METHOD FOR FORMING COMPOSITE COMPONENTS HAVING CAVITIES

Abstract

A method (200) for forming a composite component (100) includes stacking a plurality of composite plies (124) around a mandrel assembly (150) to form a composite plie core (122). The mandrel assembly (150) includes a mandrel (152) having an internal structural body (154) and an external release layer (156). The mandrel (152) is configured to form a cavity (116) in the composite plie core (122). The method (200) also includes processing the composite plie core (122) containing the mandrel assembly (150) to compact the plurality of composite plies (124) together. Furthermore, the method (200) includes withdrawing the mandrel assembly (150) from the composite plie core (122) to form the cavity (116). Figure for the abstract: Fig 2

Description

Title of the invention: METHOD FOR FORMING COMPOSITE COMPONENTS HAVING CAVITIES Technical field

[0001] The present disclosure relates generally to composite turbomachinery components and more particularly to methods of forming composite components having cavities, such as cooling cavities, in rotor blades and stator vanes of turbomachinery. TECHNICAL CONTEXT

[0002] A gas turbine engine generally includes a fan and a core arranged in fluid communication with each other. Further, the core of the gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is supplied from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and combusted in the combustion section to provide combustion gases. The combustion gases are conveyed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then conveyed through the exhaust section, for example, to the atmosphere.

[0003] In general, turbine performance and efficiency can be improved by increasing combustion gas temperatures. However, increased combustion temperatures can negatively impact gas turbine engine components, for example, by increasing the likelihood of hardware failures. Thus, while increased combustion temperatures may be beneficial to turbine performance, certain gas turbine engine components may require cooling features or reduced exposure to combustion gases to reduce the negative impacts of increased temperatures on the components.

[0004] Typically, the turbine section includes one or more stages of stator and rotor blade vanes, and each stage of stator and rotor blade vanes includes a plurality of airfoils, e.g., nozzle airfoils in the stator blade portion and blade airfoils in the rotor blade portion. Because the airfoils are downstream of the combustion section and positioned in the combustion gas flow, the airfoils Airfoils typically include one or more cooling features to minimize the effects of relatively hot combustion gases, such as, for example, film holes, cooling holes, or slots, which may provide cooling within or on the surface of the airfoils. For example, cooling apertures may be provided throughout a component that allow a flow of coolant from within the component to be directed onto the exterior surface of the component. In addition, airfoils typically include cavities or ducts to provide compressed cold air to the cooling features, for example, from the compressor section. DESCRIPTION OF FIGURES

[0005] A complete and enabling disclosure of the present disclosure, including the best mode thereof, addressed to a person skilled in the art, is set forth in the specification, which refers to the appended figures, in which:

[0006] [Fig.l] is a schematic sectional view of a gas turbine engine, in accordance with one embodiment of the present disclosure;

[0007] [Fig.2] is a perspective view of an embodiment of a component of a gas turbine engine according to aspects of the present disclosure, particularly illustrating the component configured as a turbine rotor blade;

[0008] [Fig.3] is a perspective view of another embodiment of a component of a gas turbine engine according to aspects of the present disclosure, particularly illustrating the component configured as a stator vane;

[0009] [Fig.4A], [Fig.4B] and [Fig.4C] are various sectional views of the component of the [Fig.2] according to the aspects of the present disclosure;

[0010] [Fig.5] is a schematic diagram of one embodiment of a ply core composites according to aspects of the present disclosure, particularly illustrating the composite ply core comprising composite plies;

[0011] [Fig.6] is a flowchart of one embodiment of a forming method of a composite component according to aspects of the present disclosure;

[0012] [Fig.7] is a flowchart of another embodiment of a method of forming a composite component according to aspects of the present disclosure;

[0013] [Fig.8A] is a perspective view of one embodiment of a chuck assembly to form a cooling cavity in a component of a gas turbine engine according to aspects of the present disclosure;

[0014] [Fig.8B] is a sectional view of the mandrel assembly of [Fig.8A] along section line 8B-8B;

[0015] [Fig.9] is a flowchart of one embodiment of a forming method of a chuck assembly according to aspects of the present disclosure;

[0016] [Fig.lOA], [Fig.lOB], [Fig.lOC] and [Fig.lOD] illustrate various components of a system for forming a composite component according to aspects of the present disclosure.

[0017] The repeated use of references in this specification and the drawings is intended to represent the same or similar features or elements of the present disclosure. DETAILED DESCRIPTION

[0018] Reference will now be made in detail to embodiments of the disclosure, one or more examples of which are illustrated in the drawings. Each example is provided for the purpose of explaining the disclosure, and not as a limitation of the disclosure. In fact, it will be apparent to those skilled in the art that various modifications and variations may be made to the present disclosure without departing from the scope or spirit of the disclosure. For example, features illustrated or described in one embodiment may be used with another embodiment to produce yet another embodiment.

[0019] The term "by way of example" is used herein to mean "serve as an example, instance, or illustration." Any implementation described herein as "by way of example" should not necessarily be construed as preferred or advantageous over other implementations. Furthermore, unless specifically stated otherwise, all embodiments described herein should be considered exemplary.

[0020] The term "at least one of" in the context of, for example, "at least one of A, B, and C" means only A, only B, only C, or any combination of A, B, and C.

[0021] The term "turbomachine" refers to a machine comprising one or more compressors, a heat generating section (e.g., a combustion section), and one or more turbines that together generate a torque output.

[0022] The term "gas turbine engine" refers to an engine having a turbomachine as all or part of its power source. Exemplary gas turbine engines include turbofan, turboprop, turbojet, turboshaft, and the like, as well as hybrid-electric versions of one or more of these engines.

[0023] The term "combustion section" refers to any heat addition system for a turbomachine. For example, the term combustion section may refer to a section comprising one or more of a combustion assembly deflagration, a rotary detonation combustion assembly, a pulsed detonation combustion assembly, or other suitable heat addition assembly. In some exemplary embodiments, the combustion section may comprise an annular combustion chamber, a tubular combustion chamber, a cannular combustion chamber, a trapped vortex combustor (TVC), or other suitable combustion system, or combinations thereof.

[0024] As used herein, the term "rotor" refers to any component of a rotating machine, such as a turbine engine, that rotates about an axis of rotation. For example, a rotor may comprise a shaft or a coil of a rotating machine, such as a turbine engine.

[0025] As used herein, the term "stator" refers to any component of a rotating machine, such as a turbine engine, that has a coaxial configuration and arrangement with a rotor of the rotating machine. A stator may be disposed radially inward or radially outward along a radial axis relative to at least a portion of a rotor. Additionally, or alternatively, a stator may be axially disposed adjacently to at least a portion of a rotor.

[0026] The terms "low" and "high," or their respective comparative degrees (e.g., with the term plus, where applicable), when used with compressor, turbine, shaft, or spool components, etc., each refer to relative speeds in an engine, unless otherwise specified. For example, a "low turbine" or "low speed turbine" defines a component configured to operate at a rotational speed, such as a maximum permissible rotational speed, lower than a "high turbine" or "high speed turbine" of the engine.

[0027] The terms "forward" and "aft" refer to relative positions in a gas turbine engine or vehicle and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, with respect to a gas turbine engine, forward refers to a position closer to an engine inlet and aft refers to a position closer to a nozzle or exhaust of the engine.

[0028] The terms "upstream" and "downstream" refer to the direction relative to a fluid flow in a fluid path. For example, the term "upstream" refers to the direction from which the fluid flows, and the term "downstream" refers to the direction toward which the fluid flows.

