As shown, the tip segment 132 extends axially along at least a portion of the chord length c and radially along a portion of the wingspan S adjacent to the tip 108. Exemplary axial positions of the tip segment 132, measured axially from the leading edge 110 to the trailing edge 112 (i.e., the leading edge 110 is at 0% of the chord length c and the trailing edge 112 is at 100% of the chord length c), may include, but are not limited to, from about 0% to about 30% of the chord length c (e.g., adjacent to the leading edge 110); from about 70% to about 100% of the chord length c (e.g., adjacent to the trailing edge 112); and from 0% to 100% of the chord length c (i.e., along the entire chord length c). Exemplary radial positions of the tip segment 132, measured radially from the root 106 (i.e., the root 106 is located at 0% of the wingspan S and the tip 108 is located at 100% of the wingspan S), may include, but are not limited to, from about 50% to 100% of the wingspan S (i.e., from about the middle wingspan to the tip 108); from about 60% to 100% of the wingspan S; from about 75% to 100% of the wingspan S; from about 80% to 100% of the wingspan S; from about 85% to 100% of the wingspan S; from about 90% to 100% of the wingspan S; and from about 95% to 100% of the wingspan S.

[0051] See Figure 6 In other embodiments, one or more chopped fiber layers 120 are disposed adjacent to the trailing edge 112 to reinforce the trailing edge 112 of the airfoil 100. For example, a plurality of chopped fiber layers 120 may be dispersed within a continuous fiber layer 118 in the trailing edge segment 134 (e.g., as shown in the figure). Figure 3 and 4 As shown, or any other suitable dispersion), while the remainder of the airfoil 100 is formed solely of a continuous fiber layer 118. For example, as Figure 6As shown, the airfoil 100 includes a trailing edge segment 134 that extends axially along a portion of the chord length c and radially along at least a portion of the wingspan S adjacent to the trailing edge 112. Exemplary axial positions of the trailing edge segment 134, measured axially from the leading edge 110 to the trailing edge 112 (i.e., the leading edge 110 at 0% of the chord length c and the trailing edge 112 at 100% of the chord length c), may include, but are not limited to, from about 60% to about 100% of the chord length c; from about 70% to about 100% of the chord length c; from about 75% to about 100% of the chord length c; from about 80% to about 100% of the chord length c; from about 85% to about 100% of the chord length c; from about 90% to about 100% of the chord length c; and from about 95% to about 100% of the chord length c. Exemplary radial positions of the trailing edge segment 134, measured radially from the root 106 (i.e., the root 106 is located at 0% of the wingspan S and the tip 108 is located at 100% of the wingspan S), may include, but are not limited to, from about 0% to 100% of the wingspan S (i.e., from the root 106 to the tip 108); from about 50% to 100% of the wingspan S (i.e., from about the middle of the wingspan to the tip 108); and from about 75% to 100% of the wingspan S.

[0052] See Figure 7 In other embodiments, one or more chopped fiber layers 120 are disposed adjacent to the leading edge 110 to reinforce the leading edge 110 of the airfoil 100. For example, a plurality of chopped fiber layers 120 may be dispersed within a continuous fiber layer 118 in the leading edge portion 136 (e.g., as shown in the image). Figure 3 and 4 (as shown or any other suitable dispersion), while the remainder of the airfoil 100 is formed solely of a continuous fiber layer 118. Similar to the tip section 132 and the trailing edge section 134, Figure 7The airfoil 100 includes a leading edge segment 136 that extends axially along a portion of the chord length c and radially along at least a portion of the wingspan S adjacent to the leading edge 110. Exemplary axial positions of the leading edge segment 136 (i.e., the leading edge 110 at 0% of the chord length c and the trailing edge 112 at 100% of the chord length c), measured axially from the leading edge 110 to the trailing edge 112, may include, but are not limited to, from about 0% to about 40% of the chord length c; from about 0% to about 30% of the chord length c; from about 0% to about 25% of the chord length c; from about 0% to about 20% of the chord length c; from about 0% to about 15% of the chord length c; from about 0% to about 10% of the chord length c; and from about 0% to about 5% of the chord length c. Exemplary radial positions of the leading edge segment 136, measured radially from the root 106 (i.e., the root 106 is located at 0% of the wingspan S and the tip 108 is located at 100% of the wingspan S), may include, but are not limited to, from about 0% to 100% of the wingspan S (i.e., from the root 106 to the tip 108); from about 50% to 100% of the wingspan S (i.e., from about the middle wingspan to the tip 108); and from about 75% to 100% of the wingspan S.