[0029] As used herein, the terms "axial" and "axially" refer to directions and orientations that extend substantially parallel to a centerline of the gas turbine engine. Further, the terms "radial" and "radially » refer to directions and orientations that extend substantially perpendicular to the centerline of the gas turbine engine. Additionally, as used herein, the terms “circumferential” and “circumferentially” refer to directions and orientations that extend arcuately around the centerline of the gas turbine engine.

[0030] The terms "coupled," "attached," "attached to," and the like refer to both direct coupling, attachment, or attachment, as well as indirect coupling, attachment, or attachment through one or more intermediate components or features, unless otherwise specified herein.

[0031] As used herein, the terms "first," "second," and so on, may be used interchangeably to distinguish one component from another and are not intended to signify the location or importance of individual components.

[0032] The term "adjacent" as used herein in reference to two walls or surfaces refers to the two walls or surfaces being in contact with each other, or the two walls or surfaces being separated only by one or more non-structural layers and the two walls or surfaces and the one or more non-structural layers being in a serial contact relationship (i.e., a first wall / surface in contact with the one or more non-structural layers, and the one or more non-structural layers in contact with a second wall / surface).

[0033] As used herein, the terms "integral," "unitary," or "monolithic" as used to describe a structure refer to the structure formed integrally of a continuous material or group of materials without seams, connections, joints, or the like. The integral, unitary structure described herein may be formed by additive manufacturing to have the described structure, or alternatively by a casting process, etc.

[0034] As used herein, the term "composite material" refers to a material produced from two or more constituent materials, at least one of the constituent materials being a non-metallic material. Exemplary composite materials include polymer matrix composites (PMCs), ceramic matrix composites (CMCs), quintuple chopped composite materials, etc.

[0035] As used herein, ceramic matrix composite or "CMC" refers to a class of materials that comprise a reinforcing material (e.g., reinforcing fibers) surrounded by a ceramic matrix phase. Generally, the reinforcing fibers provide structural integrity to the ceramic matrix. Some examples of CMC matrix materials may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), oxide ceramics (e.g., silicon oxycarbides, silicon oxynitrides, aluminum oxide (A12O3), silicon dioxide silicon (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) may also be incorporated into the CMC matrix.

[0036] Some examples of CMC reinforcing fibers may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), non-oxide carbon-based materials (e.g., carbon), oxide ceramics (e.g., silicon oxycarbides, silicon oxynitrides, aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates such as mullite, or mixtures thereof), or mixtures thereof.

[0037] In general, particular CMCs may be designated by their fiber type / matrix type combination. For example, C / SiC for carbon fiber reinforced silicon carbide; SiC / SiC for silicon carbide fiber reinforced silicon carbide, SiC / SiN for silicon carbide fiber reinforced silicon nitride; SiC / SiC-SiN for a silicon carbide fiber reinforced silicon carbide / silicon nitride matrix mixture, etc. In other examples, CMCs may include a matrix and reinforcing fibers comprising oxide-based materials such as aluminum oxide (Al2O3), silicon dioxide (SiO2), aluminosilicates, and mixtures thereof. Aluminosilicates may include crystalline materials such as mullite (3Al2O3-2SiO2), as well as glassy aluminosilicates.

[0038] As used herein, a polymer matrix composite or "PMC" refers to a composite material composed of a variety of short or continuous reinforcing fibers bonded together by a matrix of organic polymers. PMCs are designed to transfer loads between fibers of a matrix. Some of the advantages of PMCs include their light weight, high resistance to abrasion and corrosion, and high stiffness and strength in the direction of their reinforcements.

[0039] In some embodiments, the reinforcing fibers of the CMCs and PMCs may be bundled or coated prior to incorporation into the matrix. For example, bundles of the fibers may be formed as a reinforced ribbon, such as a unidirectional reinforced ribbon. A plurality of the ribbons may be stacked together to form a preform component. The fiber bundles may be impregnated with a slurry composition prior to forming the preform or after forming the preform. The preform may then undergo heat treatment and subsequent chemical treatment to result in a component formed from a material having a desired chemical composition. For example, the preform may undergo curing or combustion to yield a high-carbon residue in the preform, and subsequent melt infiltration with silicon, or curing or pyrolysis to yield a silicon carbide matrix in the preform, and subsequent chemical vapor infiltration with silicon carbide. Additional steps may be taken to enhance densification of the preform, either before or after chemical vapor infiltration, by injecting a liquid resin or polymer into the preform, followed by a heat treatment step to fill voids with silicon carbide.A CMC or PMC material as used herein may be formed using any method known or hereinafter developed, including, but not limited to, melt infiltration, chemical vapor infiltration, polymer impregnation pyrolysis (PIP), or any combination thereof.

[0040] Such materials, as well as certain monolithic ceramics (i.e., ceramic materials without reinforcing material) or polymers, are particularly suitable for higher temperature applications. In addition, these materials are lightweight compared to superalloys, but can still provide strength and durability to the component made from them. Therefore, such materials are currently being considered for many gas turbine components used in higher temperature sections of gas turbine engines, such as airfoils (e.g., turbines and blades), combustors, shrouds, and other similar components, which would benefit from the lighter weight and higher temperature capability that these materials can offer.

[0041] In general, turbine performance and efficiency can be improved by increasing combustion gas temperatures. Non-traditional high-temperature materials, such as ceramic matrix composite (CMC) or polymer matrix composite (PMC) materials, are more commonly used for various components in gas turbine engines. For example, because such materials can withstand relatively extreme temperatures, there is particular interest in replacing components in the combustion gas flow path with these materials.However, even if these components can withstand more extreme temperatures than typical components, these components may still require cooling features or reduced exposure to flue gases to reduce a likelihood of negative impacts from increased flue gas temperatures, e.g., hardware failures or the like.

[0042] Although many advantages can be realized by using high-temperature components, these materials may have disadvantages. For example, some high-temperature components may have lower thermal conductivity than similar components formed, for example, from nickel alloys. The reduced thermal conductivity of high-temperature components may require that cooling features or cold compressed air supply cavities or conduits be located closer to the airfoil surface. In addition, cooling features, cavities, or conduits with complex and complicated geometries may be particularly difficult to form in high-temperature components. Thus, high-temperature components with cooling geometry as well as associated methods of producing such components would be useful.

[0043] Accordingly, the present disclosure relates to systems and methods that utilize a dual-material, ephemeral / fusible mandrel assembly to form one or more cavities in a composite component, such as a CMC or PMC motor component. In one embodiment, the composite component may be a rotor blade or a stator vane. In particular, the mandrel assembly includes at least one mandrel having an inner structural body and an outer release layer. More particularly, in one embodiment, the inner structural body is generally formed of a metal alloy, such as a tin-zinc alloy, that defines and maintains cavity dimensions by compacting and consolidating a green core (also known as a preform) of the composite component.Thus, after deposition of the green core, the inner structural body may be melted from the green core, for example, before final processing of the green core. Further, in one embodiment, the outer release layer is generally formed of a silicone sheath to prevent contamination of the green core by residual material left after melting of the inner structural body. In one embodiment, the outer release layer may be completely removed after removal of the inner structural body, for example, to ensure that molten metal does not contaminate the green core during removal. In some embodiments, for example, the inner release layer may be mechanically stripped after melting of the inner structural body.

[0044] In another embodiment, in addition to the composite component described above, the mandrel assembly according to the present disclosure may be used in the manufacture of larger composites (e.g., PMC, CMC, etc.). In additional embodiments, the mandrel assembly according to the present disclosure may be used as a stacking support for green plies, which are then processed to a secondary green state (e.g., CMC) or to completion (e.g., PMC).

[0045] It should be noted that, although the present disclosure is generally described herein with reference to a gas turbine engine, the disclosed systems and methods may generally be used on components in any suitable type of turbine engine, including aircraft turbine engines, ground turbine engines, or steam turbine engines. Further, although the present disclosure is generally described with reference to stators and rotors in a turbine section, the disclosed systems and methods may generally be used on any component subject to increased temperatures where cooling may be desirable.

[0046] Referring now to the drawings, in which like reference numerals indicate like elements throughout, [Fig. 1] is a schematic sectional view of a gas turbine engine 10 in accordance with one embodiment of the present disclosure. More particularly, for the embodiment of [Fig. 1], the gas turbine engine 10 is configured as a high bypass ratio turbojet engine. However, in other embodiments, the gas turbine engine 10 may be configured as a low bypass ratio turbojet engine, a turbojet, a turboprop, a turboshaft engine, or other turbomachines known in the art. As shown in [Fig. 1], the gas turbine engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference) and a radial direction R.Generally, the gas turbine engine 10 includes a fan section 14 and a core turbine engine 16 disposed downstream of the fan section 14.

[0047] The core turbine engine 16 shown generally comprises a substantially tubular outer casing 18 which defines an annular inlet 20. The outer casing 18 encloses, in a series flow relationship, a compressor section 21 comprising a supercharger or low pressure (LP) compressor 22 and a high pressure (HP) compressor 24; a combustion section 26; a turbine section 27 comprising a high pressure (HP) turbine 28 and a low pressure (LP) turbine 30; and a jet exhaust nozzle section 32. The gas turbine engine 10 includes at least one rotating shaft 33 drivingly coupled between the compressor section 21 and the turbine section 27. For example, a high pressure (HP) shaft or spool 34 may drivingly connect the HP turbine 28 to the HP compressor 24. Similarly, a low pressure (LP) shaft or spool 36 may drivingly connect the LP turbine 30 to the LP compressor 22.

[0048] For the illustrated embodiment, the fan section 14 includes a variable pitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced apart manner. As shown, the fan blades 40 extend outwardly from the disk 42 generally along the direction radial R. Each fan blade 40 is rotatable relative to the disc 42 about a pitch axis P because the fan blades 40 are operatively coupled to a suitable actuator 44 configured to vary the pitch of the fan blades 40. The fan blades 40, the disc 42 and the actuator 44 are rotatable together about the centerline 12 by the LP shaft 36 through a power transmission box 46. The power transmission box 46 has a plurality of gears for lowering the rotational speed of the LP shaft 36 to a more efficient fan rotational speed.

[0049] Still referring to the embodiment of [Fig. 1], the disk 42 is covered by a rotatable forward nacelle 48 aerodynamically profiled to promote airflow through the plurality of fan blades 40. In addition, the fan section 14 includes an annular fan housing or outer nacelle 50 that circumferentially surrounds the fan 38 or at least a portion of the core turbine engine 16. It should be noted that the outer 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. Furthermore, a downstream section 54 of the outer nacelle 50 may extend over an outer portion of the core turbine engine 16 so as to define a bypass airflow passage 56 therebetween.

[0050] During operation of the gas turbine engine 10, a volume of air 58 enters the gas turbine engine 10 through an associated inlet 60 of the external nacelle 50 or fan section 14. As the volume of air 58 passes through the fan blades 40, a first portion of the volume of air 58 as indicated by the arrows 62 is directed or routed into the bypass airflow passage 56 and a second portion of the air 58 as indicated by the arrows 64 is directed or routed into the LP compressor 22. The ratio of the first portion of air 62 to the second portion of air 64 is commonly referred to as a bypass ratio. The pressure of the second portion of air 64 is then increased as it is passed through the HP compressor 24 and into the combustion section 26, where it is mixed with fuel and combusted to provide a hot combustion gas 66.

[0051] The hot combustion gases 66 are routed through the HP turbine 28 where a portion of the thermal or kinetic energy of the hot combustion gases 66 is extracted via sequential stages of HP turbine stator vanes 68 which are coupled to the outer casing 18 and HP turbine rotor blades 70 which are coupled to the HP shaft or spool 34, thereby causing rotation of the HP shaft or spool 34, thereby supporting operation of the HP compressor 24. The hot combustion gases 66 are then routed through the LP turbine 30 where a second portion of thermal and kinetic energy is extracted from the hot combustion gases 66 via sequential stages of LP turbine stator vanes 72 which are coupled to the casing external 18 and LP turbine rotor blades 74 which are coupled to the LP shaft or coil 36, thereby causing rotation of the LP shaft or coil 36, thereby supporting operation of the LP compressor 22 or rotation of the variable pitch fan 38.

[0052] The hot combustion gas 66 is then routed through the jet exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. Simultaneously, the pressure of the first portion of air 62 is substantially increased as the first portion of air 62 is routed through the bypass airflow passage 56 before it is exhausted from a fan nozzle exhaust section 76 of the gas turbine engine 10, also providing propulsive thrust. At least one of the combustion section 26, the HP turbine 28, the LP turbine 30, or the jet exhaust nozzle section 32 at least partially defines a flow path 78 for routing the hot combustion gas 66 through the core turbine engine 16.Various components may be positioned in the flow path 78 such as HP turbine stator vanes 68, HP turbine rotor blades 70, LP turbine stator vanes 72, or LP turbine rotor blades 74. Further, such components may require cooling to withstand the increased temperatures of the hot flue gas 66.

[0053] Referring now to [Fig. 2], a perspective view of one embodiment of a component 100 of the gas turbine engine 10 ([Fig. 1]) is illustrated in accordance with aspects of the present disclosure. In particular, [Fig. 2] illustrates the component 100 configured as a turbine rotor blade. In other embodiments, the component 100 may be any other component of the gas turbine engine 10 such as various shrouds, linings, bands, etc. of the gas turbine engine 10. For example, the component 100 may be a structure that at least partially defines the flow path 78 for the hot combustion gas 66, abuts the flow path 78 for the hot combustion gas 66, or extends into the flow path 78 for the hot combustion gas 66.

[0054] As shown in [Fig. 2], the component 100 may include a core 101 and an outer enclosure 103. The core 101 may include an outer surface 112 (Figures 4A-4C) extending a length between a first end 111 and a second end 113. Further, the core 101 may at least partially define a cavity 116, such as a cooling cavity, extending from the first end 111 along at least a portion of the length of the core 101. For example, as shown in [Fig. 2], the core 101 includes two cavities 116 that extend from the first end 111 and stop short of the second end 113. In another embodiment, the cavities 116 described herein may extend the entire length of the core 101 (e.g., from the first end 111 to the second end 113) or for less than the entire length of the core 101, depending, for example, on whether it is desirable to pressurize the cavity 116. In one embodiment, for example, to pressurize the cavity 116, the cavity 116 stops short of the second end 113, as shown in [Fig. 2]. In such embodiments, the air within the cavity 116 may be used to feed smaller cooling channels (not shown) that connect to the outer enclosure 103.

[0055] The core 101 and the cavity 116 will be described in more detail below with reference to, for example, Figures 2 and 3 and 4A-4C. The cavity 116 may be fluidly coupled to an air supply (not shown) at the first end 111 to provide cold air 115 to the component 100.Further, as illustrated, the outer enclosure 103 may abut the flow path 78 such that the hot combustion gas 66 flows past or through the component 100. The outer enclosure 103 may have an outer surface 107. The outer enclosure 103 is positioned outside the core 101 and extends from the first end 111 of the core 101 along at least a portion of the length of the core 101.