[0053] See now Figure 8 In a further embodiment, multiple chopped fiber layers 120 are disposed within multiple continuous fiber layers 118 adjacent to various portions of the airfoil 100 to reinforce the various segments and / or features of the airfoil 100. For example, as Figure 8 As shown, the airfoil 100 includes, as per the description... Figure 2 and Figure 6 The described tip segment 132 and trailing edge segment 134. Therefore, for the depicted embodiment, a plurality of chopped fiber layers 120 are disposed within region 138, which extends axially along at least a portion of the chord length c near the tip 108 and radially along at least a portion of the span S near the trailing edge 112. The axial and radial lengths of region 138 can be described above regarding... Figure 2 and 6 The axial position of the tip section 132 and the radial position of the trailing edge section 134 are described within the range.

[0054] like Figure 8 As further shown, the airfoil 100 includes a leading edge segment 136. In some embodiments, Figure 8 The leading edge segment 136 may include one or more chopped fiber layers 120 disposed between a plurality of continuous fiber layers 118, for example, as per [reference to...]. Figure 7 As described. In other embodiments, Figure 8The leading edge segment 136 may include a protective element, such as a metal cap 142, disposed on the leading edge 110, instead of reinforcing the leading edge segment 136 with chopped fiber layers 120. It should be understood that in some embodiments, the metal cap 142 may be used in conjunction with one or more chopped fiber layers 120 embedded in the composite component 100 (e.g., the depicted airfoil 100) to further reinforce one or more segments of the composite component 100 (e.g., the leading edge 110 and / or the trailing edge 112 of the airfoil 100).

[0055] Furthermore, the dispersion of the chopped fiber layers 120 can vary across the entire thickness of the composite component 100. In some embodiments, the plurality of chopped fiber layers 120 are biased toward the outer surface of the airfoil 100, such as the outer pressure surface 114 and / or the outer suction surface 116, such that at least a majority of the plurality of chopped fiber layers 120 are arranged between a plurality of continuous fiber layers 118 near the outer surface. In other words, the chopped fiber content of the composite component 100 can be graded such that it gradually decreases from the outer surface of the component 100 inwards. For example, regarding... Figure 2 In the described tip section 132, a plurality of chopped fiber layers 120 may be biased toward the outer pressure surface 114 and the outer suction surface 116 in the tip section 132, such that the chopped fiber layers 120 in the tip section 132 are disposed primarily toward the outer pressure surface 114 and the outer suction surface 116 of the airfoil 100. As another example, a plurality of chopped fiber layers 120 may be disposed throughout the airfoil 100 (i.e., not only in one or more of the tip section 132, trailing edge section 134, and leading edge section 136) among a plurality of continuous fiber layers 118, but the chopped fiber layers 120 may be biased toward one or both of the outer pressure surface 114 and the outer suction surface 116, such that the chopped fiber layers 120 are concentrated at or near the outer surface of the airfoil 100. It should be understood that concentrating the chopped fiber layers 120 at or near the outer surface of the composite component 100 (e.g., the depicted airfoil 100) can provide targeted reinforcement. For example, a component may typically be subjected to damaging forces along its outer surface rather than its interior, such that a chopped fiber layer 120 at or near the outer surface helps prevent or reduce damage to the most vulnerable areas without having to distribute the chopped fiber layer 120 throughout the component.

[0056] Of course, it should be understood that the chopped fiber layer 120 can also be dispersed in a more targeted manner within the continuous fiber layer 118. For example, instead of distributing the multiple chopped fiber layers 120 across the entire outer surface of the composite component 100, the chopped fiber layers 120 or multiple chopped fiber layers 120 can be disposed at or near one or more portions or segments of the outer surface rather than across the entire outer surface. As several non-limiting examples, the plurality of chopped fiber layers 120 may be biased toward the outer pressure surface 114 of the airfoil 100 instead of both the outer pressure surface 114 and the outer suction surface 116; the plurality of chopped fiber layers 120 may be biased toward the outer suction surface 116 of the airfoil 100 instead of both the outer pressure surface 114 and the outer suction surface 116; the plurality of chopped fiber layers 120 may be biased toward one or both of the outer pressure surface 114 and the outer suction surface 116 of the airfoil 100 from a mid-span radial position to the tip 108 (e.g., from about 50% to about 100% of the wingspan S); or the plurality of chopped fiber layers 120 may be biased toward the outer suction surface 116 within a mid-span region of the airfoil (e.g., from about 40% to about 60% of the wingspan S).