[0056] Still referring to [Fig. 2], in one embodiment, the component 100 may be a turbine rotor blade. For example, the turbine rotor blade may be the LP turbine rotor blade 74 or the HP turbine rotor blade 70. In other embodiments, the component 100 may be any other turbine rotor blade of the gas turbine engine 10 ([Fig. 1]), such as an intermediate turbine blade. In such embodiments, the component 100 may include an inner band 102 positioned at the first end 111 with an inner band surface 105. For example, the inner band surface 105 may at least partially define the flow path 78 such that the hot combustion gas 66 flows through the flow path 78. Thus, the inner band surface 105 may define an innermost boundary of the flow path 78 in a radial direction R.In a particular embodiment, the internal band 102 may be configured as a platform.

[0057] Additionally, as illustrated, a blade root 86 may be coupled to a turbine rotor disk (not shown), which is in turn coupled to the rotating shaft 33 (e.g., [Fig. 1]). It will be readily understood that, as shown in [Fig. 2] and as generally known in the art, the blade root 86 may define a projection 89 having a dovetail or other shape for receipt in a complementary shaped slot in the turbine rotor disk to couple the turbine rotor blade 70, 74 to the disk. Of course, each turbine rotor blade 70, 74 may also be coupled to the turbine rotor disk or rotating shaft 33 in other ways. Generally, the hot combustion gases 66 may flow from the section of combustion 26 ([Fig.l]) upstream of the component 100 beyond or through the component 100. It will be recognized that the flow path 78 may further be defined by the outer casing 18 as described with reference to [Fig.l] or adjacent components 100 having respective inner bands 102. The inner band 102 may be heated by the hot combustion gas 66 flowing past the inner band 102.

[0058] In any event, as shown in [Fig. 2], the turbine rotor blades 70, 74 may be coupled to the turbine rotor discs such that a row of circumferentially adjacent turbine rotor blades 70, 74 extend radially outward from the perimeter of each disc, i.e., into the flow path 78 ([Fig. 1]). Further, as shown in [Fig.l], the hot combustion gas 66 flowing through the flow path 78 may create a pressure differential across the turbine rotor blades 70, 74 causing the turbine rotor blades 70, 74 and thus the rotating shaft 33 to rotate. Thus, the turbine rotor blades 70, 74 may transform the kinetic or thermal energy of the hot combustion gas 66 into rotational energy to drive other components of the gas turbine engine 10 (e.g., one or more compressors 22, 24 via one or more rotating shafts 33).

[0059] It will be recognized that the outer enclosure 103 (Figures 2 and 3) may be configured as an airfoil 80. Accordingly, in such embodiments, as shown in Figures 2 and 3, the outer surface 107 may include an airfoil surface 85. Further, the airfoil surface 85 may include a pressure side 82 and a suction side 84. The airfoil surface 85 may also include a leading edge 88 at a forward position (e.g., FWD) of the airfoil 80 in the axial direction A. The airfoil surface 85 may further include a trailing edge 90 at a rearward position (e.g., AFT) of the airfoil 80 in the axial direction A. Furthermore, as illustrated, the airfoil 80 may extend from the blade root 86 to a blade tip 87 along an extent S.For example, the blade root 86 may be approximately at the first end 111 of the core 101, and the blade tip 87 may be approximately at the second end 113. Thus, the airfoil 80 may extend into the flow path 78 of the hot combustion gas 66. Further, the hot combustion gas 66 may flow over a combination of the pressure side 82, the suction side 84, the leading edge 88, or the trailing edge 90 and thereby heat the airfoil 80. The airfoil 80 may define the chord C extending axially between the opposing leading and trailing edges 88, 90. Furthermore, the airfoil 80 may define a width W between the side . pressure side 82 and the suction side 84. The width W of the aerodynamic profile 80 can vary along the extent S.

[0060] Furthermore, as shown in [Fig. 2], the cavity(ies) 116 may be at least partially defined by the core 101. Thus, cold air 115 may be directed toward and cool the component 100. In particular, it will be recognized that the cavity 116 may be fluidly coupled to the air supply and receive pressurized cold air 115 from the compressor section 21 (see, for example, [Fig. 1]). In such embodiments, the air supply may be the cold air supply 115. In other embodiments, the cold air 115 may be pressurized cold air 15 from another component of the gas turbine engine 10, such as a pump or the bypass airflow passage 56. The cold air 115 received in the cavity 116 is generally cooler than the hot combustion gas 66 flowing against or over the outer surface 107 of the outer enclosure 103.

[0061] Referring now to [Fig. 3], a perspective view of one embodiment of a component 100 is illustrated according to aspects of the present disclosure. In particular, [Fig. 3] illustrates the component 100 configured as a stator vane. For example, the component 100 may be the HP turbine stator vane 68 of the HP turbine 28 or the LP turbine stator vane 72 of the LP turbine 30. In such embodiments, the component 100 may include an outer band 108 positioned at the second end 113 of the turbine stator vane, for example, radially outward from the inner band 102. Further, the outer band 108 may include an outer band surface 109. For example, the outer band surface 109 may at least partially define the flow path 78 for the hot combustion gas 66. Thus, the outer band surface 109 may define an outermost boundary of the flow path 78.

[0062] Each turbine stator blade 68, 72 may have the outer enclosure 103 configured as the airfoil 80, or, more particularly, configured as a vane, that extends from the first end 111, such as between the first end 111 and the second end 113. For example, the outer enclosure 103 or the airfoil 80 may extend between the inner band 102 and the outer band 108. Each turbine stator blade 68, 72 may have the same characteristics as the airfoil 80 described above with reference to the turbine rotor blade 70, 74. For example, the stator blade 68, 72 may have a pressure side 82 opposite a suction side 84. The opposing pressure and suction sides 82, 84 of each airfoil 80 may extend radially along the extent S from a blade root at the inner band 102 to a blade tip at the outer band 108.Furthermore, the pressure sides and . suction 82, 84 of the airfoil 80 may extend axially between a leading edge 88 and an opposite trailing edge 90.

[0063] It will be appreciated that, although the airfoil 80 of the turbine stator vane 68, 72 may have the same characteristics as the airfoil 80 of the turbine rotor blade 70, 74, the airfoil 80 of the turbine stator vane 68, 72 may have a different configuration than the airfoil 80 of the turbine rotor blade 70, 74. For example, airfoils 80 of the LP turbine stator vanes 72 or airfoils 80 of the HP turbine rotor blades 70 may differ in size, shape, or configuration from the airfoils 80 of the HP turbine stator vanes 68 and the LP turbine rotor blades 74.However, it should also be understood that, although the airfoils 80 may differ in size, shape, or configuration, the disclosure described herein may be applied to any airfoil 80 in the gas turbine engine 10, as well as other suitable components 100 of the gas turbine engine 10.

[0064] The turbine stator vanes 68, 72 may direct the hot combustion gases 66 through the flow path 78. Further, the turbine stator vanes 68, 72 may increase the velocity of the hot combustion gas 66, thereby increasing the dynamic pressure while decreasing the static pressure of the hot combustion gas 66. In such embodiments, the outer band 108 may at least partially define the flow path 78. Further, the airfoil surface 85 or the outer band surface 109 may be heated by the hot combustion gas 66 flowing through the flow path 78.

[0065] Referring now to Figures 4A-4C, various cross-sectional views of one embodiment of component 100 (e.g., turbine rotor blades 70, 74) of [Fig. 2]) are illustrated in accordance with aspects of the present disclosure. It should be recognized that component 100 may be configured as a turbine rotor blade or a turbine stator vane, as well as any of turbine stator vanes 68, 72 as described with reference to [Fig. 3]. Thus, outer enclosure 103 may be configured as airfoil 80 and outer surface 107 may be airfoil surface 85. However, in other embodiments, component 100 or outer enclosure 103 may comprise any other structure exposed to hot combustion gas 66.

[0066] Further, as shown, component 100 may include core 101 and outer enclosure 103. In one embodiment, as shown, core 101 may generally have the same shape as outer enclosure 103. For example, in embodiments where outer enclosure 103 is an airfoil 80, core 101 may also be generally shaped as an airfoil. In other embodiments, additional layers or materials may envelop the outer enclosure 103 or may be incorporated into the outer enclosure 103. For example, the outer enclosure 103 may be formed of various layers, or various thermal coatings may be applied or sprayed onto the outer enclosure 103. Still referring to Figures 4A-4C, in one embodiment, the outer enclosure 103 may be adhered to at least a portion of the outer surface 112 using any suitable means, such as adhesives, tape, welding, or mechanical fasteners (e.g., bolts, screws, and rivets).

[0067] Further, the core 101 may at least partially define the cavity 116 described herein. For example, in one embodiment, the cavity 116 may extend the entire length of the core 101. Further, in one embodiment, the cavity 116 is fluidly coupled to the cold air supply 115 at the first end 111 (see, for example, Figures 2 and 3). Thus, in one embodiment, the cavity 116 may provide the fresh air 115 along the entire length of the core 101 or along the entire length of the outer enclosure 103. In addition, Figures 4A-4C, generally illustrate sectional views of the turbine rotor blade 70, 74 of [Fig. 2] along various locations along the extent thereof, i.e., to indicate cross-sections of the cavity(ies) 116 diverging along the extent of the turbine rotor blade 70, 74.

[0068] In still other embodiments, as shown in Figures 4A-4C, the core 101 may define a plurality of cavities 116, such as two or more cavities 116, such as three or more, configured generally identically to one another. It should be recognized that a portion of the cavities 116 may extend the entire length of the core 101 from the cold air supply 115 at the first end 111 to the second end 113. In other embodiments, some of the cavities 116 may extend along only a portion of the core 101.

[0069] In one embodiment, at least one of the core 101 or the outer enclosure 103 may be formed from a composite material such as a CMC material or other suitable composite material having high temperature capability. Composite materials generally include a fibrous reinforcing material embedded in a matrix material, such as a polymeric or ceramic material. The reinforcing material serves as the load-bearing constituent of the composite material, while the matrix of a composite material serves to bind the fibers together and act as a medium through which externally applied stress is transmitted and distributed to the fibers. For example, both the core 101 and the outer enclosure 103 may be formed from a CMC material.

[0070] In embodiments where the core 101 is formed from a CMC material, the cavity 116 may be formed in the core 101 while the core 101 is in a green state. For example, the core 101 in the green state may not be cured or may only be partially cured. It should be recognized that the core 101 in the green state may be more flexible, less brittle, and less rigid than a fully cured CMC core. The process of forming the component 100 is more fully described below with reference to Figures 5-8 below.

[0071] Exemplary CMC materials may include silicon carbide (SiC), silicon, silica, or alumina matrix materials and combinations thereof. Ceramic fibers may be incorporated into the matrix, such as oxidation-stable reinforcing fibers, including monofilaments such as sapphire and silicon carbide (e.g., SCS-6 from Textron), as well as rovings and yarns comprising silicon carbide (e.g., NICALON® from Nippon Carbon, TYRANNO® from Ube Industries and SYLRAMIC® from Dow Corning), alumina silicates (e.g., 440 and 480 from Nextel) and whiskers and chopped fibers (e.g., 440 and SAFFIL® from Nextel), and 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).For example, in some embodiments, bundles of the fibers, which may include a coating of ceramic refractory material, are formed as a reinforced ribbon, such as a unidirectional reinforced ribbon. A plurality of the ribbons may be stacked together (e.g., as plies) to form a preform component. The fiber bundles may be impregnated with a slurry composition before forming the preform or after forming the preform. The preform may then undergo a thermal treatment, such as curing or burning to provide a high-carbon residue in the preform, and a subsequent chemical treatment, such as melt infiltration with silicon, to result in a component formed from a CMC material having a desired chemical composition.In other embodiments, the CMC material may be formed as, for example, a carbon fiber fabric rather than a ribbon.

[0072] As indicated, it may be desirable to form components 100 of the gas turbine engine 10, such as components within or defining the flow path 78, e.g., stator vanes 68, 72, turbine rotor blades 70, 74, or other components, from composite materials such as CMC materials. The components 100 may be formed from a plurality of plies of the composite material, which are stacked together or assembled with other subassemblies, such as bundles of plies, preforms, or a stack of composite plies, to define the composite component. However, the cooling cavity(ies) 116 may be difficult to form in the core 101 of these components 100 after thermal or chemical processing. More particularly, the core 101 may be too rigid or brittle to form internal or complicated cooling cavities 116 as described herein. In addition, a ply stacking process for manufacturing these components may not be precise enough to form these types of internal or complicated passages. Furthermore, the outer enclosure 103 or plies forming the outer enclosure 103 may adhere more easily to a core 101 that is only partially processed (e.g., in the green state).

[0073] Referring now to [Fig. 5], an embodiment of a composite ply core 122 is shown in accordance with aspects of the present disclosure. It should be recognized that the composite ply core 122 may become the core 101 described with reference to Figures 1-4, for example after the component formation process is complete. As illustrated, the composite ply core 122 comprises a plurality of composite plies 124. As illustrated, the plurality of composite plies 124 may comprise a composite material such as a CMC material or a PMC material. The composite plies 124 may be stacked on a tool, mandrel, mold, or other suitable support device or surface. Preferably, each composite ply 124 of the plurality of composite plies 124 is cut, for example, from a ribbon as previously described, such that each composite ply 124 is oversized.That is, each of the plurality of composite plies 124 may be longer than a final ply length to provide machine stock for machining the green batch of plies to predetermined dimensions as described in more detail herein. Further, as shown in [Fig. 5], the plurality of composite plies 124 forms a composite ply stack 126. In some embodiments, the composite ply stack 126 may be a ply batch stack (which may also be referred to as a composite preform) or the like, which may be generally referred to as the composite ply stack 126. In some embodiments, at least one of the composite plies 124 may be a prepreg ply. For example, at least some or all of the composite plies 124 used to form the composite ply core 122 may be prepreg plies. In one embodiment, all of the composite plies 124 may be prepreg plies.In another embodiment, at least one of the composite plies 124 may be a CMC or PMC ply, such as all of the composite plies 124. In some embodiments, at least one of the composite plies 124 may be a CMC or PMC prepreg ply.

[0074] Referring now to [Fig. 6], a flowchart of one embodiment of a method 200 of forming a composite component according to aspects of the present disclosure is illustrated. In some embodiments, the component may be component 100 for a gas turbine engine 10 as generally described with reference to Figures 1-5. For example, component 100 may be any of turbine rotor blades 70, 74 or turbine stator vanes 68, 72 as described above with reference to Figures 1-3. In addition, although [Fig. 6] depicts steps performed in a particular order for purposes of illustration and discussion, the methods described herein are not limited to any particular order or arrangement. Those skilled in the art, using the disclosures provided herein, will understand that various steps of the methods disclosed herein may be omitted, rearranged, combined, or adapted in various ways without departing from the scope of the present disclosure.

[0075] As illustrated at reference numeral (202), the method 200 may include stacking a plurality of composite plies around a mandrel assembly to form a composite ply core (such as composite ply core 122), the mandrel assembly including a mandrel having an inner structural body and an outer release layer. Thus, the mandrel is configured to form a cavity in the composite ply core as the composite component is formed.