[0057] Now for reference Figure 9 A flowchart is provided illustrating method 900 for forming composite components of a gas turbine engine. (Example) Figure 9 As shown, method 900 includes laying (902) a plurality of continuous fiber layers 118. As described herein, each continuous fiber layer 118 includes continuous fibers 122 disposed in a first matrix material 124. In various embodiments, the continuous fibers 122 and the first matrix material 124 may be any of the various fiber and matrix materials described herein with respect to polymer matrix composite (PMC) materials or ceramic matrix composite (CMC) materials.

[0058] Method 900 further includes forming (904) one or more chopped fiber layers 120. In some embodiments, each chopped fiber layer 120 includes a plurality of chopped metal fibers 126, said plurality of chopped metal fibers 126 being infused with a second matrix material 130 to form a film sheet 128. In other embodiments, each chopped fiber layer 120 includes a plurality of non-metallic chopped fibers 126, such as ceramic chopped fibers, which are infused with the second matrix material 130 to form a film sheet 128. As described herein, one or more chopped fiber layers 120 may be formed from film sheets 128, for example, each film sheet 128 may form a single chopped fiber layer 120 or the film sheet 128 may be cut to form two or more chopped fiber layers 120.

[0059] like Figure 9As shown, method 900 further includes stacking one or more chopped fiber layers 120 with a plurality of continuous fiber layers 118 (906). The stack of the plurality of continuous fiber layers 118 and one or more chopped fiber layers 120 forms a reinforcing stack 140. Figure 3 , 4 That is, the stack of continuous fiber layers 118 is reinforced with one or more chopped fiber layers 120, such that the stack of layers 118 and 120 can be referred to as a reinforced stack 140.

[0060] Still refer to Figure 9 Method 900 further includes heat-treating and / or chemically treating the reinforcing laminate 140 to form the composite component 100. For example, method 900 may include compacting (908) and curing (910) the reinforcing laminate 140 in an autoclave at elevated temperatures and elevated pressures to adhere or laminate multiple layers together, including laminating a continuous fiber layer 118 onto one or more chopped fiber layers 120. Thus, the continuous fiber layer 118 and the chopped fiber layer 120 are co-cured to form the composite component 100. As described herein, the first matrix material 124 of the continuous fiber layer 118 and the second matrix material 130 of the chopped fiber layer 120 may be the same matrix material or different matrix materials. Co-curing the continuous fiber layer 118 with one or more chopped fiber layers 120 can promote blending and / or sharing of the first matrix material 124 and the second matrix material 130 between the continuous fiber layer 118 and the chopped fiber layer 120. Furthermore, it should be understood that the elevated temperatures are above ambient temperatures and the elevated pressures are above ambient pressures.

[0061] In some embodiments, the composite component 100 may undergo further processing after the autoclave process, such as densification and finishing processes. As described herein, for CMC components 100, such as CMC airfoils 100, the composite airfoil stack can be compacted and, if appropriate, cured while subjected to elevated pressure and temperature to produce a cured preform. For example, the stack or preform can be cured in an autoclave to form an autoclave body. In some embodiments, the autoclave body is then heated (fired) in a vacuum or inert atmosphere to decompose the binder, remove the solvent, and transform the precursor into the desired ceramic matrix material. Due to the decomposition of the binder, the result of the preform is a porous CMC sintered body, which can be densified, for example, by melt infiltration (MI), to fill the pores and produce the corresponding CMC component. The specific processing techniques and parameters for the heat treatment and / or chemical treatment of the airfoil stack will depend on the specific composition of the material. As an example, other known methods or techniques for curing the composite layer and for densifying the CMC component can be used.

[0062] Furthermore, method 900 may include one or more protective elements, such as Figure 8 The metal cap 142 described herein, applied to the leading edge segment 136, is applied to or attached (912) to the composite component 100. As described herein, the protective element may be applied to a portion of the composite component 100 that does not have the chopped fiber layer 120, or the protective element may be used in conjunction with one or more chopped fiber layers 120, for example, the metal cap 142 may be applied to a portion of the composite component 100 that includes one or more chopped fiber layers 120.