[0076] In particular, as shown in Figures 8A and 8B, a perspective view and a sectional view of an embodiment of a mandrel assembly 150 according to the present disclosure are illustrated, respectively. More specifically, as shown, the mandrel assembly 150 includes a mandrel 152 having an inner structural body 154 and an outer release layer 156. Thus, the mandrel 152 is configured to form one or more cooling cavities in the composite ply core 122 as the composite component is formed. For example, the cooling cavities may be generally configured like the cooling cavities 116 of Figures 2-4.Further, in one embodiment, the cavity 116 may be formed in the composite ply core 122 from the first end 111 as illustrated in Figures 2 and 3, such that the cavity 116 may be fluidly coupled to the cold air supply 115 after final processing. It should be recognized that the cooling cavities 116 described herein may be machined into any shape. For example, the cavity 116 may define a circular cross-section, an arcuate cross-section, an elongated cross-section, or any other suitable shape having one or more curved or straight segments.

[0077] Referring again to [Fig. 6], as illustrated at reference numeral (204), the method 200 may comprise processing the composite ply core containing the mandrel assembly. For example, in one embodiment, the method 200 may comprise partially processing the composite ply core containing the mandrel assembly to form a green core. Thus, the composite ply layer 126 may be partially processed to form a green layer of a green core 130 (see, for example, Figures 10C and 10D). In one embodiment of the method 200, the partial processing of the composite ply core may comprise (206) compacting the composite ply core. In another embodiment of the method 200, the partial processing of the composite ply core 122 may comprise (208) autoclaving the composite ply core. In yet another embodiment of the method 200, the partial processing of the composite ply core may comprise both compacting and autoclaving the composite ply core. For example, the composite ply core 122 ([Fig. 5]) may be compacted and then processed in an autoclave.

[0078] The compaction may be performed in the atmosphere, i.e., at ambient temperature and pressure or at an elevated temperature, and may use vacuum pressure, shrink film, or other means to apply the compaction force / pressure. The autoclave processing may be performed at a reduced temperature, at a reduced pressure, or for a shorter time compared to a standard autoclave cycle. In some embodiments, the partial processing of the composite ply stack 126 ([Fig. 5]) may involve only compaction, i.e., the composite ply stack 126 may be compacted without also undergoing a reduced autoclave cycle. In other embodiments, to partially process the composite ply stack 126, the stack may undergo a reduced autoclave cycle without being separately compacted.

[0079] In such embodiments, after partial processing, the composite plies 124 forming the composite ply stack 126 retain some flexibility and malleability. Such flexibility and malleability may assist in machining the composite ply stack 126. That is, the partial processing of the composite ply layer 126 achieves an adequate level of consolidation and hardening to achieve appropriate strength for subsequent handling, manipulation, and machining of the green stack. In contrast, a standard autoclave cycle is typically performed as part of processing a final ply or stack assembly to achieve final component dimensions and to stiffen the green core.More specifically, the standard autoclave cycle imparts rigidity to the final assembly of plies or stacks by complete drying or curing of the composite constituents and produces the final dimensions of the composite component by perfect consolidation of the plies or subassemblies.

[0080] Further, in embodiments in which the stack of composite plies 126 is processed in an autoclave, the stack of composite plies 126 may be autoclaved using soft or hard tooling. For example, the stack of composite plies 126 may be autoclaved using metal tooling, i.e., hard tooling, which is shaped to impart a desired shape to the composite ply stack 126. As another example, the composite ply stack 126 may be autoclaved using soft tooling such as a vacuum bag, for example, the composite ply stack 126 may be supported on a metal tool, and then the composite ply stack 126 and tool may be bagged and the air removed from the bag to apply pressure to the composite plies 124, and compact them, forming the composite ply stack 126 before the composite ply stack 126 is processed in a reduced autoclave cycle as previously described.

[0081] As shown, after the composite ply stack 126 is partially processed, the composite ply stack 126 is in a green state and thus forms a green state stack, which may be a subassembly used to form the composite component, such as the component 100 of Figures 2-4. Thus, as shown at reference numeral (210) of [Fig. 6], the method 200 may include removing the mandrel assembly from the composite ply core to form the cavity. More specifically, in one embodiment, the method 200 may include removing the mandrel assembly from the green state core to form the cavity. In such embodiments, removing the mandrel assembly may include melting the inner structural body 154 and then removing (e.g., by mechanical stripping) the outer release layer 156 after melting the inner structural body 154.Accordingly, in one embodiment, the inner structural body 154 may be formed of a low melting temperature alloy such as a tin-zinc alloy. Further, in such embodiments, a separate heat treatment may be required at an increased temperature to melt the low melting temperature metal alloy (e.g., when the melting temperature of the metal alloy is higher than temperatures used during the process to complete formation of the partially formed component 117). Accordingly, in one embodiment, the inner structural body 154 of the mandrel assembly 150 is configured to define and maintain the dimensions of the cavity 116 by compaction or consolidation, while the inner release layer 156 is configured to ensure that the molten metal alloy does not contaminate the composite ply core during removal of the inner structural body 154.Thus, after removal of the mandrel assembly 150, the cavity 116 described herein is formed.

[0082] Referring now to [Fig. 7], a flowchart of additional optional steps of the method 200 of forming the composite component according to aspects of the present disclosure is illustrated. In particular, in embodiments where the mandrel assembly 150 is used to form a CMC component, as shown at reference numeral (212), the method 200 may include wrapping one or more additional composite plies around the green core to form an outer enclosure. For example, the additional composite plies may be configured like the composite plies 124 used to form the core of composite plies 122. It will be recognized that the outer enclosure may generally be configured like the outer enclosure 103 as described with reference to Figures 2-4. Further, a portion of the composite plies 124 used to form the outer enclosure 103 may be prepreg plies. In yet another embodiment, at least one of the composite plies 124 may be a CMC prepreg ply, such as all of the composite plies 124. Thus, it will be recognized that at least one of the composite plies 124 may be bonded to the green core 130 during a final processing step of the method 200.

[0083] Referring now to [Fig. 9], a flowchart of one embodiment of a method 300 of forming the mandrel 152 of the mandrel assembly 150 according to the present disclosure is illustrated. Thus, in such embodiments, the mandrel 152 may be formed to have a shape corresponding to a desired shape of the one or more cavities 116. For example, as shown in Figures 8A and 8B, the mandrel 152 may be formed to have a single cross-sectional shape that may vary along the extent of the component 100 to be formed.

[0084] More specifically, in one embodiment, as shown in [Fig. 9], forming the mandrel 152 may include (302) machining a metal alloy 158 to form the inner structural body 154 (see, for example, Figures 8A and 8B), for example, via casting. Further, as shown in [Fig. 9], forming the mandrel 152 may include (304) coating the inner structural body 154 with an impermeable non-metallic material 160 to form the outer release layer 156 on the inner structural body 154 (see, for example, Figures 8A and 8B).In particular embodiments, for example, coating the inner structural body 154 with the impermeable non-metallic material 160 to form the outer release layer 156 on the inner structural body 154 may comprise (306) injection molding, (308) dip coating, (310) spraying, (312) winding, or any other suitable application technique. Further, in such embodiments, by way of example, the metal alloy 158 may be a tin-zinc alloy, while the impermeable non-metallic material 160 may be a silicone coating, a polytetrafluoroethylene (PTFE) / Teflon coating, or a sol-gel coating.

[0085] Accordingly, in particular embodiments, the impermeable non-metallic material 160 generally has thermal stability by melting the internal structural body 154, as well as non-stick properties. Thus, the external release layer 156 ensures that molten metal does not contaminate the preform. green during mandrel removal. After melting, the outer release layer 156 can be extracted from the cavity 116. As used herein, “thermal stability” is generally characterized by a melting temperature 10 degrees Celsius (°C) higher than the melting temperature of the inner structural body 154. Additionally, in one embodiment, the melting temperature of the inner structural body 154 is limited by the curing temperature of the composite and a thermal degradation temperature of the impermeable non-metallic material 160. For example, for the initially selected material (e.g., silicone), the melting temperature may be about 260°C.

[0086] Referring particularly to Figures 10A-10D, various components of a system 400 for forming a composite component according to aspects of the present disclosure are illustrated. In particular, as shown in Figures 10C and 10D, the partially formed component 117 is illustrated according to aspects of the present disclosure. In particular, the partially formed component 117 may be produced using any or all of the steps of the method 200 prior to the final processing step. Further, it should be recognized that the partially formed component 117 may form the component 100, as illustrated in Figures 2-4, after the final processing step.

[0087] The partially formed component 117 may generally be configured like the completed component 100 as described with reference to Figures 2-4. For example, the partially formed component 117 may include a core, such as the green core 130, and the outer enclosure 103. The core may include an outer surface 112 and may at least partially define the cavity 116 described herein. Further, the cavity 116 may be configured to be fluidly coupled to a cold air supply 115, for example after the partially formed component 117 is finally processed to form the component 100. Further, the partially formed component 117 may be configured to become a turbine rotor blade or a turbine stator vane. Thus, the outer enclosure 103 may be configured as an airfoil 80 and an outer surface 107 may be an airfoil surface 85.

[0088] However, the partially formed component 117 may include additional features not present in the finalized component 100. In particular, as shown in Figures 10C and 10D, the partially formed component 117 may include the mandrel assembly 150 defining and maintaining the dimensions of the cavity 116 by compacting or consolidating the preform. Further, in one embodiment, the core may include the green core 130 formed from a carbon matrix composite material. However, in other embodiments, the green core 130 may be formed from any suitable composite material.

[0089] It should also be recognized that the partially formed component 117 may not include all of the features of the completed component 100. For example, the cavity 116 may only be fluidly coupled to the cold air supply 115 after the component is finalized, e.g., after the mandrel assembly 150 has been removed.

[0090] The outer enclosure 103 may be positioned outside the core 101 and extend from the first end 111 of the core along at least a portion of the length of the core. It will be recognized that the outer enclosure 103 may be formed of one or more composite plies 124, which may be in an uncured or untreated state in the partially formed component 117 shown. Further, it will be recognized that multiple layers of composite plies 124 may be wrapped around the core, such as the green core 130, to form the outer enclosure 103. For example, at least one of the composite plies 124 may contact the exterior surface 112 of the core 101. In addition, various fillers or intermediate layers of other materials may be incorporated between composite plies 124.

[0091] Referring again to [Fig. 6], as indicated at reference numeral (214), the method 200 may include processing the green core and the outer enclosure to form the composite component. In one embodiment, the processing of the green core 130 and the outer enclosure 103 may include (216) autoclaving the green core and the outer enclosure to form an autoclaved body. Further, a further step may include (218) curing the autoclaved body to form a cured body. In some embodiments, the processing of the green core 130 and the outer enclosure 103 may include an additional combustion step, for example (220) for burning the composite plies. An additional step may include (222) densifying the cured body to form the composite component.In some embodiments, the treatment of the green core 130 and the outer enclosure 103 may include at least one of melt infiltration, polymer infiltration and pyrolysis, or chemical vapor infiltration (CVI).

[0092] For example, the processing may include autoclaving the assembled partially formed component 117 using a standard autoclave cycle, rather than a reduced autoclave cycle as previously described, to form an autoclaved body. In embodiments in which the composite material is a CMC material, the autoclaved body may then undergo baking (or combustion) to form a baked body, followed by densification to produce a densified CMC component that is a one-piece component, meaning the component is a continuous piece of CMC material.

[0093] For example, after autoclaving, the component may be placed in a furnace to burn off any mandrel-forming materials or solvents used in forming the CMC plies and to decompose binders in the solvents, and then placed in a furnace with silicon to convert a ceramic matrix precursor of the plies into the ceramic matrix material of the CMC component. The silicon melts and infiltrates into any porosity created with the matrix as a result of binder decomposition during burning / firing; melt infiltration of the CMC component with silicon densifies the CMC component. However, densification may be performed using any known densification technique, including, but not limited to, Silcomp, melt infiltration (MI), chemical vapor infiltration (CVI), polymer infiltration and pyrolysis (PIP), and oxide / oxide processes.In one embodiment, the densification and firing may be performed in a vacuum furnace or an inert atmosphere having an atmosphere set at temperatures above 1200°C to allow the silicon or other suitable material or materials to moltenly infiltrate the component.

[0094] Referring now to Figures 10A-10D, schematic diagrams of one embodiment of components of the system 400 for forming a composite component according to the present disclosure are illustrated. More specifically, as shown in Figures 10A, 10C, and 10D, the system 400 includes the mandrel assembly 150 described herein. Further, as particularly shown in [Fig. 10A], the mandrel assembly 150 may include one or more mandrels 152, such as a first mandrel 162 and a second mandrel 164, each having the inner structural body 154 and outer release layer 156 described herein. Additionally, as illustrated, the first mandrel 162 has a shape corresponding to a desired shape of a first cooling cavity, while the second mandrel 164 has a shape corresponding to a desired shape of a second cooling cavity.Thus, the first and second mandrels 162, 164 may be arranged together in a desired configuration prior to stacking a first plurality of composite plies 402 so as to define and maintain the dimensions of the first and second cooling cavities by compacting or consolidating the preform. More specifically, as mentioned, the inner structural body 154 of the first and second mandrels 162, 164 may be formed of a metal alloy, while the outer release layer 156 of the first and second mandrels 162, 164 may be formed of an impermeable non-metallic material, which may also be deformable or flexible.

[0095] Additionally, as shown in [Fig. 10B], the system 400 includes the first plurality of composite plies 402, such as the composite plies 124, to be placed around the mandrel assembly 150 to form the composite ply core 122 ([Fig. 10C]). Further, as shown in [Fig. 10C], the system 400 includes a consolidation apparatus 404 for processing the composite ply core 122 containing the mandrel assembly 150. Further, as shown in [Fig. 10B], the system 400 may include a second plurality of composite plies 406, such as the composite plies 124, to be wrapped around the green core 130 to form the outer enclosure 103. Furthermore, the consolidation apparatus 404 is further configured to process the green core 130 and the outer enclosure 103 to form the composite component. In one embodiment, the consolidation apparatus 404 may comprise a compaction device, an autoclave device, a melt infiltration device, a polymer infiltration and pyrolysis device, or a chemical vapor infiltration (CVI) device.

[0096] Optionally, after processing, the composite component may undergo finish machining, if and when necessary, and be coated with one or more coatings, such as an environmental barrier coating (EBC). For example, the composite plies 124 that are wrapped around the core 101 may be oversized such that a portion of the composite plies 124 extend beyond the desired trailing edge 90 of the airfoil 80. Accordingly, after processing, the composite plies 124 may be machined to define the trailing edge 90. In other embodiments, the composite plies 124 may be machined after autoclaving the outer enclosure 103, but before curing and densifying the outer enclosure 103.

[0097] The mandrel assembly and associated mandrel assembly described herein provide numerous advantages not present in the prior art. The fabrication of the inner structural body defines and maintains the cavity dimensions through compaction and consolidation of the preform, while the outer release layer ensures that molten metal from melting the inner structural body does not contaminate the preform (e.g., Sn contamination) during mandrel removal with residual material. The outer release layer can also be easily manually removed after melting the inner structural body.