[0063] Therefore, as described herein, this subject matter provides composite components having at least one chopped fiber layer, such as composite airfoils including gas turbine engine fan blades, and methods for forming composite components having at least one chopped fiber layer. For example, this subject matter provides composite components comprising one or more chopped fiber layers, each comprising a plurality of chopped fibers formed as a thin film sheet, wherein the one or more chopped fiber layers are stacked with and co-cured with continuous fiber layers, each continuous fiber layer comprising a plurality of continuous fibers. The chopped fiber layers can help reinforce the composite component, and in at least some embodiments, targeted reinforcement can be provided by stacking with continuous fiber layers in certain target areas rather than throughout the entire component. The benefits of this subject matter include reducing or eliminating separate metal tip caps, leading-edge guards, and / or trailing-edge guards; reducing or eliminating debonding or peeling between the metal tip cap or guard and the composite component; eliminating field replacement of tip caps; reducing or eliminating separate bonding operations during manufacturing (i.e., bonding the metal tip cap or guard to the composite component); optimizing material use in target areas (e.g., optimizing chopped fiber reinforcement in target fan blade areas); and reducing the weight and cost of each component (e.g., each fan blade) by removing the metal cladding. Other advantages of the subject matter described herein will also be appreciated by those skilled in the art.

[0064] Further aspects of this disclosure are provided by the subject matter of the following clauses:

[0065] 1. A composite airfoil for a gas turbine engine, comprising: opposing pressure and suction sides extending radially from a root to a tip along the wingspan, the root defining a first radial end of the composite airfoil and the tip defining a second radial end of the composite airfoil; opposing leading and trailing edges extending radially along the wingspan, the pressure and suction sides extending axially between the leading and trailing edges; a plurality of continuous fiber layers, each continuous fiber layer comprising a plurality of continuous fibers; and a chopped fiber layer comprising a plurality of chopped fibers formed as a thin film sheet, wherein the chopped fiber layer is stacked with the plurality of continuous fiber layers to form the composite airfoil.

[0066] 2. The composite airfoil according to any of the preceding clauses, wherein the plurality of chopped fibers are randomly oriented in the film sheet.

[0067] 3. The composite airfoil according to any of the preceding clauses, wherein the plurality of continuous fibers in each of the plurality of continuous fiber layers are disposed in a first matrix material, and wherein the film sheet comprises the plurality of chopped fibers injected with a second matrix material.

[0068] 4. The composite airfoil according to any of the preceding clauses, wherein the second matrix material is the same as the first matrix material.

[0069] 5. The composite airfoil according to any of the preceding clauses, further comprising: a plurality of chopped fiber layers, wherein the chopped fiber layer is one of the plurality of chopped fiber layers, and wherein the plurality of chopped fiber layers are stacked with the plurality of continuous fiber layers such that the plurality of chopped fiber layers are dispersed within the plurality of continuous fiber layers.

[0070] 6. The composite airfoil according to any of the preceding clauses, further comprising: a plurality of chopped fiber layers, wherein the chopped fiber layer is one of the plurality of chopped fiber layers, and wherein the fiber diameter varies among the chopped fibers of the plurality of chopped fiber layers to change the fiber density among the plurality of chopped fiber layers.

[0071] 7. The composite airfoil according to any of the preceding clauses, further comprising: a plurality of chopped fiber layers, wherein the chopped fiber layer is one of the plurality of chopped fiber layers, and wherein the fiber length varies among the chopped fibers of the plurality of chopped fiber layers to change the fiber density among the plurality of chopped fiber layers.

[0072] 8. The composite airfoil according to any of the preceding clauses, further comprising: a plurality of chopped fiber layers, wherein the chopped fiber layer is one of the plurality of chopped fiber layers, and wherein the areal weight varies among the chopped fibers of the plurality of chopped fiber layers to change the fiber density among the plurality of chopped fiber layers.

[0073] 9. The composite airfoil according to any of the preceding clauses, wherein the chopped fiber layer is disposed adjacent to the tip to reinforce the tip of the composite airfoil.

[0074] 10. The composite airfoil according to any of the preceding clauses, wherein the chopped fiber layer is disposed adjacent to the trailing edge to reinforce the trailing edge of the composite airfoil.