[0098] The following additional aspects of the invention are provided:

[0099] A method of forming a composite component, the method comprising: stacking a plurality of composite plies around a mandrel assembly to form a composite ply core, the mandrel assembly comprising a mandrel having an inner structural body and an outer release layer, the mandrel configured to form a cavity in the composite ply core; processing the composite ply core containing the mandrel assembly to compact the plurality of composite plies together; and removing the mandrel assembly from the ply core composites to form the cavity, wherein removing the mandrel assembly from the composite ply core includes melting the inner structural body and removing the outer release layer after melting the inner structural body.

[0100] The method may further comprise forming the mandrel to have a shape corresponding to a shape of the cavity.

[0101] Forming the mandrel further comprises: fabricating a metal alloy to form the inner structural body; and coating the inner structural body with an impermeable non-metallic material to form the outer release layer on the inner structural body.

[0102] The metal alloy has a melting temperature above a curing temperature of the composite component and below a barrier layer temperature.

[0103] The impermeable non-metallic material is a polymeric material comprising at least one of silicone, polytetrafluoroethylene, or a sol-gel.

[0104] Coating the inner structural body with the impermeable non-metallic material to form the outer release layer on the inner structural body further comprises using at least one of injection molding, dip coating, spraying, or winding to coat the inner structural body with the impermeable non-metallic material to form the outer release layer.

[0105] Partially processing the composite ply core containing the mandrel assembly to form a green core prior to processing the composite ply core; wrapping one or more additional composite plies around the green core to form an outer enclosure; and processing the green core and the outer enclosure to form the composite component.

[0106] Partial processing of the composite ply core comprises at least one of compacting the composite ply core or autoclaving the composite ply core.

[0107] Processing the green core and outer enclosure includes: autoclaving the green core and outer enclosure to form an autoclaved body; curing the autoclaved body to form a cured body; and densifying the cured body to form the composite component.

[0108] The treatment of the green core and the outer enclosure comprises: at least one of melt infiltration, polymer infiltration and pyrolysis, or chemical vapor infiltration.

[0109] At least one of the plurality of composite plies is a ceramic matrix composite ply.

[0110] The composite component is a gas turbine engine component.

[0111] The composite component is at least one of a turbine rotor blade or a turbine stator vane.

[0112] A system for forming a composite component, the system comprising: a mandrel assembly comprising a first mandrel having an inner structural body and an outer release layer; a first plurality of composite plies to be placed around the mandrel assembly to form a composite ply core, the first mandrel being configured to form a cavity in the composite ply core; and a consolidation apparatus for processing the composite ply core containing the mandrel assembly, wherein, after processing the composite ply core, the inner structural body is melted from the composite ply core and the outer release layer is removed after melting the inner structural body.

[0113] The system further comprising a second plurality of composite plies to be wrapped around a green core formed by consolidating the core of composite plies to form an outer enclosure, wherein the consolidation apparatus is further configured to process the green core and the outer enclosure to form the composite component.

[0114] The first mandrel has a shape corresponding to a desired shape of the cavity, and wherein the inner structural body of the first mandrel is formed of a metal alloy and the outer release layer is formed of an impermeable non-metallic material.

[0115] The mandrel assembly further includes a second mandrel for forming a second cavity in the composite ply core.

[0116] At least one of the plurality of composite plies is a ceramic matrix composite ply.

[0117] The consolidation apparatus comprises at least one of a compaction device, an autoclave device, a melt infiltration device, a polymer infiltration and pyrolysis device, or a chemical vapor infiltration (CVI) device.

[0118] This written description uses exemplary embodiments to disclose the disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods.

Claims

Claims

1. A method (200) of forming a composite component (100), the method (200) comprising: stacking a plurality of composite plies (124) around a mandrel assembly (150) to form a core of composite plies (122), the mandrel assembly (150) comprising a mandrel (152) having an inner structural body (154) and an outer release layer (156), the mandrel (152) being configured to form a cavity (116) in the core of composite plies (122); processing the core of composite plies (122) containing the mandrel assembly (150) to compact the plurality of composite plies (124) together; and removing the mandrel assembly (150) from the composite ply core (122) to form the cavity (116), wherein removing the mandrel assembly (150) from the composite ply core (122) includes melting the inner structural body (154) and removing the outer release layer (156) after melting the inner structural body (154).

2. The method (200) of claim 1, further comprising forming the mandrel (152) to have a shape corresponding to a desired shape of the cavity (116).

3. The method (200) of claim 2, wherein forming the mandrel (152) further comprises: fabricating a metal alloy (158) to form the inner structural body (154); coating the inner structural body (154) with an impermeable non-metallic material (160) to form the outer release layer (156) on the inner structural body (154).

4. The method (200) of claim 3, wherein the metal alloy (158) has a melting temperature greater than a curing temperature of the composite component (100) and less than a barrier layer temperature.

5. The method (200) of claim 3, wherein the impermeable non-metallic material (160) is one of: silicone, polytetrafluoroethylene, or a sol-gel.

6. The method (200) of claim 3, wherein coating the internal structural body (154) with the non-metallic material waterproof (160) to form the outer release layer (156) on the inner structural body (154) further comprises using at least one of injection molding, dip coating, spraying or winding to coat the inner structural body (154) with the waterproof non-metallic material (160) to form the outer release layer (156).

7. The method (200) of any preceding claim, further comprising: partially processing the composite ply core (122) containing the mandrel assembly (150) to form a green core (130) prior to processing the composite ply core (122); wrapping one or more additional composite plies (124) around the green core (130) to form an outer enclosure (103); and processing the green core (130) and the outer enclosure (103) to form the composite component (100).

8. The method (200) of claim 7, wherein the partial processing of the composite ply core (122) comprises at least one of compacting the composite ply core (122) or autoclaving the composite ply core (122).

9. The method (200) of claim 7, wherein processing the green core (130) and the outer enclosure (103) comprises: autoclaving the green core (130) and the outer enclosure (103) to form an autoclaved body; curing the autoclaved body to form a cured body; and densifying the cured body to form the composite component (100).

10. The method (200) of claim 7, wherein the treatment of the green core (130) and the outer enclosure (103) comprises: at least one of melt infiltration, polymer infiltration and pyrolysis, or chemical vapor infiltration.

11. The method (200) of any preceding claim, wherein at least one of the plurality of composite plies (124) is a ceramic matrix composite ply.

12. A method (200) according to any preceding claim, wherein the composite component (100) is at least one of a turbine rotor blade or a turbine stator vane.

13. A system (400) for forming a composite component (100), the system (400) comprising: a mandrel assembly (150) comprising a first mandrel (152) having an inner structural body (154) and an outer release layer (156); a first plurality of composite plies (124) to be placed around the mandrel assembly (150) to form a core of composite plies (122), the first mandrel (152) being configured to form a first cavity (116) in the core of composite plies (122); and a consolidation apparatus (404) for processing the composite ply core (122) containing the mandrel assembly (150), wherein, after processing the composite ply core (122), the inner structural body (154) is melted from the composite ply core (122) and the outer release layer (156) is removed after melting the inner structural body (154).

14. The system (400) of claim 13, further comprising a second plurality of composite plies (124) to be wrapped around a green core (130) formed by consolidating the core of composite plies (122) to form an outer enclosure (103), wherein the consolidation apparatus (404) is further configured to process the green core (130) and the outer enclosure (103) to form the composite component (100).

15. The system (400) of claim 13, wherein the inner structural body (154) of the first mandrel (152) is formed of a metal alloy and the outer release layer is formed of an impermeable non-metallic material.

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