[0075] 11. The composite airfoil according to any of the preceding clauses, further comprising: a plurality of chopped fiber layers, wherein the chopped fiber layer is one of the plurality of chopped fiber layers, and wherein the plurality of chopped fiber layers are biased toward the outer surface of the composite airfoil such that at least a majority of the plurality of chopped fiber layers are disposed between the plurality of continuous fiber layers near the outer surface.

[0076] 12. The composite airfoil according to any of the preceding clauses, wherein the chopped fiber layer is disposed on the outer surface of the composite airfoil to define the outer layer of the airfoil.

[0077] 13. The composite airfoil according to any of the preceding clauses, wherein the chopped fiber layer is disposed on the inner surface of the airfoil to define the surface layer of the airfoil.

[0078] 14. The composite airfoil according to any of the preceding clauses, wherein the plurality of chopped fibers are metal fibers.

[0079] 15. The composite airfoil according to any of the preceding clauses, wherein the metal fiber is at least one of titanium, aluminum and nickel fibers.

[0080] 16. The composite airfoil according to any of the preceding clauses, wherein the plurality of chopped fibers are ceramic fibers or ceramic alloy fibers.

[0081] 17. The composite airfoil according to any of the preceding clauses, wherein the ceramic or ceramic alloy fiber comprises silicon.

[0082] 18. A method for forming a composite component of a gas turbine engine, the method comprising: laying a plurality of continuous fiber layers, each continuous fiber layer comprising a plurality of continuous fibers disposed in a first matrix material; stacking a chopped fiber layer with the plurality of continuous fiber layers, the plurality of continuous fiber layers and the chopped fiber layer forming a reinforcing stack; and curing the reinforcing stack.

[0083] 19. The method according to any of the preceding clauses, further comprising: forming the chopped fiber layer, wherein the chopped fiber layer comprises a plurality of chopped metal fibers injected with a second matrix material to form a thin film sheet.

[0084] 20. The method according to any of the preceding clauses, wherein the first matrix material and the second matrix material are the same matrix material.

[0085] 21. The method according to any of the preceding clauses, wherein the first matrix material and the second matrix material are different matrix materials.

[0086] 22. The method according to any of the preceding clauses further includes forming a plurality of chopped fiber layers, wherein the chopped fiber layer is one of the plurality of chopped fiber layers.

[0087] 23. The method according to any of the preceding clauses, wherein the plurality of chopped fiber layers are formed of one or more film sheets, each film sheet comprising a plurality of chopped metal fibers injected with a matrix material.

[0088] 24. The method according to any of the preceding clauses further includes cutting the film sheet to form two or more chopped fiber layers, wherein the chopped fiber layer is one of the two or more chopped fiber layers.

[0089] 25. The method according to any of the preceding clauses, wherein curing the reinforcing laminate comprises compacting and curing the reinforcing laminate in an autoclave at elevated temperature and elevated pressure.

[0090] 26. The method according to any of the preceding clauses, wherein the composite component is a composite airfoil reinforced with chopped metal fibers.

[0091] 27. A composite component for a gas turbine engine, comprising: a plurality of continuous fiber layers, each continuous fiber layer being formed of a plurality of continuous fibers disposed in a matrix material; and a chopped fiber layer comprising a plurality of chopped fibers formed as a thin film sheet, wherein the chopped fiber layer is stacked with the plurality of continuous fiber layers to form the composite component.

[0092] This written description uses examples to disclose embodiments of this disclosure, including best practices, and also enables any person skilled in the art to practice this disclosure, including making and using any device or system and performing any combined methods. The patentable scope of this disclosure is defined by the claims and may include other examples that would occur to a person skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that are not indistinguishable from the literal language of the claims, or if they include equivalent structural elements that are not substantially different from the literal language of the claims.

Claims

1. A composite airfoil for a gas turbine engine, characterized by, include: Opposite pressure side and suction side, the pressure side and the suction side extend radially along the wingspan from root to tip, the root defining a first radial end of the composite airfoil, and the tip defining a second radial end of the composite airfoil; Opposite leading and trailing edges, the leading and trailing edges extending radially along the wingspan, the pressure side and the suction side extending axially between the leading and trailing edges; Multiple continuous fiber layers, each continuous fiber layer containing multiple continuous fibers; and A chopped fiber layer comprising a plurality of chopped fibers formed into a thin film sheet. The chopped fiber layer is stacked with the plurality of continuous fiber layers to form the composite airfoil.

2. The composite airfoil according to claim 1, characterized in that, The plurality of chopped fibers are randomly oriented in the film sheet.

3. The composite airfoil according to claim 2, characterized in that, The plurality of continuous fibers in each of the plurality of continuous fiber layers are disposed in a first matrix material, and the film sheet includes the plurality of chopped fibers injected with a second matrix material.

4. The composite airfoil according to claim 3, characterized in that, The second matrix material is the same as the first matrix material.

5. The composite airfoil according to claim 1, characterized in that, Further includes: Multiple short-cut fiber layers, The chopped fiber layer is one of the plurality of chopped fiber layers, and The plurality of chopped fiber layers are stacked with the plurality of continuous fiber layers, such that the plurality of chopped fiber layers are dispersed within the plurality of continuous fiber layers.

6. The composite airfoil according to claim 1, characterized in that, Further includes: Multiple short-cut fiber layers, The chopped fiber layer is one of the plurality of chopped fiber layers, and The fiber diameter varies among the chopped fibers in the plurality of chopped fiber layers to change the fiber density among the plurality of chopped fiber layers.

7. The composite airfoil according to claim 1, characterized in that, Further includes: Multiple short-cut fiber layers, The chopped fiber layer is one of the plurality of chopped fiber layers, and The fiber length varies among the chopped fibers in the plurality of chopped fiber layers to change the fiber density among the plurality of chopped fiber layers.

8. The composite airfoil according to claim 1, characterized in that, Further includes: Multiple short-cut fiber layers, The chopped fiber layer is one of the plurality of chopped fiber layers, and The area weight varies among the chopped fibers in the plurality of chopped fiber layers to change the fiber density among the plurality of chopped fiber layers.

9. The composite airfoil according to claim 1, characterized in that, The chopped fiber layer is positioned adjacent to the tip to reinforce the tip of the composite airfoil.

10. The composite airfoil according to claim 1, characterized in that, The chopped fiber layer is disposed adjacent to the trailing edge to reinforce the trailing edge of the composite airfoil.

11. The composite airfoil according to claim 1, characterized in that, Further includes: Multiple short-cut fiber layers, The chopped fiber layer is one of the plurality of chopped fiber layers, and The plurality of chopped fiber layers are biased toward the outer surface of the composite airfoil such that at least a majority of the chopped fiber layers are disposed between the plurality of continuous fiber layers near the outer surface.

12. The composite airfoil according to claim 1, characterized in that, The plurality of chopped fibers are metal fibers.

13. The composite airfoil according to claim 12, characterized in that, The metal fiber is at least one of titanium, aluminum and nickel fibers.

14. The composite airfoil according to claim 1, characterized in that, The plurality of chopped fibers are ceramic fibers or ceramic alloy fibers.

15. The composite airfoil according to claim 14, characterized in that, The ceramic fiber or ceramic alloy fiber contains silicon.

16. A method for forming a composite airfoil for a gas turbine engine, characterized in that, The method includes: Multiple continuous fiber layers are laid, each continuous fiber layer containing multiple continuous fibers disposed in a first matrix material; The chopped fiber layer is stacked with the plurality of continuous fiber layers, and the plurality of continuous fiber layers and the chopped fiber layer form a reinforcing laminate; and The reinforcing stack is cured to form the composite airfoil.

17. The method according to claim 16, characterized in that, Further includes: Forming the chopped fiber layer, The chopped fiber layer comprises a plurality of chopped metal fibers injected with a second matrix material to form a thin film sheet.

18. The method according to claim 16, characterized in that, The curing of the reinforcing layer includes compacting and curing the reinforcing layer in an autoclave under elevated temperature and pressure.

19. The method according to claim 16, characterized in that, The composite airfoil is a composite airfoil reinforced with chopped metal fibers.

20. A composite airfoil for a gas turbine engine, characterized in that, include: Multiple continuous fiber layers, each continuous fiber layer being formed by multiple continuous fibers disposed in a matrix material; and A chopped fiber layer comprising a plurality of chopped fibers formed into a thin film sheet. The chopped fiber layer is stacked with the plurality of continuous fiber layers to form the composite airfoil.

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